Boeing 757
Systems Course
1. General
2. Air
3. Electrics
4. Fire Protection
5. Flight Controls
6. Fuel
7. Hydraulics
8. Ice and Rain
9. Landing Gear and Brakes
10. Oxygen System
11. Pneumatics
12. Potable Water
13. Waste Disposal System
14. APU
15. Doors
16. P&W Engine
17. RR RB211
AeroEd LLC – Aviation Education Resource
B757 GENERAL FAMILIARIZATION
This course covers an overview of the Mechanical Systems to include: Description and Operation, Controls and
Indications, Component Location, & Servicing.
OBJECTIVES
Upon completion of this training, using the study guide provided and appropriate Maintenance Manuals, the
student will be able to:
1) Describe the safety precautions to be observed when working on or near the aircraft and its systems.
2) Describe the locations of principle components.
3) Describe the normal functions of each major system, including terminology and nomenclature.
4) Using the proper maintenance manual reference, perform all aircraft system servicing tasks.
5) Interpret reports provided by the crew members.
757 General Familiarization (7-2005)
ATA 06 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 06 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 06
B757 GENERAL FAMILIARIZATION SELF-PACED ........................................................................... 1
OBJECTIVES ............................................................................................................................................. 1
MANUAL ARRANGEMENT AND NUMBERING SYSTEM ............................................................... 5
Chapter Numbering............................................................................................................................. 5
Effectivity and Configuration Numbering .......................................................................................... 6
Page Numbering ................................................................................................................................. 8
LIST OF ABBREVIATIONS ..................................................................................................................... 9
REFERENCE PLANES AND LINES ..................................................................................................... 15
Standard Abbreviations and Definitions ........................................................................................... 15
Fuselage ............................................................................................................................................ 15
Wing.................................................................................................................................................. 15
Vertical Stabilizer ............................................................................................................................. 16
Horizontal Stabilizer ......................................................................................................................... 17
Power Plant ....................................................................................................................................... 18
PRIMARY AIRCRAFT DIMENSIONS .................................................................................................. 23
DIMENSIONS .......................................................................................................................................... 25
Overall Airplane: .............................................................................................................................. 25
Wing: ................................................................................................................................................ 25
Horizontal Stabilizer: ........................................................................................................................ 25
Vertical Stabilizer: ............................................................................................................................ 25
Fuselage: ........................................................................................................................................... 26
Areas: ................................................................................................................................................ 26
BODY STATION DIAGRAM ................................................................................................................. 28
VERTICAL STABILIZER AND RUDDER STATION DIAGRAM ..................................................... 30
HORIZONTAL STABILIZER AND ELEVATOR STATION DIAGRAM ........................................... 31
WING STATION DIAGRAM ................................................................................................................. 32
ENGINE AND NACELLE STATION DIAGRAM ................................................................................ 33
ZONE DIAGRAMS ................................................................................................................................. 34
Major Zones ...................................................................................................................................... 34
SERVICE INTERPHONE SYSTEM ....................................................................................................... 36
Component Details ........................................................................................................................... 36
Audio Amplifier................................................................................................................................ 36
Service Interphone Switch ................................................................................................................ 36
Operation .......................................................................................................................................... 38
Control .............................................................................................................................................. 38
Cabin Interphone System .................................................................................................................. 40
Cabin Interphone Handsets ............................................................................................................... 40
Pilot’s Call Panel ............................................................................................................................... 40
Operation .......................................................................................................................................... 40
Control .............................................................................................................................................. 42
757 General Familiarization (7-2005)
ATA 06 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 06 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
MANUAL ARRANGEMENT AND NUMBERING SYSTEM
The Maintenance Manual is divided into chapters and groups of chapters. Each group and every chapter has a tab
provided for ease of location. The chapterization separates the manual into the primary functions and systems of
the airplane. The chapters are further divided into sections and subjects to provide for subsystem and individual
unit breakout. Each chapter, section and subject is identified by an assigned number. Each page carries the
assigned subject number, page number, page code and the revision date.
In addition, the Power Plant chapters are issued in a self-contained set or sets (as applicable, if you have more
than one engine type in your model fleet). These pages are further identified by an engine sub-logo, for example
PW2000 SERIES ENGINES or RB211-535 SERIES ENGINES, placed to the right of the Maintenance Manual logo at
the top of the page. The numbering system is described in detail in the paragraphs that follow.
Chapter Numbering
Chapterization of the maintenance manual has provided a functional breakdown of the entire airplane. The chapter
breakdown numbering system uses a three element number (XX-XX-XX). It provides for dividing the material into
Chapters, Sections, and Subjects.
The three elements of the indicator each contain two digits.
For example:
757 General Familiarization (7-2005)
ATA 06 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Chapter Numbering (Continued):
The chapter number (1st element) and the first number of the section number (2nd element) are assigned by ATA
Specification No. 100. Material which is applicable to a system as a whole uses zeros in the 2nd and 3rd elements
of the numbers. That is, the chapter number followed by "-00-00".
For example:
AMM 22-00-00/001 (Auto Flight) is used for general description information which provides an outline breakdown
of the sections in the chapter.
Effectivity and Configuration Numbering
On each page, there is effectivity data at the lower, inner margin (Fig. 1). When a page applies to all airplanes, the
word ALL is in the effectivity block. If the data does not apply to all airplanes, then the effectivity will be one of
these types:
1. Physical description - A description of the differences that you can see.
When a physical description is used, a reference to the applicable service bulletin and PRR (production
change) are provided when that is possible. This is done primarily for the benefit of airline engineering,
and maintenance planning groups.
For example: AIRPLANES WITH VALVE INSTALLED AWAY FROM THE FILTER (POST-SB 28A-17 OR
PRR 54009) AIRPLANES WITH VALVE INSTALLED NEAR THE FILTER (PRE-SB 28A-17)
2. Component dash number - The last digits of the identification number that are on an electrical box.
3. Airplane effectivity numbers - The airline three-letter code, and the numbers or letters that Boeing and
each airline agreed on to identify each airplane. If the effectivity is applicable to all subsequent airplanes,
the last digits will be 999.
For example:
205-999 indicates airplane 205 and all subsequent airplanes.
Each paragraph can have an effectivity. Each effectivity is in upper-case letters, on the first line of the
paragraph.
757 General Familiarization (7-2005)
ATA 06 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 06 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Effectivity and Configuration Numbering (Continued):
When effectivity differences are extensive and the preceding method becomes cumbersome and distracting from
the continuity of subject matter, new page blocks are created. These added page blocks are identified by the
addition of a configuration code (CONFIG) immediately above the page number. A previously issued page block is
re-issued to incorporate the configuration code as shown in Fig. 1. Configuration codes are issued at page block
level only. They are usually used when a change to the airplane results in a major change to the manual.
Configuration codes are typically used when there are multiple configurations of page block applicable to a
customer's fleet.
In some instances, you can have CONFIGs that are provided as place holders. These procedures will be indicated
as "NOT USED" in the effectivity block in the lower left corner of the page (Fig. 1).
For the effectivity information in the power plant (70 series) chapters of the manual, two situations can exist. The
word ALL placed in the effectivity block on a page means that the page pertains to either all airplanes or all
engines, whichever the case may be. When the effectivity is limited to a system or component that remains with
the airplane during the power plant replacement, the effectivity is expressed in a manner described in the
preceding paragraphs. When a manual section, page, step or illustration is limited to an engine type or component,
the effectivity is given using the engine model, physical difference, or part number.
The word "ALL" in the effectivity block on a page means that the page pertains to all airplanes (if you have only
one engine type in your model fleet) or 2) All engines (if you have multiple engine types in your model fleet),
whichever the case may be.
Page Numbering
Each page block has its own page numbers. The page numbers are in the lower right corner of each page. The
page blocks categorize the tasks that they contain. The page blocks are defined by ATA Specification 100:
NOMENCLATURE
DESCRIPTION AND OPERATION (D&O)
FAULT ISOLATION (FI)
MAINTENANCE PRACTICES (MP)
SERVICING (SRV)
REMOVAL/INSTALLATION (R/I)
ADJUSTMENT/TEST (A/T)
INSPECTION/CHECK (I/C)
CLEANING/PAINTING (C/P)
APPROVED REPAIRS (AR)
PAGE BLOCK
1 to 99
101 to 199
201 to 299
301 to 399
401 to 499
501 to 599
601 to 699
701 to 799
801 to 899
When it is convenient for the user to have different types of tasks in one page block, MAINTENANCE PRACTICES,
the 201-to-299 page block, is used.
757 General Familiarization (7-2005)
ATA 06 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LIST OF ABBREVIATIONS
A/C: air conditioning
A/G: air/ground
A/L: auto land
A/P: autopilot
A/S: airspeed
A/T: auto throttle, adjustment/test
ABNORM: abnormal
AC: alternating current
ACARS ARINC: Communications Addressing and
Reporting System
ACCEL: acceleration, accelerate
ACM: air cycle machine
ADC: air data computer
ADF: automatic direction finder
ADI: attitude director indicator
ADP: air driven pump, air driven hydraulic pump
ADV: advance
AFCS: automatic flight control system
AGL: above ground level
AI: anti-ice
AIDS: aircraft integrated data system
AIL: aileron
ALT: altitude
ALTM: altimeter
ALTN: alternate
ALTNT: alternate
AMB: ambient
AMM: Airplane Maintenance Manual
ANN: announcement
ANNUNC: annunciator
ANT: antenna
AOA: angle of attack
APB: auxiliary power breaker
APD: approach progress display
APL: airplane
APPR: approach
APPROX: approximately
APU: auxiliary power unit
ARINC: Aeronautical Radio Incorporated
ARINC IO ARINC: I/O error
ARNC STP ARINC I/O UART: data strip error
ASA: auto land status annunciator
ASP: audio selector panel
ASYM: asymmetrical
757 General Familiarization (7-2005)
ATC: air traffic control
ATC/DABS: air traffic control/discrete address
beacon system
ATT: attitude
ATTND: attendant
AUTO: automatic
AUX: auxiliary
AVM: airborne vibration monitor
B/CRS: back course
BARO: barometric
BAT: battery
BFO: beat frequency oscillator
BITE: built-in test equipment
BK: brake
BKGRD: background
BPCU: bus power control unit
BRKR: breaker
BRT: bright
BTB: bus tie breaker
BTL: bottle
C/B: circuit breaker
C: center
°C: degrees Centigrade
CADC: central air data computer
CAPT: captain
CB: circuit breaker
CCA: central control actuator
CCW: counterclockwise
CDU: control display unit
CH: channel
CHAN : channel
CHG: change
CHR: chronograph
CHRGR: charger
CK: check
CKT: circuit
CL: close
CLB: climb
CLR: clear
CLSD: closed
CMD: command
CMPTR: computer
CNX: cancelled
ATA 06 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COL: column
COMM: communication
COMP: compressor
COMPT: compartment
CON: continuous
COND: condition
CONFG: configuration
CONFIG: configuration
CONN: connection
CONT: control
CP: control panel
CPCS: cabin pressure control system
CPS: cycles per second
CRS: course
CRT: cathode ray tube
CRZ: cruise
CSEU: control system electronics unit
CT: current transformer
CTN: caution
CTR: center
CU: control unit
CUST: customer
CW: clockwise
CWS: control wheel steering
DA: drift angle
DADC: digital air data computer
DC: direct current
DEC: decrease, decrement
DECEL: decelerate
DECR: decrease
DEG: degree
DEPR: depressurize
DEPT: departure
DEST: destination
DET: detector
DETNT: detent
DEV: deviation
DFDR: digital flight data recorder
DG: directional gyro
DH: decision height
DIFF: differential
757 General Familiarization (7-2005)
DIR: direct
DISC: disconnect
DISCH: discharge
DISCONT: discontinued
DISENG: disengage
DISP: dispatch
DIST: distance
DK: deck
DME: distance measuring equipment
DMU: data management unit
DN: down
DPCT : differential protection current
transformer
DR: door
DSCRT IO: discrete I/O error
DSPLY: display
DSPY: display
EADI: electronic attitude director indicator
ECON: economy
ECS: environmental control system
EDP: engine driven pump, engine hydraulic
pump
EEC: electronic engine control
EFDARS: expanded flight data acquisition
and reporting system
EFI: electronic flight instruments
EFIS: electronic flight instrument system
EGT: exhaust gas temperature
EHSI: electronic horizontal situation
indicator
EICAS: engine indicating and crew alerting
system
ELEC: electrical
ELEV: elevation
EMER: emergency
ENG: engage, engine
ENT: entrance, entry
ENTMT: entertainment
EPC: external power contactor
EPR: engine pressure ratio
EPRL: engine pressure ratio limit
ATA 06 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EQUIP: equipment
ERR: error
ESS: essential
EVAC: evacuation
EVBC: engine vane and bleed control
EXH: exhaust
EXT: external
EXTIN: extinguish, extinguished
EXTING: extinguishing
F/D: flight director
F/F: fuel flow
F/O: first officer
°F: degrees Fahrenheit
FAA: Federal Aviation Administration
FCC: flight control computer
FCEU: flight controls electronic unit
FCU: fuel control unit
FDR: feeder
FIM: Fault Isolation Manual
FL: flow
FL/CH: flight level change
FLD: field
FLT: flight
FLUOR: fluorescent
FMC: flight management computer
FMS: flight management system
FREQ: frequency
FRM: Fault Reporting Manual
FSEU: flap/slat electronic unit
FT: feet, foot
FWD: forward
G/S: glide slope, ground slope
GA: go-around
GB: generator breaker
GCB: generator circuit breaker
GCR: generator control relay
GCU: generator control unit
GEN: generator
GHR: ground handling relay
GND: ground
GP: group
757 General Familiarization (7-2005)
GPWS: ground proximity warning system
GR: gear
GRD: ground
GS: ground speed
GSSR: ground service select relay
GSTR: ground service transfer relay
GW: gross weight
H/L: high/low
HDG: heading
HF: high frequency
HORIZ: horizontal
HP: high pressure
HSI: horizontal situation indicator
HTR: heater
HYD: hydraulic
IAS: indicated airspeed
IDENT: identification
IDG: integrated drive generator
IGN: ignition
ILLUM: illuminate, illuminated
ILS: instrument landing system
IMP: imperial
IN: in, input
INBD: inboard
INC: incorporated, increase, increment
INCR: increase
IND: indicator
INFC: interface
INFLT: inflight
INHIB: inhibit
INIT: initiation
INOP: inoperative
INPH: interphone
INST: instrument
INT: interphone
INTLK: interlock
INTPH: interphone
INTMT: intermittent
IP: intermediate pressure
IRS: inertial reference system
IRU: inertial reference unit
ATA 06 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ISLN: isolation
ISOL: isolation
IVSI: instantaneous vertical speed indicator
KG: kilograms
KIAS: knots indicated airspeed
KTS: knots
L: left
L/R: left/right
L-NAV: lateral navigation
LAV: lavatory
LB: pound
LBS: pounds
LCD: liquid crystal display
LCR: left-center-right
LDG: landing
LDG GR: landing gear
LE: leading edge
LED: light emitting diode
LF: left front
LGT: light
LH: left hand
LIM: limit
LOC: localizer
LN: left nose
LR: left rear
LRRA: low range radio altimeter
LRU: line replaceable unit
LSB: lower side band
LVR: lever
LW: left wing
LWR: lower
M-SPD: manual speed
MAG: magnetic
MAINT: maintenance
MALF: malfunction
MAN: manual
MAX: maximum
MCDP: maintenance control display panel
MCP: mode control panel
MCU: modular concept unit
MDA: minimum decision altitude
757 General Familiarization (7-2005)
MIC: microphone
MIN: minimum
MM: Maintenance Manual
MOD: module
MON: monitor
MOT: motion
MPU: magnetic pickup
MSG: message
MSTR: master
MSU: mode selector unit
MTG: miles to go
MU: management unit
MUX: multiplexer
N/A: not applicable
NAC: nacelle
NAV: navigation
NCD: no computed data
NEG: negative
NEUT: neutral
NLG: nose landing gear
NO: number
NORM: normal
NRM: normal
NVMEM RD: non-volatile memory read error
NVMEM WR: non-volatile memory write
error
02: oxygen
OBS: observer
OK: okay
OPR: operate
OPT: option
OPRN: operation
OUT: output
OUTBD: outboard
OVHD: overhead
OVHT: overheat
OVRD: override
OXY: oxygen
P/RST: press to reset
P/S: pitot/static
PA: passenger address
ATA 06 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PASS: passenger
PCA: power control actuator
PCT: percentage
PDI: pictorial deviation indicator
PES: passenger entertainment system
PLA: power level angle
PLT: pilot
PMG: permanent magnet generator
PNEU: pneumatic
PNL: panel
POR: point of regulation
POS: position, positive
PPOS: present position
PRESS: pressure
PRG FLOW: program flow error
PRIM: primary
PROC: procedure
PROG MEM ROM: memory error
PROJ: projector
PROT: protection
PS: pitot static
PSI: pounds per square inch
PSS: passenger service system
PSU: passenger service unit
PTT: push to talk
PTU: power transfer unit
PWR: power
QAD: quick-attach-detach
QTS: quarts
QTY: quantity
R/T: rate of turn
R/W MEM RAM: memory error
R: right
RA: radio altimeter, radio altitude
RAT: ram air turbine
RCVR: receiver
RDMI: radio distance magnetic indicator
REC: recorder
RECIRC: re-circulate
REF: reference
REFRIG: refrigeration
757 General Familiarization (7-2005)
REG: regulator
REL: release
REP: representative
REQ: required
RES : reserve
RESSTART: power interrupt restart error
REV: reverse
RF: right front
RH: right hand
RLSE: release
RLY: relay
RLY/SW: relay/switch
RMI: radio magnetic indicator
RMT OUT: high-speed ARINC output error
RN: right nose
ROT: rotation
RPM: revolutions per minute
RPTG: reporting
RR: right rear
RST: reset
RTO: rejected takeoff
RUD: rudder
RW: right wing
RWY: runway
SAM: stabilizer trim/elevator asymmetry
limit module
SAT: static air temperature
SEC: second
SEI: standby engine indicator
SEL: select
SELCAL: selective calling
SERV: service
SG: signal generator
SLCTD: selected
SLCTR: selector
SOV: shut off valve
SP: speed
SPD: speed
SPD BK: speed brake
SQL: squelch
SSB: single side band
ATA 06 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STA: station
STAB: stabilizer
STBY: standby
STS: system status
SURF: surface
SW: switch
SWITCH IN: switch input error
SYNC: synchronous
SYS: system
SYST: system
T/R: thrust reverser
T.O. : takeoff
TACH: tachometer
TAI: thermal anti-ice
TAS: true airspeed
TAT: total air temperature
TCC: turbine case cooling
TE: trailing edge
TEMP: temperature
TFR: transfer
THR: thrust
THROT: throttle
THRSH: threshold
THRT: thrust
THRU: through
TIE: bus tie
TLA: thrust lever angle
TMC: thrust management computer
TMS: thrust management system
TMSP: thrust mode select panel
TO: takeoff
TOL: tolerance
TR: transformer rectifier
TRP: thrust rating panel
TUNE: tuner
TURB: turbine
TURBL: turbulent, turbulence
757 General Familiarization (7-2005)
UBR: utility bus relay
UPR: upper
USB: upper side band
V/NAV: vertical navigation
V/S: vertical speed
VERT: vertical
VERT: SPD vertical speed
VFY: verify
VG: vertical gyro
VHF: very high frequency
VIB: vibration
VLD: valid
VLV: valve
VOL: volume
VOLT: voltage
VOR VHF: omni range receiver
VOX: voice
VTR: video tape reproducer
W/D: wiring diagram
W/W: wheel well
WARN: warning
WG: wing
WHL: wheel
WHLS: wheels
WPT: waypoint
WSHLD: windshield
WX: weather
WXR: weather
X-CH: cross channel
X-CHAN: cross channel
XDCR: transducer
XMISSION: transmission
XMIT: transmit
XMTR: transmitter
XPNDR: transponder
Y/D: yaw damper
ATA 06 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
REFERENCE PLANES AND LINES
The airplane is divided into reference planes (stations), waterlines and buttock lines. These are measured in
inches from fixed points of reference. This provides a means of quickly identifying the location of components,
the center of gravity and the distribution of the weight.
Standard Abbreviations and Definitions
Fuselage
B STA, BS, or STA: Body (Fuselage) Station.
This is a plane perpendicular to the fuselage centerline, It is located 159.00 inches forward of the nose.
BBL or BL: Body (Fuselage) Buttock Line.
This is a vertical plane parallel to the fuselage vertical centerline plane, BBL 0.00 located by its distance
outboard from the fuselage centerline plane.
BRP: Body (Fuselage) Reference Plane.
This is a plane perpendicular to the BBL plane and passes through the top of the main deck floor beams
(BWL 208.10).
BWL or WL: Body (Fuselage) Waterline.
This is a plane perpendicular to the BBL plane. It is located by its distance from a parallel imaginary
plane (BWL 0.00). BWL 0.00 is 133.00 inches below the lowest fuselage surface.
LBL: Left Buttock Line
RBL: Right Buttock Line
Wing
FS:
The principal spanwise transverse member of the wing structure.
It is perpendicular to the wing reference plane.
ISS: Inboard Slat Stations.
These are planes perpendicular to inboard leading edge slats. They are measured from the intersection of
the slat rotation axis and a plane perpendicular to the wing reference plane.
LES: Leading Edge Station.
These are planes perpendicular to the wing reference plane and the leading edge. They are measured
from the intersection of the leading edge extension and the wing buttock line 0.00.
757 General Familiarization (7-2005)
ATA 06 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Wing Definitions (Continued):
MAC: Mean Aerodynamic Chord.
This is the chord of a section of an imaginary airfoil which would have vectors throughout the flight range
identical to those of the actual wing.
OSS: Outboard Slat Stations.
These are planes perpendicular to the outboard leading edge slats. They are measured from the
intersection of the slat rotation axis and a plane perpendicular to the wing reference plane.
RS:
See definition for FS.
W STA or WS: Wing Station.
These are planes perpendicular to the wing reference plane and the plane of the outboard rear spar. They
are measured from the intersection of the extended leading edge and wing buttock line 0.00.
WBL: Wing Buttock Line.
This is a plane perpendicular to the wing reference plane and parallel to the trace of the fuselage
centerline. It is measured from intersection of wing reference plane and body buttock line 0.00.
WRP: Wing Reference Plane.
This is the datum plane of the wing. It is inclined up 5 degrees with respect to the BWL plane and passes
through the intersection of the BBL 0.00 and BWL 178.187909.
WTS: Wing Tip Station.
This is a plane perpendicular to the wing reference plane and wing buttock line 0.00. It is measured from
the intersection of the leading edge and wing buttock line 0.00.
Vertical Stabilizer
ASS: Auxiliary Spar Station.
This is a plane perpendicular to the vertical stabilizer auxiliary spar. It is measured from the Auxiliary
Spar Station 0.00, intersection of the auxiliary spar centerline extension and body waterline 228.99 (757
ROOT CHORD).
757 General Familiarization (7-2005)
ATA 06 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Vertical Stabilizer (Continued):
FIN STA: Fin Station.
This is a plane perpendicular to the centerline of the vertical stabilizer rear spar. It is measured from Fin
Station 0.00, intersection of rear spar centerline extension and body waterline 228.99 (757 ROOT CHORD).
FSS: Front Spar Station.
This is a plane perpendicular to the vertical stabilizer front spar. It is measured from the fin front spar
station 0.00, intersection waterline 228.99 (757 ROOT CHORD).
LES: Leading Edge Station.
These are planes perpendicular to the vertical stabilizer leading edge. They are measured from the leading
Edge Station 0.00, intersection of the leading edge line extension and body waterline 228.99 (757 ROOT
CHORD).
LFFS: Lower Front Spar Station.
These are planes perpendicular to the vertical stabilizer lower front spar. They are measured from the
Lower Front Spar Station 0.00, intersection of the lower front spar centerline extension and body waterline
228.99 (757 ROOT CHORD).
RUD STA: Rudder Station.
These are planes perpendicular to the rudder hinge centerline. They are measured from Rudder Station
0.00, intersection of rudder hinge centerline and body waterline 228.99 (757 ROOT CHORD).
Horizontal Stabilizer
AUX SPAR STA: Auxiliary Spar Station.
These are planes perpendicular to the horizontal stabilizer auxiliary spar. They are measured from
Auxiliary Spar Station 0.00, intersection of auxiliary spar extension and stabilizer buttock line 0.00.
ELEV STA: Elevator Station.
These are planes perpendicular to the elevator hinge centerline. They are measured from the intersection
of elevator hinge centerline and stabilizer buttock line 0.00.
757 General Familiarization (7-2005)
ATA 06 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Horizontal Stabilizer (Continued):
FS STA: Front Spar Station.
These are planes perpendicular to the horizontal stabilizer front spar. They are measured from Front Spar
Station 0.00, intersection of front spar and trace of body buttock line 0.00 at horizontal stabilizer reference
plane.
HSBL: Stabilizer Buttock Line.
This is a plane perpendicular to the horizontal stabilizer reference plane and parallel to the trace of the
fuselage centerline. It is measured from stabilizer buttock line 0.00, intersection of horizontal stabilizer
reference plane and body buttock line 0.00.
HSRP: Horizontal Stabilizer Reference Plane.
This is the datum plane of the horizontal stabilizer. It is inclined 7° up with respect to the BWL plane and
passes through the intersection of the BBL 0.00 and BWL 238.015 planes.
LE STA: Leading Edge Station.
This is a plane perpendicular to the horizontal stabilizer leading edge. It is measured from Stabilizer
Leading Edge Station 0.00, intersection of leading edge line extension and stabilizer buttock line 0.00.
RS STA: Rear Spar Station.
This is a plane perpendicular to the horizontal stabilizer rear spar. It is measured from Rear Spar Station
0.00, intersection of rear spar and trace of body buttock line 0.00 at horizontal stabilizer reference plane.
STAB STA: Stabilizer Station. This is a plane perpendicular to the stabilizer rear spar and the horizontal
stabilizer reference plane. Stabilizer station 0.00 is at the intersection of the leading edge extension, body
buttock line 0.00 and the horizontal stabilizer reference plane.
Power Plant
PPBL: Power Plant Buttock Line.
This is a plane perpendicular to the wing reference plane. It is measured from a parallel plane (PPBL
0.00) that intersects the WBL 255.0 plane at the wing leading edge and angles 1.5 degrees inboard just
forward of the wing leading edge.
757 General Familiarization (7-2005)
ATA 06 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Power Plant (Continued):
PPWL: Power Plant Waterline.
This is a plane perpendicular to the PPBL datum plane and inclined 2.4072 degrees upward from the wing
reference plane. The PP WL 100.00 (centerline of engine) is measured 61.70 inches down from the wing
leading edge at WBL 255.00.
PPS: or PPSTA Power Plant Station.
This is a plane perpendicular to the engine centerline. The zero position is located 72.30 inches forward of
the forward edge of the fan cowl panel.
757 General Familiarization (7-2005)
ATA 06 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 06 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-1
757 General Familiarization (7-2005)
ATA 06 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-2
757 General Familiarization (7-2005)
ATA 06 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PRIMARY AIRCRAFT DIMENSIONS
IF 06-3
757 General Familiarization (7-2005)
ATA 06 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 06 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DIMENSIONS
Overall Airplane:
Length -- 155 feet-3 inches
Width -- 124 feet-6 inches
Height (vertical stabilizer tip, top of the fairing to the ground) -- 44 feet-6 inches
Wing:
Root Chord (calculated, at body centerline) -- 360.85 inches
Basic Chord (calculated) -- 286.50 inches
Tip Chord (calculated) -- 68.00 inches
Planform Taper Ratio
Tip Chord/Basic Chord -- 0.237
Tip Chord/Root Chord -- 0.188
Dihedral (wing reference plane in relation to the body reference plane) -- 5 degrees
Sweepback (25 percent chord line) -- 25 degrees
Aspect Ratio -- 7.95
Mean Aerodynamic Chord (basic wing only) -- 199.70 inches
Horizontal Stabilizer:
Span -- 600 inches
Taper Ratio -- 0.347
Sweepback (25 percent chord line) -- 30.186 degrees
Dihedral (horizontal stabilizer reference plane in relation to body reference plane) -- 7 degrees
Aspect Ratio -- 4.496
Vertical Stabilizer:
Height -- 293.374 inches
Taper Ratio -- 0.346
Sweepback (25 percent chord line) -- 40 degrees
Aspect Ratio -- 1.615
757 General Familiarization (7-2005)
ATA 06 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fuselage:
Height of body reference plane (top of floor beam WL 208.10) above ground at main gear -- 152.10 inches
Height (constant cross section) Above body reference plane -- 98.4 inches
Below body reference plane -- 75.10 inches
Height to centerline of windows above body reference plane -- 38 inches
Length -- 1846 inches
Areas:
Wing (basic) -- 1951 square feet
Horizontal Stabilizer Surfaces (total, includes the area within fuselage) -- 545 square feet
Vertical Stabilizer Surfaces (total) -- 370 square feet
757 General Familiarization (7-2005)
ATA 06 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-4
757 General Familiarization (7-2005)
ATA 06 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
BODY STATION DIAGRAM
IF 06-5
757 General Familiarization (7-2005)
ATA 06 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-6
757 General Familiarization (7-2005)
ATA 06 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
VERTICAL STABILIZER AND RUDDER STATION DIAGRAM
IF 06-7
757 General Familiarization (7-2005)
ATA 06 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HORIZONTAL STABILIZER AND ELEVATOR STATION DIAGRAM
IF 06-8
757 General Familiarization (7-2005)
ATA 06 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WING STATION DIAGRAM
IF 06-9
757 General Familiarization (7-2005)
ATA 06 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE AND NACELLE STATION DIAGRAM
IF 06-10
757 General Familiarization (7-2005)
ATA 06 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ZONE DIAGRAMS
The 757 airplane is divided into 8 major zones to help you find and identify the airplane components and parts.
The major zones then are divided into the sub-zones and the sub-zones into zones.
The zones are numbered in the sequence that follows:
1.
2.
3.
4.
Wings - inboard to outboard and front to back
Horizontal Stabilizer and Elevator - inboard to outboard and front to back
Vertical Stabilizer and Rudder - root to tip of vertical stabilizer
Fuselage - front to back and away from floor line
Each of the structural components, passenger compartment doors, cargo doors, landing gear doors, rudders,
elevators, flaps, ailerons, spoilers, leading edge devices, and equivalent components has a different zone
number.
A three-digit number identifies the major zones, sub-zones, and zones as follows:
1. Major Zone - the first digit is a number from 1 to 8 followed by two zeroes
2. Sub-zone - the first digit represents the major zone, the second digit is a number from 1 to 6 or 9, and
the third digit is a zero
3. Zone - the first two digits represent the sub-zone number and the third digit shows a component or
group of components that are in the sub-zone
Major Zones
100
200
300
400
500
600
700
800
Lower Half of Fuselage
Upper Half of Fuselage
Empennage and Body Section 48
Power Plants and Nacelle Struts
Left Wing
Right Wing
Landing Gear and Landing Gear Doors
Doors
757 General Familiarization (7-2005)
ATA 06 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-11
757 General Familiarization (7-2005)
ATA 06 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
SERVICE INTERPHONE SYSTEM
The service interphone system provides facilities for interphone communication between servicing stations
during ground operations. In addition, communication is extended to the flight compartment and attendant
stations.
The service interphone system includes phone jacks located at convenient service locations around the airplane.
The system also includes amplifiers and mixing circuits that are housed inside the audio accessory unit. It is
located on shelf E4 of the main equipment center. The SERV INTPH switch activates the system and is located
on the right sidewall panel, P61.
The system gets power from the 28-volts DC battery bus, through a circuit breaker on overhead panel P11.
Component Details
Audio Amplifier
Two identical audio amplifiers with parallel inputs are located in the audio accessory unit on shelf E4-3 in the
main equipment center. The amplifiers are used by both the service interphone and the cabin interphone
systems. A placard on the audio accessory unit identifies and locates the two amplifiers. There are two outputs
from each amplifier. One output goes to the cabin interphone audio on the audio selector panels. The other
output goes to the cabin interphone handsets and service interphone jacks. The amplifiers have internal
adjustments preset for normal compression, squelch, and volume.
Service Interphone Switch
The SERV INTPH switch is located on right sidewall panel P61. In the ON position, the switch connects the
microphone lines from the service interphone jacks to the input of the interphone amplifiers. The OFF position
disconnects the microphone lines to isolate the service interphone jacks during flight.
757 General Familiarization (7-2005)
ATA 06 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-12
757 General Familiarization (7-2005)
ATA 06 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Operation
The INTERPHONE CABIN SERVICE circuit breaker on overhead panel P11 controls power to the system.
The SERV INTPH switch on panel P61 connects the mic lines from the phone jacks to the input of the interphone
amplifiers. This switch activates the service interphone system. Then the interphone amplifiers distribute the
amplified audio to the audio selector panels in the flight compartment, as well as to the cabin handsets, and to
the audio line of the service interphone jacks located at convenient service locations. All microphone inputs from
the handsets and from the audio selector panels mixes with that of the service interphone jacks onto a singleparty system.
The flight crew talks with the service interphone stations by pushing the MIC SELECTOR labeled CAB on the
audio selector panels. This switch, along with the SERV INTPH switch on panel P61, connects a flight crew
headset to the service interphone system. On the audio selector panel, the CAB switch is a push-on/push-off
type and is lighted when on. The same knob also turns a volume control for that particular station.
Personnel at the cabin attendant stations use the service interphone system by picking up the attendant handset
and talking.
Personnel at service locations use the service interphone system by plugging a headset into the phone jack and
talking. First, the SERV INTPH switch must be set to ON.
Control
To place the system in operation, supply electrical power. On overhead panel P11, make sure the INTERPHONE
CABIN SERVICE circuit breaker is closed. On right sidewall panel P61, make sure the SERV INTPH switch is on.
757 General Familiarization (7-2005)
ATA 06 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-13
757 General Familiarization (7-2005)
ATA 06 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Cabin Interphone System
The cabin interphone system provides facilities for interphone communication among cabin attendants, and
between the flight compartment crewmembers and attendants. The cabin interphone system microphone circuits
can be switched to the input of the passenger address (PA) system to permit announcements.
Cabin Interphone Handsets
Each attendant station has a handset. Each handset has the following components: a noise-canceling microphone
to decrease the input of airplane noise, a small speaker to listen, buttons to operate the handset, and a magnetoperated hook switch to disconnect the handset from the system when the attendant places the handset back on
the hook.
In operation, 4 of the handset buttons call specific stations (FWD, MID, AFT, PILOT). The P.A. button connects
the handset to the passenger address speakers to make announcements. The RESET button permits the
attendant to make more calls without the need to put the handset back on the hook after each call. The button
operates the same as when the handset is put on the hook. The ALERT button connects the handset to all
stations at the same time. The P.A. PUSH TO TALK switch momentarily connects the microphone to transmit
voice during passenger announcements.
Pilot’s Call Panel
The pilots' call panel is on overhead panel P5. The cabin interphone system connects to the pilots' call panel
through three blue lighted call switches which indicate an attendant call (FWD, MID, AFT), and one blue lighted
call switch that indicates an alert call (ALERT).
The call switches (FWD, MID, AFT) are used to signal the related attendant station. The ALERT call switch is
used to make an alert call.
Operation
Cabin interphone calling from one station to another is done by single-digit dialing. Pressing a call switch on a
handset activates the system by sending two tones to the audio accessory unit. The tones are processed in the
accessory unit and a signal is provided to the called station to turn on the call light. A signal is also sent to the
PA system to generate a hi/lo chime for attendant calls or to the aural warning system to generate a hi chime in
the flight compartment for pilot calls.
The FWD, MID, AFT, and ALERT call switches on the pilots' call panel make it possible for the pilot to make and
receive cabin interphone calls. The light in each call switch comes on when there is an incoming call and go off
when the pilot pushes the switch. Power failures will automatically connect the microphones of the cabin
interphone handsets to the PA system.
757 General Familiarization (7-2005)
ATA 06 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 06-14
757 General Familiarization (7-2005)
ATA 06 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Operation (Continued):
The controller in the audio accessory unit monitors busy stations, and prevents incoming calls when a called
station is busy. The controller also activates the PA hi/lo chime output, and causes the lamp driver to turn on
the attendant call light. Attendant to attendant calling can be accomplished except under the following
conditions:
1. Calls can not be placed between the two mid attendant stations
2. Any station that has a handset off hook cannot be called
3. No station can call another station that has been called but not answered; however, the station that
placed the call can repeat the call
4. Only one call can be made from a handset without first performing a reset
Any attendant can call the pilot. Multiple calls by attendants to the pilot are possible. After a call, the attendant
must put the handset back on the hook (cradle) or push the RESET switch to reset the unit. When the attendant
at the called station removes the handset from the hook, a magnetic switch closes and connects the microphone
to the system.
Control
To place the system in operation, supply electrical power. On overhead panel P11, make sure the INTERPHONE
CABIN SERVICE circuit breaker is closed.
757 General Familiarization (7-2005)
ATA 06 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOTES:
757 General Familiarization (7-2005)
ATA 06 - Page 43
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 21
AIR-CONDITIONING SYSTEM .............................................................................................................. 4
Cooling ............................................................................................................................................... 4
Temperature Control ........................................................................................................................... 4
Distribution ......................................................................................................................................... 4
Heating ................................................................................................................................................ 4
Pressurization Control......................................................................................................................... 4
Control and Indication ........................................................................................................................ 6
COOLING PACK SYSTEM .................................................................................................................... 11
Operation .......................................................................................................................................... 14
The Air Cooling Pack Operation ...................................................................................................... 14
The Pack Flow Control ..................................................................................................................... 17
The Pack Temperature Control ......................................................................................................... 18
The AUTO Mode .............................................................................................................................. 18
The STBY-N mode ........................................................................................................................... 19
The STBY-W Mode.......................................................................................................................... 19
The STBY-C Mode........................................................................................................................... 19
The OFF Mode ................................................................................................................................. 19
AIR-CONDITIONING SYSTEM COMPONENTS ................................................................................ 23
Flow Control and Shutoff Valve....................................................................................................... 23
Primary Heat Exchanger ................................................................................................................... 24
Secondary Heat Exchanger ............................................................................................................... 24
Plenum/Diffuser Assembly ............................................................................................................... 24
Air Cycle Machine ............................................................................................................................ 26
Re-heater ........................................................................................................................................... 28
Condenser ......................................................................................................................................... 28
Water Extractor ................................................................................................................................. 30
Split-Duct Water Separator............................................................................................................... 30
Low Limit Control Valve ................................................................................................................. 32
Temperature Control Valve .............................................................................................................. 33
Automatic-Pack-Temperature Controller ......................................................................................... 34
Standby-Pack-Temperature Controller ............................................................................................. 34
Pack Temperature Sensor ................................................................................................................. 37
Altitude Switches .............................................................................................................................. 37
Compressor Outlet Temperature Sensor ........................................................................................... 37
Compressor Outlet Overheat Switch ................................................................................................ 38
Pack Outlet Overheat Switch ............................................................................................................ 38
Cabin Air Supply Check Valve ........................................................................................................ 38
Flow Control Card ............................................................................................................................ 40
ZONE TEMPERATURE CONTROL SYSTEM ..................................................................................... 41
Functional Description.............................................................................................................................. 44
System Temperature Integration (Pack Demand) ............................................................................. 44
757 General Familiarization (7-2005)
ATA 21 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Alternate Temperature Control ......................................................................................................... 44
APU Output Control ......................................................................................................................... 44
Zone Temperature Control (Auto Mode).......................................................................................... 45
Off Mode........................................................................................................................................... 46
Trim Air Pressure Regulation ........................................................................................................... 46
ZONE TEMPERATURE COMPONENTS.............................................................................................. 48
Zone Temperature Sensors ............................................................................................................... 48
Duct Temperature Sensors ................................................................................................................ 48
Duct Temperature Sensors for the Zone Inlet................................................................................... 48
Duct Temperature Sensors for the Mix Manifold............................................................................. 48
Duct Overheat Switches.................................................................................................................... 48
Check Valves for the Trim Air Supply ............................................................................................. 50
Zone Temperature Selectors ............................................................................................................. 50
Pressure Regulating Valve for the Trim Air System ........................................................................ 52
The Modulating Valve for the Trim Air ........................................................................................... 53
Zone Temperature Controller ........................................................................................................... 54
CONDITIONED AIR DISTRIBUTION SYSTEM ................................................................................. 55
Mix Manifold .................................................................................................................................... 56
Ground Conditioned Air Connector ................................................................................................. 56
Flight Deck Conditioned Air Distribution ........................................................................................ 59
Floor Registers .................................................................................................................................. 59
Overhead Register............................................................................................................................. 59
Windshield Diffuser .......................................................................................................................... 59
Gaspers.............................................................................................................................................. 60
Air Inlets ........................................................................................................................................... 60
Passenger Cabin Conditioned Air Distribution ................................................................................ 62
Sidewall Riser Ducts......................................................................................................................... 62
Overhead Duct .................................................................................................................................. 62
Sidewall Outlet Ducts ....................................................................................................................... 62
Return Air Grills ............................................................................................................................... 62
Overhead Outlets .............................................................................................................................. 62
Sidewall Outlets ................................................................................................................................ 62
Galley and Lavatory Outlets ............................................................................................................. 62
CABIN AIR RECIRCULATION SYSTEM ............................................................................................ 64
Control Panel .................................................................................................................................... 67
Failure Indication .............................................................................................................................. 67
Recirculation Air Fan........................................................................................................................ 68
Recirculation Air Filters ................................................................................................................... 68
Recirculation Air Check Valve ......................................................................................................... 68
THE LAVATORY AND GALLEY VENTILATION SYSTEM ............................................................ 71
Ventilation ........................................................................................................................................ 71
Functional Description...................................................................................................................... 71
Failure Indication .............................................................................................................................. 71
Aft Equipment, Lavatory, Galley Ventilation Fan ........................................................................... 72
Lavatory and Galley Ventilation Check Valve ................................................................................. 72
Galley Ventilation Filter ................................................................................................................... 72
757 General Familiarization (7-2005)
ATA 21 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fan Current Sensors .......................................................................................................................... 72
HEATING SYSTEM - FORWARD & AFT CARGO COMPARTMENTS & SUPPLEMENTAL ....... 74
FORWARD CARGO COMPARTMENT HEATING SYSTEM .................................................... 74
Functional Description...................................................................................................................... 74
Forward Cargo Heat Exchanger (only on some aircraft) .................................................................. 76
Forward Cargo Heating Flow Fan .................................................................................................... 76
Forward Cargo Temperature Thermal Switch .................................................................................. 76
Forward Cargo Fan Current Sensor .................................................................................................. 76
Forward Cargo Heating Inlet Grill.................................................................................................... 76
AFT CARGO COMPARTMENT HEATING SYSTEM ......................................................................... 78
Functional Operation ........................................................................................................................ 78
Aft Cargo Heater............................................................................................................................... 80
Aft Cargo Heating Flow Fan ............................................................................................................ 80
Aft Cargo Temperature Control Switch............................................................................................ 80
Aft Cargo Fan Current Sensor .......................................................................................................... 80
Aft Cargo Heating Inlet Grill ............................................................................................................ 80
SUPPLEMENTAL HEATING SYSTEM ................................................................................................ 82
Foot Electric Surface Heaters ........................................................................................................... 82
Shoulder Air Supply Heater.............................................................................................................. 82
PRESSURIZATION SYSTEM ................................................................................................................ 85
Functional Description...................................................................................................................... 86
Automatic Pressure Controller.......................................................................................................... 88
Outflow Valve................................................................................................................................... 89
Cabin Pressure Selector Panel .......................................................................................................... 90
PRESSURE RELIEF SYSTEM ............................................................................................................... 92
Positive Pressure Relief Valve.......................................................................................................... 92
Vacuum Relief Valve........................................................................................................................ 94
PRESSURIZATION INDICATION AND WARNING SYSTEM.......................................................... 95
Altitude Warning System ................................................................................................................. 95
Altitude Switch ................................................................................................................................. 95
Differential Pressure Sensor ............................................................................................................. 96
Pressurization and Indication Warning Module ............................................................................... 97
EQUIPMENT COOLING SYSTEM ....................................................................................................... 99
Functional Description – Equipment Cooling System Normal Operation ..................................... 100
Fan Check Valve............................................................................................................................. 102
Overboard Exhaust Valve ............................................................................................................... 103
Equipment Cooling Air Fan............................................................................................................ 104
Equipment Cooling Air Cleaner ..................................................................................................... 105
Fan Fail Sensor (Current Sensor).................................................................................................... 106
Equipment Cooling Low Flow Sensor ........................................................................................... 107
Equipment Cooling Smoke Sensor ................................................................................................. 108
Equipment Cooling Control Card ................................................................................................... 109
757 General Familiarization (7-2005)
ATA 21 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AIR-CONDITIONING SYSTEM
The air-conditioning system maintains airplane environmental control for the comfort of passengers and crew. The
total system is made up of the conditioned air distribution system, pressurization control system, cargo
compartment heating system, cooling system and temperature control system.
Cooling
Cooled air is supplied to the distribution system by two air cooling packs. The pneumatic system provides source
air to the packs. The ram air system assures pack temperature control. The pack indicating system allows flight
deck monitoring of pack operation.
The equipment cooling system removes heat generated by electrical and electronic equipment in the flight
compartment, forward equipment area, main equipment racks, and aft equipment area.
Temperature Control
The temperature control system regulates the environment within the airplane's three primary zones. Cooled air
mixes with hot "trim" air to obtain the desired temperature at each zone. Monitoring of the system is provided by
valve position indication and the zone temperature indication system.
Distribution
The distribution system channels temperature-controlled air to the flight deck and passenger cabin through a
network of distribution ducts. The main mix manifold combines cooled air from the air cooling packs with
recirculation air drawn from the passenger cabin. The air is channeled to the flight deck through ducts and to the
passenger compartment through risers, overhead ducting and outlets. Separate exhaust ducting provides positive
ventilation for all lavatory and galley areas.
Heating
Independent cargo compartment heating systems provide temperature control for the forward, aft and bulk cargo
compartments.
The supplemental heating system channels heated air to the captain's and first officer's foot and shoulder areas.
Pressurization Control
The air conditioning system provides pressurization control by regulating the amount of air discharged from the
airplane. Backup positive and negative pressurization relief systems prevent the cabin pressure from exceeding
established limits. Pressurization indication and warning systems allow the operator to monitor the pressurization
control system.
757 General Familiarization (7-2005)
ATA 21 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-1
757 General Familiarization (7-2005)
ATA 21 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Control and Indication
Control selectors and switch lights on the pilot's overhead panel P5, and one control selector on the right side
panel P61, allow control of all air-conditioning systems. Warning lights, position indicators and other gages on the
P5 panel provide partial indication of system operation. The Engine Indication and Crew Alerting System provide
the remaining indication for the air conditioning system.
EICAS provisions include two cathode-ray tube (CRT) display screens on the pilot's center instrument panel P2, a
display select panel on the forward electrical control stand P9, and an EICAS maintenance panel on the right side
panel P61. EICAS displays primary engine parameters on the upper CRT. Warning, caution, and advisory messages
also appear on the upper CRT during certain system malfunctions. Aural tones occur in the flight deck with each of
the above types of messages.
Two other types of messages may also appear; status or maintenance. These two types of messages both appear
on the lower CRT only when called upon by an operator. Status messages display data on the current status of
operation and may be called upon by selecting the STATUS switch on the display select panel. The airplane may
be either in-flight or on the ground. Maintenance messages are available only on the ground by selecting the
ECS/MSG switch on the EICAS MAINT panel. Current maintenance information is usually displayed unless a
recorded page is called up. A recorded page is a page with information about the status of the system in-flight
when certain failures occurred. These are called AUTO EVENTS and are called up by selecting the AUTO EVENT
READ switch on the EICAS MAINT panel.
757 General Familiarization (7-2005)
ATA 21 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-2
757 General Familiarization (7-2005)
ATA 21 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-3
757 General Familiarization (7-2005)
ATA 21 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EICAS ECS MAINTENANCE PAGE
IF 21-4
757 General Familiarization (7-2005)
ATA 21 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COOLING PACK SYSTEM
The cooling pack system is used to decrease the temperature of a supply of air. The output of the cooling pack
system is cold air. This cold air is supplied to all of the necessary airplane compartments by the conditioned-airdistribution systems.
Other systems are necessary for the cooling pack system to operate. The cooling pack system uses a supply of air
from the pneumatic system. This is hot air from the engines, the APU, or a ground air source.
The cooling pack system also uses a supply of air from the ram air system. This is cold air which does not enter
the air cooling packs. It is used to decrease the temperature of the air in two of the heat exchangers in each air
cooling pack.
The temperature control system sends a signal for the necessary output of the cooling pack system.
The cooling pack system is made of these components:
1. Two air cooling packs; each has these components:
a)
b)
c)
d)
e)
f)
g)
h)
i)
j)
k)
l)
m)
2.
3.
4.
5.
6.
7.
8.
A primary heat exchanger
A secondary heat exchanger
An air cycle machine (which has a turbine, a compressor, and a fan)
A split-duct water separator
A re-heater
A condenser
A water extractor
A low-limit control valve
A temperature control valve
A compressor-outlet overheat switch
Two compressor-outlet temperature sensors
Two pack temperature sensors
A plenum/diffuser assembly
Two automatic pack-temperature controllers
One standby pack-temperature controller
Two flow control cards
Two flow-control-and-shutoff valves
Two pack-outlet overheat switches
Two cabin-air-supply check valves
Two altitude switches.
Most of the components in the cooling pack system are interchangeable between the left and right systems. The
only component that is not interchangeable is the standby pack temperature controller, because there is only one of
them on the airplane.
757 General Familiarization (7-2005)
ATA 21 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-5
757 General Familiarization (7-2005)
ATA 21 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-6
757 General Familiarization (7-2005)
ATA 21 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Operation
The cooling pack system puts out a specified amount of air at a specified temperature.
The Functional Description of the cooling pack system is divided into three parts:
1. The Air Cooling Pack Operation tells what happens to the air as it flows through the air cooling pack.
2. The Pack Flow Control tells how the cooling pack system gets the necessary amount of air from the air
cooling pack.
3. The Pack Temperature Control tells how the cooling pack system gets the necessary air temperature from
the air cooling pack.
The Air Cooling Pack Operation
The air cooling pack is pneumatically operated. It operates only when a supply of air goes into it. The flow-controland- shutoff valve lets air enter the air cooling pack. Thus, when the flow-control-and-shutoff valve is open, the air
cooling pack is ON. When the flow-control-and-shutoff valve is closed, the air cooling pack is OFF.
Air from the flow-control-and-shutoff valve first flows into the primary heat exchanger. The primary heat exchanger
causes the air temperature to decrease. Then the pressure and the temperature of the air are increased as it flows
through the air-cycle-machine compressor. The secondary heat exchanger again causes the air temperature to
decrease.
Then the water is removed from the air. The split-duct water-separator removes an initial amount of water as the
air moves from the secondary heat exchanger. Then the air flows through the re-heater and the condenser, which
both cause the air temperature to decrease. The low temperature of the air causes water drops. Subsequently, as
the air flows through the water extractor, the water extractor removes the water. The water extractor puts the water
in the ram air system or lets it flow overboard.
After the water is removed from the air, the air flows in the other direction through the re-heater. This causes the
temperature to increase before it goes into the air-cycle- machine turbine.
The air-cycle-machine turbine decreases the pressure and the temperature of the air. The energy from the
expanded air makes the air cycle machine operate. Subsequently, the air again flows through the condenser. This
causes the air temperature to increase as it goes from the air cooling pack.
The low-limit-control valve and the temperature control valve lets some of the air go around the air cycle machine.
This hot air flows to the turbine outlet to prevent ice. Also, because a smaller amount of air flows through the air
cycle machine, the speed of the air cycle machine decreases. This causes the compressor outlet temperature to
decrease and the air- cooling-pack-outlet temperature to increase.
757 General Familiarization (7-2005)
ATA 21 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-7
757 General Familiarization (7-2005)
ATA 21 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Cooling Pack Operation (Continued):
Small amounts of air flow through the tubes in the air cooling pack. This air has special uses. Some tubes supply
air to the low-limit-control valve to cause it to open or close. Some air goes from the secondary heat exchanger
through a tube to get water from the water extractor. Also, some air goes from the compressor flows to the
condenser and the turbine outlet to prevent ice.
The Pack Flow Control
The amount of air that flows through the air cooling pack is controlled by the flow control cards. Each flow control
card controls one flow-control-and-shutoff valve. The flow-control- and-shutoff valves permit a controlled amount
of air into each air cooling pack.
The flow-control-and-shutoff valves open when the L or R PACK selector, on the pilot's overhead panel P5, is not
in the OFF position. The selector sends a signal to the flow control card to open the flow-control-and-shutoff valve.
The flow control card causes the flow-control-and-shutoff valves to go into the normal flow mode or the high flow
mode.
A compressor overheat condition occurs when the compressor discharge temperature is more than 490°F (254°C).
At 490°F (254°C), the compressor outlet overheat switch closes and sends an overheat signal to the flow control
card. Also at 490°F (254°C), the compressor outlet sensor sends an overheat signal through the pack temperature
controller to the flow control card. The flow control card then latches in the overheat position and causes the
following conditions to occur:
1.
2.
3.
4.
The flow-control-and-shutoff valve closes.
The PACK OFF and INOP lights, on the P5 panel, come on.
The advisory messages L (R) PACK OFF and L (R) PACK TEMP show on the EICAS display.
The AUTO/MANUAL relays are not energized (this prevents control of the air cooling pack by the
automatic-pack- temperature controller).
5. The backup-temperature-control card causes the air cooling pack to go to the cool mode (the ram air
doors and the temperature control valve open).
A pack overheat condition occurs when the pack discharge temperature is more than 190°F (88°C). At 190°F
(88°C), the pack overheat switch closes and sends an overheat signal to the flow control card. The flow control
card then latches in the overheat position and causes the following conditions to occur:
1. The backup temperature control card causes the pack to go the full cold mode (the ram air doors fully
open and the temperature control valve closes).
2. The AUTO/MANUAL relays are not energized (this prevents control of the air cooling pack by the
automatic-pack- temperature controller).
3. The INOP light, on the P5 panel, comes on.
4. The advisory messages L (R) PACK TEMP show on the EICAS display.
The air cooling packs operate as usual when the temperature returns to the normal range and the RESET switch,
on the P5 panel, is pushed.
757 General Familiarization (7-2005)
ATA 21 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
The Pack Temperature Control
The air-cooling-pack-outlet temperature is controlled by the automatic-pack-temperature controllers, the standbypack- temperature controller, or the backup-temperature-control cards. Each component controls the air cooling
packs for a different temperature mode. The temperature mode is controlled by the position of the L and R PACK
selectors, on the P5 panel (the L and R PACK selectors can be in different temperature modes at the same time).
The temperature modes are AUTO, STBY-N, STBY-W, STBY-C, and OFF.
The AUTO Mode
1. In the AUTO mode, the automatic-pack-temperature- controllers control the air cooling packs. AC power is
always supplied to each controller. When the L or R PACK selector is in AUTO, the controller sends signals
through the AUTO/MANUAL relays to control the position of the ram air doors and the temperature control
valve. Also, because the low-limit-control valve relay is not energized, the controller can send signals to
the low-limit-control valve.
2. The necessary air-cooling-pack-outlet temperature is sent as a signal to the automatic pack temperature
controller by the zone temperature controller. This is the pack demand signal. The automatic-packtemperature controller compares the pack demand signal to the pack- temperature-sensor signal. The
difference between the signals causes an error. The error controls the signal that is sent to the ram air
door actuators and the temperature control valve.
3. When the airplane is on the ground, the ram air doors stay fully open. The temperature control valve
position changes to control the air-cooling-pack-outlet temperature.
4. The compressor outlet temperature sensor supplies input to the controller to prevent an overheat
condition.
5. The altitude switch supplies input to the controller to change the lower water separator temperature limit
from 35°F (1.7°C) to 0°F (-18°C). The controller closes at 31,000 feet. This sends a signal to the controller
to change the limit.
757 General Familiarization (7-2005)
ATA 21 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
The STBY-N mode
1. In the STBY-N mode, the backup-temperature-control card and the standby-pack-temperature-controller
controls the air cooling packs. When the L or R PACK selector is in STBY-N, the AUTO/MANUAL relays
are not energized. This lets AC power flow to the controller. It also supplies 28-volts DC to energize the
low-limit-control- valve-control relay (ground is supplied in the backup- temperature-control card). The
controller sends an open or a close signal through the control relay to the low-limit-control valve. The
backup-temperature-control card causes the ram air doors to be fully open and the temperature control
valve to be fully closed. Also, no current is supplied to the torque motor of the flow-control-and-shutoff
valve.
2. The demand signals for the standby-pack-temperature controller are programmed to be 40°F (4.5°C) (for
the pack temperature) and 450°F (232°C) (for the compressor outlet temperature). The controller
compares the temperature sensor signals to the set demand signals. The difference of the signals causes
two error signals. The greatest error signal controls the position of the low limit control valve.
The STBY-W Mode
1. In the STBY-W mode, the backup-temperature-control card controls the air cooling packs. 28-volts DC is
always supplied to the backup-temperature-control card. When the L or R PACK selector is in STBY-W, the
AUTO/MANUAL relays are not energized. This lets the card send signals through the AUTO/MANUAL
relays to the ram air doors and the temperature control valve.
2. In the STBY-W mode, the backup-temperature-control card causes the ram air doors and the temperature
control valve to be fully open. The low-limit-control valve stays closed unless it is pneumatically opened.
The STBY-C Mode
1. In the STBY-C mode, the backup-temperature-control card controls the air cooling packs. 28-volts DC is
always supplied to the backup-temperature-control card. When the L or R PACK selector is in STBY-C, the
AUTO/MANUAL relays are not energized. This lets the card send signals through the AUTO/MANUAL
relays to the ram air doors and the temperature control valve.
2. In the STBY-C mode, the backup-temperature-control card causes the ram air doors to be fully open and
the temperature control valve to be fully closed. The low- limit-control valve stays closed unless it is
pneumatically opened.
The OFF Mode
1. When the L or R PACK selector is OFF, the ram air inlet doors and the temperature control valve is
controlled by the air/ground relays. When the airplane is on the ground, the ram air doors will be fully
open and the temperature control valve will be fully closed. When the airplane is in the air, the ram air
doors will be fully closed and the temperature control valve will be fully open. The low-limit-control valve
will stay closed unless it is pneumatically opened.
2. Whenever the left or right flow control valve is closed, the applicable (L or R) PACK OFF light, on the P5
panel, will come on and the EICAS message L or R PACK OFF will be shown on the upper display.
757 General Familiarization (7-2005)
ATA 21 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-8
757 General Familiarization (7-2005)
ATA 21 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-9
757 General Familiarization (7-2005)
ATA 21 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AIR-CONDITIONING SYSTEM COMPONENTS
Flow Control and Shutoff Valve
The flow-control-and-shutoff valve lets a controlled amount of air flow from the pneumatic system into the air
cooling pack. There are two valves, one for each air cooling pack. Each valve is installed forward and outboard of
its air cooling pack.
The flow-control-and-shutoff valve lets air into the air cooling pack at two different flow rates. Each cooling pack
can supply about 80 percent of the air that is usually supplied by two cooling packs.
The mechanical-manual-override valve supplies a way for the flow control valve to be manually closed.
IF 21-10
757 General Familiarization (7-2005)
ATA 21 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Primary Heat Exchanger
The primary heat exchanger is a plate-fin, single-pass, cross-flow, air-to-air heat exchanger.
Hot air from the flow-control-and-shutoff valve enters the top of the primary heat exchanger. It flows first to the aft
end of the heat exchanger. Then it flows forward, through some of S-turns, to the heat exchanger outlet. Cold air
flows at a right angle to the hot air. The air temperatures change as the heat transmits through the thin plate fins.
An access door on each side of the heat exchanger supplies an access to clean it.
Secondary Heat Exchanger
The secondary heat exchanger is almost the same as the primary heat exchanger. One difference is that the
secondary heat exchanger is wider. Also, it has a heavy fin on the cold air inlet side, to prevent damage from hail.
Hot air for the secondary heat exchanger comes from the compressor outlet on the air cycle machine.
Plenum/Diffuser Assembly
The plenum/diffuser assembly has two main parts; the plenum and the diffuser.
The fan in the air cycle machine and a space around the fan are in the diffuser. The plenum, which is the outer
part, is made of fiberglass. An access door in the plenum supplies access to clean the primary heat exchanger.
The plenum/diffuser assembly connects the air cycle machine, the primary heat exchanger, and the ram air outlet
duct. It lets air flow from the primary heat exchanger outlet to the ram air outlet door. The plenum makes the air
from the primary-heat-exchanger outlet flow to the diffuser inlet. The air-cycle-machine fan blows air through the
diffuser to the ram air exhaust duct. The space around the fan, when necessary, lets more air flow through the
diffuser. The plenum and the space are made so that air cannot re-circulate back into the plenum.
757 General Familiarization (7-2005)
ATA 21 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-11
757 General Familiarization (7-2005)
ATA 21 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Cycle Machine
The air cycle machine is divided into three parts; a fan, a compressor, and a turbine. Each part has a different
impeller. All three impellers are attached to the same shaft. The shaft is held in its position by three air bearings.
Lubrication of the air bearings is not necessary.
Each part of the air cycle machine has its own air inlet and outlet. The turbine also has an extra inlet and three
tubes connected to it.
The air-cycle-machine fan is used to draw air through the primary and secondary heat exchangers. When the
airplane is in flight, air flows through the heat exchangers because the ram air inlet door is open. When the
airplane is not in flight, air will only flow through the heat exchangers when the fan pulls it through.
The air-cycle-machine compressor increases the temperature and the pressure of air from the primary heat
exchanger. The outlet from the compressor flows through a water-removal process. The water is removed more
easily because of the high air pressure.
The air-cycle-machine turbine causes the shaft to turn. The turbine decreases the temperature and the pressure of
the air. The low temperature at the turbine outlet may cause ice on the turbine outlet. To prevent ice or to
decrease the air temperature in the cooling pack, hot air is supplied to the turbine outlet. The supply of hot air is
controlled by the low-limit-control valve and the temperature control valve.
One of the tubes connected to the turbine also supplies hot air from the compressor outlet. The second tube
supplies the turbine outlet pressure to the low limit control valve. The third tube supplies cold air to the water
extractor inlet.
757 General Familiarization (7-2005)
ATA 21 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-12
757 General Familiarization (7-2005)
ATA 21 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Re-heater
The re-heater is a plate-fin, single-pass, cross-flow, air-to-air heat exchanger. The plate-fins are at right angles to
each other.
The re-heater decreases the air that flows from the secondary heat exchanger to the condenser. It also increases
the temperature of the air that flows from the water extractor to the turbine. The temperatures are changed in the
re-heater as the heat of the air is transmitted through the thin plate-fins.
Condenser
The condenser is a plate-fin, single-pass, cross-flow, air-to-air heat exchanger. The core has two heat transfer
matrices, which are separated by a space. The cold air side of the core also has tubes of hot air. Each heat
transfer matrix is brazed together as a unit. The housing assembly has two inlets and two outlets. It also has two
tubes connected to it. The housing assembly is welded to the core. All the parts of the condenser are made of
aluminum alloys.
The condenser changes the temperature of two supplies of air. The heat of the air supplies is transmitted through
the thin plate-fins.
The temperature of the air that flows from the re-heater to the water extractor is decreased. This causes the water
in the air to condense so it can be easily removed in the water extractor.
The temperature of the air that flows from the turbine outlet to the pack outlet is increased. This causes the air to
be above the temperature that would make ice.
One of the tubes is connected to the hot air inlet of the condenser. The tube supplies hot air from the compressor
outlet to the hot air tubes on the cold air side of the condenser. The hot air tubes prevent ice on the cold air inlet
of the condenser.
The other tube is connected to the cold air outlet of the condenser. This tube supplies air pressure to the lowlimit-control valve.
757 General Familiarization (7-2005)
ATA 21 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-13
757 General Familiarization (7-2005)
ATA 21 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Water Extractor
The water extractor has two main parts; the inlet duct and the collector. The inlet duct has a set of helical vanes,
brazed to the inner wall of the duct. The helical vanes do not move. The collector has a perforated inner shell, an
outer shell, and a sump. The collector isolates the water from the air and collects the water in the sump. Two
tubes are connected to the outer shell. A drain tube is connected to the sump. All of the parts of the water
extractor are made of aluminum alloys.
The air from the condenser has many small droplets of water. As the air flows into the water extractor, the helical
vanes cause the air to turn. As the air turns, the centrifugal force pushes the water out to the duct wall (the water
is pushed farther out than the air because the water is heavier). In the collector part of the water extractor, the
inner shell isolates the air/water near the duct wall. The air can flow through the perforations in the inner shell,
but the water cannot. Thus, the water flows through the outer shell and into the sump.
The water in the sump then flows to the water spray nozzle of the ram air cooling system. If the sump becomes
full, the water flows overboard through one of the ports to the outer shell. The other port connected to the outer
shell supplies water to the sump from the split-duct water separator.
Split-Duct Water Separator
The split-duct water separator removes water from the air that flows from the secondary heat exchanger to the reheater. The water that collects in the split-duct water separator is supplied to the sump in the water extractor.
757 General Familiarization (7-2005)
ATA 21 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-14
757 General Familiarization (7-2005)
ATA 21 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Low Limit Control Valve
The low-limit-control valve controls the flow of hot air from the primary heat exchanger inlet to the air-cyclemachine turbine outlet. Air that flows through the low limit valve does not go through the air cycle machine. This
air is used to prevent ice in the turbine outlet and to prevent an overheat condition of the air cycle machine.
The low-limit-control valve is a 2-inch diameter, pneumatically-controlled, butterfly valve. It has two differential
pressure servo assemblies, an electromagnetic control valve, a pneumatic actuator, a reference pressure regulator,
a valve flow section, and a valve position indicator.
IF 21-15
757 General Familiarization (7-2005)
ATA 21 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Temperature Control Valve
The temperature control valve is used to control the temperature of the air-cycle-machine-compressor outlet. To do
this, it controls the amount of air that flows through the air cycle machine. If more air flows through the
temperature control valve, then less air flows through the air cycle machine. The less air that flows through the air
cycle machine, the slower it operates. Thus, the temperature at the compressor outlet decreases.
The temperature control valve is a 3-inch diameter, electrically controlled and operated, butterfly-type valve. It has
a rotary electromechanical actuator, a valve flow section, and a linear variable resistor assembly.
When power is supplied to the electric motor, it turns the gears, which turns the output shaft. The output shaft
turns the butterfly shaft. The butterfly shaft turns the butterfly plate, the visual position indicator, and the wiper
arm of the linear variable resistor assembly. The 1K ohm resistor supplies a position signal to the EICAS
computers. The 10K ohm resistor supplies a position signal to the automatic pack temperature controller.
The valve can open or close manually. When the manual override knob is turned, the clutch lets the output shaft
turn. The clutch holds the butterfly plate in the set position.
IF 21-16
757 General Familiarization (7-2005)
ATA 21 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Automatic-Pack-Temperature Controller
Two interchangeable automatic-pack-temperature controllers are used to control the air cooling packs. One
controls the right cooling pack system and the other controls the left cooling pack system.
The automatic-pack-temperature controller is a box that has a microprocessor. The microprocessor controls the
air-cooling-pack- outlet temperature and monitors the system for failures. The front panel of the controller has
lamps to show if a failure occurs. The front panel also has switches and instructions to do the built-in test
equipment (BITE).
The controllers are in the E6 rack of the aft equipment center, or they can be found on the E3 rack of the main
equipment center.
Each automatic-pack-temperature controller controls these components for each air cooling pack:
1.
2.
3.
4.
The temperature control valve
The ram-air-inlet-door and the ram-air-exhaust-door actuators of the ram-air-cooling system
The torque motor of the flow-control-and-shutoff valve
The low-limit-control valve, when the system is not in the STBY-N mode
Standby-Pack-Temperature Controller
The standby-pack-temperature-controller is a box which has two microprocessors. Each microprocessor controls
the left or the right air cooling pack. They only operate when the applicable cooling pack system is in the STBY-N
mode. The microprocessors also monitor parts of the cooling pack system for failures. The front panel of the
standby pack temperature controller has lights that come on when a failure occurs. It also has switches and
instructions to do BITE.
In the STBY-N mode, the standby-pack-temperature controller controls the low-limit-control valve. The position of
the low-limit-control valve controls how fast the air cycle machine turns. The speed of the air cycle machine
controls the compressor outlet temperature and the air-cooling-pack-outlet temperature.
757 General Familiarization (7-2005)
ATA 21 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-17
757 General Familiarization (7-2005)
ATA 21 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pack Temperature Sensor
For each air cooling pack, two pack temperature sensors supply a resistive signal of the temperature at their
positions. One pack temperature sensor sends a signal to the standby-pack-temperature controller. The other pack
temperature sensor sends a signal to the applicable (left or right) automatic-pack-temperature controller. The pack
temperature sensors are installed on the outlet of the water extractor.
Each pack temperature sensor has two series-connected glass-sealed thermisters in a stainless steel probe.
The automatic-pack-temperature controller compares the pack demand signal (from the zone temperature
controller) with the pack temperature sensor signal. The difference between the signals is the pack error. Then it
uses the pack error to control the positions of the ram air door actuators and the temperature control valve. The
standby-pack-temperature controller compares the pack temperature sensor signal to a value of 40°F (4°C) to get
the pack error. Then it uses the pack error to control the position of the low-limit-control valve.
Altitude Switches
Each altitude switch sends a signal to the applicable (left or right) automatic pack temperature controllers. The
controllers change the minimum temperature limits of the air cooling packs to make allowance for the altitude. The
switches are installed in the forward end of the left ECS bay, to the left of the cabin air supply check valve.
The altitude switch is a pressure sensing aneroid, at an altitude of 31,000 feet, the aneroid closes the electrical
switch.
When the altitude switch closes, the automatic-pack-temperature controller changes the limit of the air cooling
pack temperature from 35°F (1.7°C) to 0°F (-18°C). At low altitude the 35°F (1.7°C) temperature limit is used to
prevent ice at the condenser and the water extractor.
Compressor Outlet Temperature Sensor
For each air cooling pack, two compressor-outlet-temperature sensors supply a resistive signal of the temperature
at their position. One sensor sends a signal to the automatic-pack-temperature controller. The other sensor sends
a signal to the standby-pack-temperature controller. The sensors are installed adjacent to each other, at the aircycle-machine-compressor outlet.
Each sensor has a platinum element which is hermetically sealed in a stainless probe.
The automatic-pack-temperature controller uses the signal to control the position of the temperature control valve.
The standby pack temperature controller uses the signal to position the low-limit-control valve. The position of
these two valves controls the compressor outlet temperature.
757 General Familiarization (7-2005)
ATA 21 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Compressor Outlet Overheat Switch
In each air cooling pack, a compressor-outlet-overheat switch is installed at the air-cycle-machine-compressor
outlet. The switch is a thermal protective device. When it closes, it sends a signal to the flow control card that an
overheat condition has occurred. The switch closes when the temperature is greater than 490°F (254°C). The
switch opens again when the temperature is less than 450°F (232°C).
Pack Outlet Overheat Switch
The pack-outlet-overheat switch is at the condenser outlet. The switch is a thermal protective device. When the
pack outlet becomes too hot, the switch sends a signal to the flow control card. The switch closes when the
temperature is greater than 190°F (88°C). The switch opens again when the temperature is less than 160°F (71°C).
The pack-outlet-overheat switch has a temperature-sensitive element with a bimetallic, snap-acting disc. At a
specified temperature, the disc moves suddenly to its opposite shape. This causes the switch to close.
Cabin Air Supply Check Valve
The cabin-air-supply-check valve prevents the flow of air from the mix manifold into the air cooling packs. The
check valve has a spring, a flapper plate, and a stop. The top of the flapper plate is attached by a horizontal hinge
to the inside of the check valve. The spring is installed on the hinge to push against the flapper plate. A stop is
installed on the top of the check valve to limit how far the flapper can move.
When no air is in the system, the spring keeps the flapper plate closed. When the air flows from the air cooling
pack to the mix manifold, it pushes the flapper plate open. The air cannot flow in the opposite direction because
the flapper plate can only move in the one direction. Thus, the check valve prevents the flow of air from the mix
manifold to the air cooling pack.
757 General Familiarization (7-2005)
ATA 21 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-18
757 General Familiarization (7-2005)
ATA 21 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flow Control Card
The flow control card is in the P50 card file in the main equipment bay. The card controls the position of the flowcontrol-and-shutoff valve. It also controls the indications that show when an overheat condition occurs.
IF 21-19
757 General Familiarization (7-2005)
ATA 21 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ZONE TEMPERATURE CONTROL SYSTEM
The primary (zone) temperature control system controls the temperature of the air given to each cabin zone. The
system uses a mixture of warm pneumatic air and cool air from the air cooling packs. This will give satisfactory
air temperature for the passenger and crew.
The primary (zone) temperature control system includes the components that follow:
1.
2.
3.
4.
5.
6.
7.
8.
9.
One zone temperature controller
Three trim air modulating valves
Two check valves for the trim air supply
One pressure regulating valve for the trim air supply
Three zone temperature sensors
Three duct overheat switches
Five duct temperature sensors
Temperature selector switches for three zones
Three amber INOP warning lights.
The temperature of the conditioned air is controlled with the trim air from the downstream side of the flow control
and shutoff valve. A mixture of trim air, cooling pack air and recirculation air are used to supply the necessary air
temperature. The air temperature can be controlled for each cabin zone; the flight compartment, forward passenger
cabin, and aft passenger cabin.
The zone temperature controller uses an input temperature from each zone to operate the temperature control
system automatically. This input temperature gives the necessary demand to the controller for an increase or
decrease in temperature for each zone. The controller uses the demand signal to operate the three modulating
valves. The modulating valves control the mixture of warm air and cool air that is supplied to each zone.
The primary (zone) temperature indication system supplies a visual indication of the temperature control system.
Three modulating valves for the trim air supply are installed in the trim air ducts for each zone, downstream of the
pressure regulating valve. The modulating valve positions are electrically adjusted. These valves supply the
necessary air temperature to the flight compartment, the forward passenger zone, or the aft passenger zone.
The temperature control system contains two check valves for the trim air supply. The check valves are installed in
the trim air ducts for each zone. The check valves will not let trim air go into the pack inlet if the air cooling pack
stops.
The pressure regulating valve for the trim air system controls the pressure of the pneumatic system air that is
supplied to the modulating valves. A TRIM AIR switch-light on the pilot's overhead panel, P5, controls the valve
position.
757 General Familiarization (7-2005)
ATA 21 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Zone Temperature Control System (Continued):
A duct overheat switch is installed in the duct that goes to each cabin zone. The overheat switch operates as a
thermal protective device and closes the applicable modulating valve if the zone duct temperature increases more
than 190°F.
There are five duct air temperature sensors. The sensors get input of the temperature of the supply air that goes
into the zone ducts and the two mix manifold outlets. The sensors also supply inputs to the zone controller and
pack controllers.
There is one temperature sensor for each of the three cabin zones. Each sensor sends a temperature signal to the
zone temperature controller. The temperature sensor for the flight compartment also sends a temperature signal to
the left pack temperature controller. The temperature sensor for the forward zone also sends a temperature signal
to the right pack temperature controller.
There is one temperature selector switch for each cabin zone, installed on the pilots' overhead panel, P5. These
switches permit the selection of the AUTO control of the temperature control system.
757 General Familiarization (7-2005)
ATA 21 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-20
757 General Familiarization (7-2005)
ATA 21 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Functional Description
System Temperature Integration (Pack Demand)
The zone temperature controller calculates a temperature demand signal for the pack. The demand signal is
calculated when the controller compares the inputs of the temperature selector switch for the zone and the zone
temperature sensor. The zone temperature controller finds the lowest of the three temperature inputs and sends
that signal to the pack temperature controllers as pack demand. The pack demand is the signal from the controller
that will change the output of the pack.
The pack temperature controllers also receive the temperature inputs from the air temperature sensors in the mix
manifold duct. These temperature inputs are then sent to the zone temperature controller. The zone controller
finds the lowest of the two mix manifold temperature signals. The lowest signal is used as a reset temperature for
the pack demand signal. This lowest temperature corrects the pack demand signal to include the effect of recirculated air mixed with conditioned air from the air conditioning packs.
The pack demand is sent to a new valve until the mix manifold temperature is the same as the zone with the
highest temperature. The zone temperature controller then closes the modulating valve for the coolest zone and
adjusts the modulating valves for the other two zones. This will supply the necessary zone inlet temperatures.
Alternate Temperature Control
The pack demand signal is gone when there is a zone controller failure or all temperature selector switches are out
of the AUTO position. When the pack demand signal is gone, the pack controllers are set to a 75°F (24°C)
alternative mode.
In the alternative mode the system is set to a two zone control operation. This causes the input from the forward
zone to control the right pack and the input from the flight compartment to control the left pack. An alternative
signal in the zone temperature sensor supplies applicable pack controllers with the flight compartment or forward
zone temperature. Each pack controller uses its alternative sensor. An alternative pack command is supplied to
compare the sensor value with the set value of 75°F (24°C).
APU Output Control
The zone controller also supplies a command to the Auxiliary Power Unit (APU) to control the quantity of bleed air
to the packs during ground operation. The APU controller uses this command for the supply temperature to give
the APU output. The input signal will stay at a minimum until the pack and trim controls give the maximum
available heat or cooling. The APU command signal increases with larger demand to give the necessary output.
757 General Familiarization (7-2005)
ATA 21 - Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Zone Temperature Control (Auto Mode)
The zone temperature controller operates with 115-volts AC power. Automatic temperature control is supplied
when these conditions occur:
1. The temperature selector switches for all three zones are in the AUTO position
2. The selector switch for the pack mode is in the AUTO position
3. The TRIM AIR switch-light is ON
The AUTO input controls the temperature between 65°F (18°C) and 85°F (29°C).
During correct automatic mode operation, the zone temperature controller adjusts the modulating valve for the
applicable zone. The controller adjusts the modulating valve to supply the conditioned air from a correct mixture of
the mix manifold air and the trim air supply. This mixture of air will supply the necessary air temperature in each
cabin zone. The temperature selector switch for each zone has many different positions to get automatic
temperature control. These positions can be the AUTO (12 o'clock position), AUTO W, AUTO C or in between the
AUTO W and AUTO C positions.
The AUTO position sends a temperature signal equivalent to 75°F (24°C) to the zone temperature controller to
supply a pack or an APU demand signal. The AUTO W position sends a temperature signal to the controller that is
equivalent to 85°F (29°C). The AUTO C position sends a temperature signal to the controller that is equivalent to
65°F (18°C).
In a condition that is too hot (190°F, 88°C), the applicable flight compartment, forward, or aft zone duct overheat
relay energizes. The energized overheat relay does not permit the zone temperature controller to adjust the
modulating valves. The relay closes the valves with 28-volts DC.
The overheat relay causes the INOP light on the pilot's overhead panel, P5, to come on for the applicable zone. The
EICAS will also show the FLT DECK TEMP, FWD CABIN TEMP or AFT CABIN TEMP level C display messages.
These EICAS messages are auto events for the environmental control systems (ECS) (Ref 31-41-00). An overheat
relay does not permit the zone temperature controller to monitor the applicable zone's temperature control
components.
An INOP condition can also occur. An INOP condition shows a failure is in a control circuit for the cabin zone. The
zone temperature controller will find the failure.
An INOP condition also removes the temperature input signal, for the applicable zone, from the pack and APU
circuits. The zone temperature controller supplies a ground to permit the INOP light to come on. This will also
cause the EICAS messages, FLT DECK TEMP, FWD CABIN TEMP, or AFT CABIN TEMP, to be shown. The duct
overheat relay for the forward or aft zone also energizes which will close the applicable modulating valve with 28volts DC. The zone temperature controller will monitor the LRUs for the applicable zone.
757 General Familiarization (7-2005)
ATA 21 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Off Mode
The off mode does not permit the trim air into the applicable zone. Through the flight compartment, forward or aft
zone overheat relay the modulating valves are closed with 28-volts DC. The zone temperature selector, in the OFF
position, will cause the INOP light for that zone to come on. The EICAS messages, FLT DECK TEMP, FWD CABIN
TEMP, or AFT CABIN TEMP, will also come on for that zone. These EICAS messages are ECS auto events. In the
off mode, the zone temperature controller will not monitor the components for that zone.
Trim Air Pressure Regulation
Trim air goes through the flow control and shutoff valves. The trim air will then push open the check valves for the
trim air supply. The air then flows to the pressure regulating valve. The pressure regulating valve controls the
pressure of the trim air to 4 PSI above the cabin pressure. This keeps the trim air pressure to a limit to decrease
the noise of the air that goes through the modulating valves and their duct.
The TRIM AIR switch-light on the P5 panel, controls the pressure regulating valve. The regulating valve operates
with 28-volts DC power. Push the switch-light to the on position to open the valve. Push the switch-light to the off
position to keep 28-volts DC power away from the valve. With the switch-light in the off position the valve closes
and the yellow OFF part of the switch-light comes on. The level C message, TRIM AIR, is also shown on the EICAS
display.
757 General Familiarization (7-2005)
ATA 21 - Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-21
757 General Familiarization (7-2005)
ATA 21 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ZONE TEMPERATURE COMPONENTS
Zone Temperature Sensors
The zone temperature sensors are dual element sensors. Each element has two glass sealed thermistors connected
in series and are fully sealed in a stainless steel probe-type housing.
The temperature sensor measures the air temperature for each cabin zone. The thermistor operates at low current
levels to keep to a minimum the temperature changes because of the effect of internal sensor heat. The resistance
in each thermistor increases when the input temperature decreases and the resistance decreases when the input
temperature increases. The input temperature is sent electrically to the zone temperature controller and to the
pack temperature controllers.
Duct Temperature Sensors
The duct temperature sensors are one element sensors that have two thermistors that are glass sealed and
connected in series. The thermistors are sealed in probe-type housing with an electrical connector that is part of
the thermistor and housing. The sensor is installed on a three-hole flange with a gasket. The sensor has a
temperature control range of 35°F (1.7°C) to 180°F (82°C) and an operating temperature range of -65°F (-54°C) to
350°F (177°C).
Duct Temperature Sensors for the Zone Inlet
A duct temperature sensor is installed in the inlet duct for each cabin zone. Each sensor supplies a temperature
signal to the zone temperature controller.
Duct Temperature Sensors for the Mix Manifold
Two duct temperature sensors are installed on the mix manifold. Each sensor supplies a temperature signal to a
pack temperature controller.
Duct Overheat Switches
The zone duct overheat switch includes a liquid-filled probe, a housing, and an electrical connector. The housing
includes a diaphragm welded to the housing, a spring and retainer, and a switch. The spring holds the diaphragm
and the retainer prevents too much extension of the spring.
As the duct temperature increases, the pressure in the probe will increase in proportion to the increased
temperature because of the expanded liquid. The diaphragm feels the pressure increase and at 190°F (88°C) the
pressure is larger than the spring force and the switch will operate. When the temperature decreases to 160°F
(71°C), the spring goes back suddenly. The switch will operate when the duct temperature increases. The switch
will operate the circuitry to close the applicable modulating valve. This will also cause a yellow INOP light on the
pilot's overhead panel, P5, to come on.
757 General Familiarization (7-2005)
ATA 21 - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-22
757 General Familiarization (7-2005)
ATA 21 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Check Valves for the Trim Air Supply
The check valve for the trim air supply includes two semicircular flappers connected to a pin that goes through the
center of the valve housing. The flappers are vertical to the valve housing axis. The flappers, pin, and flapper stop
pin are installed in the one-piece valve housing. A flow arrow, on the side of the valve housing, shows the direction
of the correct airflow.
The flow of air in the direction of the flow arrow will open the flappers until they touch the stop pin. The opposite
flow of air will close the flappers. When the valve is closed, it prevents the flow of air in the opposite direction.
Zone Temperature Selectors
Each zone temperature selector has a selector shaft at one end of the housing. A knob and a twelve-pin connector
are installed on opposite ends of the selector shaft. A name plate and a wire diagram label are attached to the
housing. The selector includes a potentiometer, switches, and a cam assembly.
Turn the temperature selector switch to get the AUTO or OFF temperature control mode. In the AUTO mode, turn
the knob between C and W to select the temperature between 65°F (18°C) and 85°F (29°C).
757 General Familiarization (7-2005)
ATA 21 - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-23
757 General Familiarization (7-2005)
ATA 21 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pressure Regulating Valve for the Trim Air System
The pressure regulating valve for the trim air system is a three-inch diameter butterfly valve. The valve is
electrically controlled and pneumatically operated.
The pressure regulating valve controls the pressure of the trim air to 4 PSI above the cabin pressure. The TRIM
AIR switch-light on the P5 panel, controls the pressure regulating valve. The regulating valve operates with 28volts DC power. Push the switch-light to the on position to open the valve. Push the switch-light to the off position
to keep 28-volts DC power away from the valve. With the switch-light in the off position the valve closes and the
yellow OFF part of the switch-light comes on. The level C message, TRIM AIR, is also shown on the EICAS display.
The valve can be manually opened if the cam is turned counterclockwise (CCW). The cam can be locked with a
lock wire, to keep the valve in this position. The valve can be manually closed if the cam is turned clockwise (CW).
The cam must be locked with a lock wire to keep the valve in this position.
IF 21-24
757 General Familiarization (7-2005)
ATA 21 - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
The Modulating Valve for the Trim Air
The modulating valve is a 1.75 inch diameter, electrically operated valve.
Three modulating valves for the trim air system are installed in the mix bay at the aft end of the forward cargo
compartment. The valves control the flow of hot air to their related zones: flight deck, forward cabin, and aft cabin.
The removal and installation for each valve is the same.
The valve may be manually opened or closed, the manual knob is turned CW to open and CCW to close.
IF 21-25
757 General Familiarization (7-2005)
ATA 21 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Zone Temperature Controller
The zone temperature controller is a microcomputer that contains all of the built-in-test-equipment (BITE)
switches, fault indication lights, and BITE instructions. Two handles at the front of the controller help in the
removal/installation. Three posts in the 125-pin connector, at the rear of the controller, are in a different position.
This will prevent the installation of the controller at the incorrect location. Nine circuit cards on a primary circuit
board contain the circuits for the controller.
The zone temperature controller controls the temperature with the AUTO adjustment of the temperature selector
switches. The controller also automatically and continuously monitors faults and supplies manual operation of the
BITE. These controls supply the indication of internal controller failures and can isolate and monitor the failure of
line replaceable units (LRUs). Red lights show defective LRUs during a BITE check. An internal memory supplies
on-line BITE data. The BITE will monitor and keep LRU faults for 10 flights.
IF 21-26
757 General Familiarization (7-2005)
ATA 21 - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
CONDITIONED AIR DISTRIBUTION SYSTEM
The main distribution manifold consists of a mix manifold, ducting, and a ground-air-service connector. The mix
manifold is supplied with cold air from the air conditioning packs and warm air from the recirculation system.
Then it mixes the air and distributes it to the passenger compartment, lavatories, and the galleys. The mix
manifold also provides humidity control by extracting water from the combined airflow.
Air from the left-air-conditioning pack can bypass the mix manifold. This conditioned air flows into the flight-deckdistribution duct and is distributed directly to the flight compartment.
Whenever the airplane is parked, the air conditioning packs do not have to be used. Conditioned air can be
supplied to the mix manifold through the ground air service connector. The conditioned air can be supplied by any
standard air-conditioning ground cart.
IF 21-27
757 General Familiarization (7-2005)
ATA 21 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Mix Manifold
The mix manifold sits immediately aft of the forward cargo compartment aft bulkhead. Cylindrical in shape with a
domed top, the manifold stands about 44 inches high and 24 inches in diameter. The manifold is constructed of
thermosetting resin-impregnated Aramid fiber fabric.
Three inlet ducts enter the mix manifold's lower section. The inlet ducts allow only tangential airflow into the
manifold. Four 8.5 inch diameter outlet ducts exit the manifold from the upper section and channel the airflow to
the conditioned air distribution ducts. A spacer between the bottom of the mix manifold and the airplane skin has
a hole to drain condensed water from the manifold.
Ground Conditioned Air Connector
The ground-conditioned-air connector allows conditioned air to be supplied directly into the mix manifold from a
ground air cart. The connector consists of an eight-inch-diameter receptacle, a cover, and a swing check valve. The
connector is on the underside of the fuselage, just forward of the left-air-conditioning pack.
Conditioned air from the ground cart forces the swing check valve to open. The check valve automatically closes,
but does not seat, whenever pressurized ground air is not available. The check valve is spring loaded so it only
seats when the air conditioning packs operate (i.e. a pressurized system).
757 General Familiarization (7-2005)
ATA 21 - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-28
757 General Familiarization (7-2005)
ATA 21 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flight Deck Conditioned Air Distribution
The flight-deck-conditioned-air-distribution system supplies conditioned air to the flight deck to provide a comfortcontrolled environment for the flight crew. The system consists of a flight-deck-conditioned- air-distribution duct,
floor registers, individual crew gaspers, windshield diffusers, and an overhead register.
The conditioned air is supplied to the flight deck area by the flight-deck-distribution duct. Then it is distributed
throughout the flight deck by the registers, gaspers, and the windshield diffuser.
Conditioned air from the left-pack-outlet duct flows into the flight-deck-distribution duct before it enters the mix
manifold. In the flight-deck-distribution duct, the conditioned air mixes with trim air and flows to the flight deck.
When the left-air-conditioning pack is not operating, the conditioned air can still enter the flight-deck-distribution
duct.
If only the right-air-conditioning pack is used, conditioned air flows into the mix manifold, and then exits the
manifold through the left-pack-outlet duct and flows into the flight-deck-distribution duct.
If only a ground cart is used, conditioned air flows into the mix manifold, and then exits the manifold through the
left-pack- outlet duct and flows into the flight-deck-distribution duct (same as the right-air-conditioning pack).
Floor Registers
The flight deck floor perimeter contains four floor registers. All outlets attach to flexible ducting. Air flows through
the flexible ducting into the outlets. The outlets then disperse the air throughout the flight deck.
Overhead Register
The overhead register mounts in the ceiling of the flight deck, in between halves of the overhead circuit breaker
panel P11. The register assembly directs airflow in three different directions within the flight deck. Silicon sponge
seals, between the plenum and the register, ensure an air-tight assembly. Conditioned air enters the overhead
outlet from two inlet ducts on the top side of the plenum. Air then flows through the baffle and the nozzle
disperses the air to the flight deck.
Windshield Diffuser
The overhead drip shield contains the forward windshield diffusers. The diffusers incorporate an airflow
straightener, a baffle plate, and a nozzle to ensure uniform distribution of air over the windshield.
The sidewall panels contain the side window diffusers. The diffusers consist of a Nomex honeycomb air grille.
757 General Familiarization (7-2005)
ATA 21 - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Gaspers
The flight deck has four adjustable gaspers. There are 2 gaspers in the forward end of the cockpit, on the outboard
side of the captain's station and first officer's station, respectively. A third gasper is located below and aft of the
number 3 window, on the left side of the cockpit. The fourth gasper is mounted in the ceiling, to the right of the
overhead circuit breaker panel P11. All gaspers consist of a drilled ball within a socket and an adjustable valve to
control airflow volume.
Air Inlets
Four air inlets, two on each side of the flight deck (L, R) provide supplemental heated air to the flight deck. The
inlets are made of epoxy resin-impregnated aramid fabric.
757 General Familiarization (7-2005)
ATA 21 - Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-29
757 General Familiarization (7-2005)
ATA 21 - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Passenger Cabin Conditioned Air Distribution
The passenger cabin conditioned air distribution system channels air from the mix manifold throughout the
passenger cabin. The distribution system provides a comfortable environment for the passengers. The distribution
system consists of sidewall risers, overhead ducts, overhead outlets, sidewall outlets, and return air grills.
Conditioned air from the mix manifold mixes with trim air and enters four sidewall risers (two on each side). The
air then flows through each sidewall riser into the overhead duct. Air is dispersed into the passenger
compartments from the overhead duct through overhead outlets located along the overhead duct, and sidewall
outlets located under the stowage bins. Gasper outlets, connected to the overhead ducts, disperse conditioned air
into each lavatory. Adjustable outlets disperse conditioned air into the galley areas. Air exits the passenger cabin
through the return air grills located in the sidewalls near the floor.
Sidewall Riser Ducts
The sidewall risers are rectangular, 1.5 by 17.5 inch ducts located in the wall lining of the passenger compartment,
just aft of the wing leading edge. Three-ply Kevlar and a thick honeycomb wall make up the riser ducts.
Overhead Duct
The overhead ducting extends the full length of the passenger cabin. The ducts are made of rigid plastic foam.
Adaptor assemblies attach the main overhead duct to each overhead outlet.
Sidewall Outlet Ducts
The sidewall outlets are attached to the main overhead duct by smaller, 1.5 inch diameter flex ducts. The ducts are
made of silicone rubber-impregnated fiberglass.
Return Air Grills
The return air grills include a grill assembly and a screen. The air vents in the return air grill allow air to escape
from the passenger compartment into the lower lobe of the airplane. This allows air to re-circulate and to leave the
passenger compartment if there is a rapid decompression.
Overhead Outlets
The overhead outlets consist of a flow divider and a blade inside a housing. The flow divider and blade break up
the airflow to prevent drafts in the passenger compartment.
Sidewall Outlets
The sidewall outlets consist of an air deflector inside a housing. The housing and air deflector disperse the air
flow as it enters the passenger compartment.
Galley and Lavatory Outlets
Separate adjustable outlets provide conditioned air to galley areas and non-adjustable gasper outlets provide
conditioned air to each lavatory.
757 General Familiarization (7-2005)
ATA 21 - Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-30
757 General Familiarization (7-2005)
ATA 21 - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
CABIN AIR RECIRCULATION SYSTEM
The recirculation system consists of two separate systems. The left recirculation system provides draw-thru
cooling for the forward exhaust Electrical/Electronic (E/E) equipment cooling system. The right recirculation
system provides recirculation of conditioned air for the passenger cabin.
Re-circulated air from both systems mixes with conditioned air from the cooling packs in the mix manifold. The
50/50 mixture of fresh conditioned air and re-circulated air is distributed through the passenger cabin. The use of
re-circulated air improves airflow without placing an undue load on the air supply source or the cooling packs.
Both recirculation systems contain a filter assembly, a recirculation check valve, a recirculation fan, a current
sensor, and ducting.
The right recirculation system draws air from the passenger cabin through the return air grilles, into the forward
cargo compartment sidewalls. From there it is drawn into the filters and fan, to be blown through the check valves
into the mix manifold.
The left recirculation system provides draw-thru cooling of the main panels, overhead panels, weather radar, and
the forward equipment. Air is drawn through the components and fan, to be blown through the check valve and
filter, into the mix manifold.
757 General Familiarization (7-2005)
ATA 21 - Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-31
757 General Familiarization (7-2005)
ATA 21 - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-32
757 General Familiarization (7-2005)
ATA 21 - Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Control Panel
Two alternate action switch/lights, on the pilots' overhead P5 panel, control operation of the recirculation air fans.
A white ON light will illuminate whenever the recirculation fan switch/light is selected ON, regardless of the
operation of the fan. Depressing the switch/light again extinguishes the ON light.
Failure Indication
An amber INOP light illuminates and an EICAS advisory message L or R RECIR FAN appears for a faulty system as
well as for the selection of the switch/light
ht to the OFF position.
IF 21-33
757 General Familiarization (7-2005)
ATA 21 - Page 67
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Recirculation Air Fan
The right recirculation fan is a single stage fan. It is located on the right side of the aft bulkhead in the forward
cargo compartment.
The left recirculation fan is a single stage fan, located in the left forward sidewall of the forward cargo
compartment.
Both fans contain a thermal switch in the winding of each phase of the fan motors. The thermal switches provide
overheat protection for the motors. The recirculation fans both require 115-volts AC, 3-phase power, and are
designed for continuous high speed operation.
Recirculation Air Filters
A pre-filter and a high efficiency particulate filter are both encased in an aluminum frame. The right recirculation
filters mount to the right of the mix manifold at the rear of the forward cargo compartment, upstream of the fan.
The left filter assembly mounts to the left and just upstream of the mix manifold.
Recirculation Air Check Valve
A recirculation air check valve is located just downstream of each recirculation fan.
Airflow, in the direction of the arrow on the valve housing, opens the flappers, reverse airflow closes the check
valve flappers. This action prevents air from passing back through the system, into the cargo compartment and
forward equipment center.
757 General Familiarization (7-2005)
ATA 21 - Page 68
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-34
757 General Familiarization (7-2005)
ATA 21 - Page 69
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 70
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THE LAVATORY AND GALLEY VENTILATION SYSTEM
The ventilation system draws smoke and odor-filled air from the lavatories and galleys through overhead ducting
and exhausts it overboard. The lavatory and galley ventilation system consists of two aft equip/lavatory/galley
ventilation fans, two lavatory and galley check valves, a ventilation filter for each galley, and two fan current
sensors.
The lavatory and galley ventilation system also provides for the overboard discharge of cooling air from the aft
electrical/electronic equipment (E6) rack.
Ventilation
Either of two ventilation fans draws air from the lavatories and galleys through overhead ducting. Air exits the
galleys through a filter and then into the overhead ducting. Lavatory air exits through a ceiling vent into overhead
ducting. The same overhead duct connects all lavatories and galleys for ventilation. Two fans draw the air through
check valves before exhausting the air at the cabin air outflow valve.
Cooling air exhausted from the E6 rack dumps into the ventilation system along with the air from the galleys and
lavatories. The air is drawn through the same check valves and fans, and is exhausted at the cabin air overflow
valve.
Functional Description
Normal Operation
The operating aft equip/lav/galley vent fan draws air out of the lavatories from beneath the toilet shroud, and out
of the galleys through a ventilation filter, into the overhead ducting. The vent fan cools the E/E racks by a draw
through process. The air from all three areas is drawn to the check valve, where it is exhausted in the area of the
outflow valve. The fan check valve prevents air from passing back through the non-operating fan.
The Aft Equipment/Lavatory/Galley Ventilation Fan 1 is the primary fan. The Aft Equip/Lav/Galley Vent Fan 1
Relay is energized if the system finds a ground through the fan 1 thermal switches. The energized relay provides
115-volts AC power to fan 1.
If fan 1 does not run, fan 2 will automatically provide back-up ventilation. The vent fan 2 relay is energized if the
thermal switches are closed and the fan 1 fail relay finds a ground. The fan 1 fail relay finds a ground, common to
both fans, if the fan 1 relay is relaxed. Once the vent fan 2 relay is energized, it supplies 115-volts AC power to
operate fan 2.
Failure Indication
An EICAS maintenance message AFT EXH FAN 1 will be displayed if fan 1 does not operate. An EICAS
maintenance message AFT EXH FAN 2 will be displayed when fan 2 fails to operate.
757 General Familiarization (7-2005)
ATA 21 - Page 71
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aft Equipment, Lavatory, Galley Ventilation Fan
The aft cargo compartment contains two identical, redundant fans just forward of the aft pressure bulkhead. The
axial-flow fan consists of an electric motor, three thermal switches and a single-stage vane-axial blade
configuration. The fan motor requires 115/200-volts AC, 3-phase, 400 Hz power. The thermostats act as thermal
switches, turning the motor off if a coil reaches 400°F (218°C), and automatically turning on the alternate motor.
The thermostat resets when the motor cools to 340°F (180°C).
Lavatory and Galley Ventilation Check Valve
A check valve, just downstream of each aft equip/lav/galley ventilation fan, consists of two semi-circular flapper
doors attached to a centered hinge pin. A torsion spring closes the check valve.
Airflow, in the direction of an arrow on the check valve body, overcomes the spring and drives the flappers open. A
stop-pin limits flapper-door travel. Reverse airflow will aid the spring in closing the valve flappers.
Galley Ventilation Filter
The galley ventilation filter, made of open-weave fiberglass, press-fits into each galley ventilation inlet. The filter
prevents entrance of foreign matter into the air ducts and into the fans.
Fan Current Sensors
Two fan current sensors, one on each of electrical equipment panels P36 and P37, monitor current flow to the aft
equipment/lavatory/galley ventilation fans. An integral transformer senses the current in the wires passing
through the current sensor.
757 General Familiarization (7-2005)
ATA 21 - Page 72
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-35
757 General Familiarization (7-2005)
ATA 21 - Page 73
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HEATING SYSTEM - FORWARD & AFT CARGO COMPARTMENTS & SUPPLEMENTAL
FORWARD CARGO COMPARTMENT HEATING SYSTEM
The forward cargo compartment heating system is a self controlled system requiring no pilot or mechanic action to
place it into operation.
The forward cargo compartment heating system provides heated air to the forward cargo compartment. The system
consists of a heat exchanger, a temperature thermal switch, a current sensor or current sense relay, a heating flow
fan and the associated ducting. The fan draws air from the forward cargo compartment and forces it through the
heat exchanger and back into the cargo compartment. The heat exchanger warms the air. The heating air is
supplied from the left recirculation system. The ducting disperses the warmed air throughout the cargo
compartment.
NOTE:
On some aircraft there is no heat exchanger.
Functional Description
The forward cargo compartment heating system draws air from the cargo compartment warms the air with a
heating flow fan and, then returns the warm air into the cargo compartment.
If the cargo compartment temperature falls below 50°F (10°C), the forward-cargo-compartment-temperaturethermal switch closes. This energizes the forward cargo fan relay providing power to the forward cargo heating
flow fan. The fan draws cool air from the cargo compartment.
NOTE:
On some aircraft the air also passes through a forward cargo heat exchanger. The heat
exchanger transfers heat from the equipment cooling system exhaust air to the cool cargo
compartment air. This cargo compartment air then passes through the heating flow fan
which warms the air and exhausts it back into the cargo compartment.
When the cargo compartment temperature reaches 70°F (21°C), the forward cargo compartment temperature
thermal switch opens, interrupting power to the heating flow fan.
If the forward cargo compartment heating flow fan does not operate when the temperature thermal switch closes,
the forward cargo fan current sensor (or current sense relay) provides a ground through the thermal temperature
switch for EICAS. This displays the EICAS status and maintenance message FWD CARGO FAN.
If the forward cargo compartment fire switch on the aft pilot's control stand P8 panel is ARMED, all power is
interrupted to the heating flow fan.
757 General Familiarization (7-2005)
ATA 21 - Page 74
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-36
757 General Familiarization (7-2005)
ATA 21 - Page 75
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Forward Cargo Heat Exchanger (only on some aircraft)
The heat exchanger incorporates two cross-flowing air circuits in which cargo heating air extracts heat from left
recirculation system intake air to reduce electrical heater requirement. The heat exchanger is located behind the
FWD cargo compartment aft end liner in the mix manifold bay.
Forward Cargo Heating Flow Fan
The axial flow fan is a single stage vane axial fan integrating the motor into the fan to provide simplicity of design
and construction and to optimize unit life and reliability. The forward cargo heating flow fan is located behind the
FWD cargo compartment aft end liner in the mix manifold bay.
The axial flow fan operates at a high shaft speed when supplied with aircraft electric power. The fan impeller
drives air through the fan assembly. If the temperature of a winding reaches 400°F (204°C), the thermal switch on
that winding de-energizes the fan power relay, removing power from the fan. If the cargo compartment ambient
temperature is below 50°F (10°C), the EICAS status and maintenance message FWD CARGO FAN will be displayed.
When the fan temperature decreases to 340°F (171°C), the thermal switch will reset, allowing the fan to restart if
the cargo compartment ambient temperature is below 50°F (10°C).
Forward Cargo Temperature Thermal Switch
The Cargo Compartment Temperature Thermal switch provides temperature sensing and heating fan control. The
switch closes whenever cargo compartment temperature falls below 50°F (10°C) providing power to the heating
flow fan. The switch opens when cargo compartment temperature reaches 70°F (21°C). The thermal switch is
located in the mix bay compartment, mounted on the cargo compartment left side aft end liner panel.
Forward Cargo Fan Current Sensor
A fan current sensor is located on the left miscellaneous electrical panel, P34. The current sensor is part of the
forward cargo heating indication system. The current sensor detects if the fan has electrical power. If the fan has a
failure, the current sensor will provide a ground to the EICAS computers and the FWD CARGO FAN message will
show.
Forward Cargo Heating Inlet Grill
The grill in the aft end of the forward cargo compartment covers the inlet to the forward cargo heating system.
757 General Familiarization (7-2005)
ATA 21 - Page 76
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-37
757 General Familiarization (7-2005)
ATA 21 - Page 77
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AFT CARGO COMPARTMENT HEATING SYSTEM
The aft cargo compartment heating system is a self controlled system requiring no pilot or mechanic action to
place it into operation.
Aft cargo heating is provided by re-circulating the aft cargo compartment air by a fan and passing the air through
an electrical heater before returning to the aft cargo distribution system. Operation of the fan and heater is
automatic with both selected ON/OFF in response to a signal from a remotely located temperature switch. Heater
element and duct overheat protection are provided and are integral with the heater.
Functional Operation
The aft cargo heating system is a self-contained system. The system uses a thermal switch for compartment
temperature control and internal temperature sensors to prevent heater overheat.
If the cargo compartment temperature switch falls below 50°F (10°C) the temperature control switch closes,
energizing the aft cargo fan relay and the aft cargo heater relay. The aft cargo heater heats the air. The aft cargo
heat fan circulates air thru the heater and into the aft cargo compartment.
If the Cargo Compartment Switch reaches 70°F (21°C) the temperature control switch opens. The aft cargo heater
shuts off immediately. The aft cargo heating fan stays on for 20 seconds and then shuts off.
If heater surface temperature exceeds 390°F (199°C) or if heater outlet temperature exceeds 180°F (82.2°C) the
heater is de-energized. The heater is re-energized when surface temperature drops below 350°F (177°C) and
heater outlet temperature drops below 160°F (71.1°C).
If the heating flow fan motor temperature reaches 400°F (204°C), the fan relay is deactivated and the fan and the
heater shut off.
An EICAS status and maintenance message AFT CARGO FAN will appear when the temperature control switch is
closed and the heating flow fan is not operating.
757 General Familiarization (7-2005)
ATA 21 - Page 78
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-38
757 General Familiarization (7-2005)
ATA 21 - Page 79
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aft Cargo Heater
The heater is an electrical resistance type, which heats air forced past heat exchange surfaces. The heater operates
on 200-volts AC nominal supply voltage.
The heater is installed in line with a 6 inch O/D duct in a pressurized area. The unit is capable of heating aft
cargo compartment air under any ground or flight condition. The outside surface of the heater will not exceed
185°F at normal operating conditions.
An over temperature switch senses the temperature of the heat exchange surface. The switch opens at a
temperature which limits heat exchange surface temperature to 400°F maximum steady state, and 450°F transient
overshoot at shutdown.
An over temperature switch, downstream of the heater, senses duct air temperature. The switch opens when the
air temperature is equal to or in excess of 185°F.
Aft Cargo Heating Flow Fan
The axial flow fan is a single stage vane-axial fan integrating the motor into the fan to provide simplicity of design
and construction and to optimize unit life and reliability.
The axial flow fan is designed to operate at a nominal shaft speed of 11,500 revolutions per minute when supplied
with aircraft electric power through the fan-mounted electrical connector. The impeller causes air to flow through
the axial flow fan. If the temperature of a winding reaches 400 degrees Fahrenheit, an internal thermal switch
deactivates the fan relay, shutting the fan off. The thermal switch resets, restoring the fan, when the winding
temperature decreases to 340 degrees Fahrenheit.
Aft Cargo Temperature Control Switch
The temperature control consists of a temperature sensor/switch which senses the cargo compartment
temperatures. Whenever the cargo compartment is below 50°F the switch closes and a signal is sent to the fan and
heater. When the cargo compartment air temperature reaches 70°F the switch interrupts power to the heater and
fan.
Aft Cargo Fan Current Sensor
A fan current sensor is located on the right miscellaneous electrical panel, P37. The current sensor detects if the
fan has electrical power. If the fan has a failure, the current sensor will provide a ground to the EICAS computers
and the AFT CARGO FAN message will show.
Aft Cargo Heating Inlet Grill
The inlet grill in the front of the aft cargo compartment covers the aft cargo heating system.
757 General Familiarization (7-2005)
ATA 21 - Page 80
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-39
757 General Familiarization (7-2005)
ATA 21 - Page 81
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
SUPPLEMENTAL HEATING SYSTEM
The supplemental heating system consists of four foot electric surface heaters and two shoulder air supply heaters
that are located in the flight deck.
Foot Electric Surface Heaters
Electric foot warmers are installed in the airplane cockpit, for the convenience of the pilot and co-pilot.
The foot heater may be operated at 115 or 200-volts AC. The surface temperature is thermostatically controlled at
60° |10°F. Another thermostat is used as a safety factor to control the over temperature cut-off at 150°F
maximum. The foot warmer is energized by 115/200-volts AC, 400 Hz input voltage.
Shoulder Air Supply Heater
A shoulder air supply heater supplies heated air to the captain's shoulder area. An identical heater is provided for
the first officer. Each heater mounts in the sidewall conditioned air distribution duct immediately below the flight
deck floor. The shoulder air supply heater is a dual range electrical element heater. The heater operates on low
range using 115-volts AC or on high using 200-volts AC. An integral heating control thermostat monitors heater coil
temperature and opens at 150°F interrupting current flow to the coil. A backup safety thermostat opens at 250°F.
Individual heaters provide heated air to the captain's and first officer's shoulder areas and surface heat to the foot
area. Switches on the captain's fwd aux panel (P13) provide individual control for the captain's foot and shoulder
heaters. Similar switches on the first officer's fwd aux panel (P14) provide control for the first officer's heaters.
Each switch is a three position toggle switch with OFF, LOW and HI control positions. Each air supply heater
operates in a similar manner.
757 General Familiarization (7-2005)
ATA 21 - Page 82
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-40
757 General Familiarization (7-2005)
ATA 21 - Page 83
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 84
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PRESSURIZATION SYSTEM
The pressurization control system maintains a comfortablee and safe cabin pressure, by controlling the rate of
airflow from the cabin, in all flight modes. The air conditioning system provides air for pressurization.
The system uses two modes of operation, automatic and manual. Positive pressure relief valves and cargo vent
doors provide backup protection against over-pressurization. Pressure indicators and a 10,000 ft altitude switch
provide cabin pressure indication and warning.
A selector panel, two automatic pressure controllers and an outflow valve provide system control. The selector
panel supplies signals to the controller. The controller then opens or closes the outflow valve to maintain safe
cabin pressures. The pressurization control system provides three independent control paths for the outflow valve.
The controller's built in test function (BITE) provides system integrity checks.
IF 21-41
757 General Familiarization (7-2005)
ATA 21 - Page 85
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Functional Description
The pressure controller modulates the outflow valve to maintain the desired cabin pressure. Opening the valve
allows more air to exhaust from the airplane. This reduces cabin pressure. Closing the valve increases cabin
pressure. Modulating the valve between the full open and full close extremes, allows the desired cabin pressure to
be maintained.
Before takeoff the crew selects the desired auto mode, landing (field) altitude and the auto rate. Placing the mode
selector to AUTO 1 or AUTO 2 will release control of the pressurization system to the selected controller. The other
controller goes into a standby mode and monitors the selected controller.
The airplane's two air data computers (ADC) supply ambient pressure signals and barometric corrections to each
pressure controller. The controller's integral pressure sensor measures cabin pressure.
A throttle advanced signal from the throttle micro switch pack assembly greater than 10.5-degrees from idle, puts
the controller in the takeoff mode. At rotation the controller receives a signal from the air/ground relay, and enters
the flight mode. Once in the flight mode, the controller creates a cabin pressure auto schedule. The auto schedule
is the ideal cabin-to-ambient differential pressure for varying airplane altitudes. The selected landing altitude and
the maximum cabin-to-ambient pressure ratio (delta P = 8.6 PSI) limit the auto schedule.
The error between the cabin pressure and the cabin pressure command determines the cabin pressure change
command. The selected auto rate limits the rate of change. In cruise, the cabin pressure follows the cabin pressure
command.
The controller positions the outflow valve to maintain the correct cabin pressure and rate of cabin pressure
change. A tachometer on the outflow valve actuator provides a feedback signal to the controller. This improves the
speed response of the actuator.
Depending on how the aircraft is configured, upon reaching a cabin altitude of either 11,000 ft or 14,500 ft the
controller automatically closes the outflow valve to increase cabin pressure.
The landing altitude initiate switch overrides the existing cabin altitude demand and sets it to the landing field
altitude command. The cabin depressurizes at the full depressurization rate limit, until it reaches the landing
altitude. The maximum cabin-to-ambient pressure ratio (P = 8.6 PSI) limits the depressurization rate. The landing
initiate function is used for rapid descent from cruise or flight from high to low altitudes.
757 General Familiarization (7-2005)
ATA 21 - Page 86
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-42
757 General Familiarization (7-2005)
ATA 21 - Page 87
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Automatic Pressure Controller
Two identical automatic pressure controllers control airplane pressurization by modulating the outflow valve. One
controller remains in a standby mode, monitoring system operation. The standby controller takes over system
control if the selected controller fails. Both controllers receive identical signals. Each controller contains the
sensory and electronic logic, and the control circuit for one outflow valve control channel. 115-volts AC powers
each E/E rack-mounted controller.
IF 21-43
757 General Familiarization (7-2005)
ATA 21 - Page 88
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Outflow Valve
The outflow valve is a double-door thrust-recovery type valve. The doors are aluminum with a Teflon coating to
prevent binding and excessive friction. Connecting rods on each side of the valve frame link the forward and aft
doors together. The valve controls the flow of cabin air overboard by modulating the valve
lve doors. In the full open
position the valve has an effective flow area of 108 sq in. The valve mounts on the left underside of the airplane
immediately aft of the bulk cargo compartment.
The Outflow valve actuator mounts directly on the valve frame and drives both doors simultaneously through a
control arm and linkage. The actuator consists of two identical 115-volts AC, 400-Hz motors, a 28-volts DC motor,
and a gearbox. The gearbox includes a feedback potentiometer and limit switches.
IF 21-44
757 General Familiarization (7-2005)
ATA 21 - Page 89
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Cabin Pressure Selector Panel
The pilot’s overhead panel (P5) contains the cabin pressure selector panel. The cabin pressure selector panel
consists of a mode selector switch, auto rate input selector, landing (field) altitude input selector, valve position
indicator and manual momentary contact switch. The mode selector switch allows selection of AUTO 1, AUTO 2 or
MAN. The auto rate input selector allows the crew to select the rate of cabin pressurization. Range of selection is
50 to 2000 FPM for climb and 30 to 1200 FPM for descent. The index position indicates 500 FPM climb and 300
FPM descent. The ratio of climb to descent is always 5 to 3. The landing (field) altitude input selector allows
landing altitude selection by the crew for controller information. Landing altitude range selection is -1000 to
+14,000 ft. The valve position indicator shows outflow valve position. A momentary contact switch allows manual
control of the outflow valve.
757 General Familiarization (7-2005)
ATA 21 - Page 90
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-45
757 General Familiarization (7-2005)
ATA 21 - Page 91
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PRESSURE RELIEF SYSTEM
The pressurization relief valves provide cabin pressure relief from extreme cabin-to-ambient differential pressure,
if the pressurization system malfunctions, or in the event of a rapid emergency airplane descent.
Two positive pressure relief valves act independently of each other to prevent excessive positive cabin-to-ambient
differential pressure by opening at 8.95 PSI differential. The valves are mounted flush with the airplane skin, aft of
the left wing-to-body fairing.
Two vacuum relief valves act independently of each other to prevent negative cabin-to-ambient differential
pressure from exceeding 0.75 PSI by opening at 0.3 PSI differential. The valves are mounted flush with the
airplane skin on the right hand side, forward of the No. 1 cargo door.
Positive Pressure Relief Valve
The positive pressure relief valve is a pneumatically-operated poppet valve. The valve mounts directly on the
fuselage skin aft of the left wing-to-body fairing. The valve discharges cabin air directly overboard when high
cabin-to-ambient differential pressures exist. The relief valve limits the cabin-to-ambient differential pressure by
controlling airflow through the poppet. When the differential pressure sensed by the metering section approaches
8.95 PSI, the metering pin lifts off its seat, releasing head pressure overboard through a dump tube. This opens
the valve to reduce cabin pressure.
The valve uses two control metering sections. The primary metering section limits the differential pressure to 8.95
PSI. The secondary metering section (backup control) limits the differential pressure to 9.42 PSI.
The outside edges of the flappers are painted red and give visual indication that the valve has opened.
Pulling the flappers open from outside the airplane and tucking the flag back into the valve resets visual indication
of valve operation.
757 General Familiarization (7-2005)
ATA 21 - Page 92
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-46
757 General Familiarization (7-2005)
ATA 21 - Page 93
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Vacuum Relief Valve
Two vacuum relief valves, located in the forward fuselage skin, provide negative pressure relief. The valves are
spring-loaded closed flapper type doors. Negative pressure of -0.3 PSI on the valve opens the door inward allowing
air into the cargo compartment.
IF 21-47
757 General Familiarization (7-2005)
ATA 21 - Page 94
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PRESSURIZATION INDICATION AND WARNING SYSTEM
The Pressurization Indicating and Warning System provides visual and aural warning if the pressurization control
system fails.
High cabin altitudes endanger passengers and airplane structure. An aneroid switch senses cabin altitude. The
switch actuates an aural warning siren, level A EICAS message, CABIN ALTITUDE, and visual warnings in the
flight deck, if the cabin pressure decreases to an equivalent altitude of 10,000 feet.
Three pressure indicators in the flight deck, show cabin-to-ambient differential pressure, cabin altitude, and cabin
pressure rate of climb or descent.
Altitude Warning System
If the cabin altitude climbs to 10,000 feet, the altitude switch closes, which supplies a ground to cause the highcabin-altitude- warning signals to occur. If the cabin altitude goes below 10,000 feet, the altitude switch opens and
the high-cabin-altitude-warning signals will go off.
The high-cabin-altitude-warning signals are:
1.
2.
3.
4.
5.
The CABIN ALTITUDE light on the P5 panel comes on
The CABIN ALT light on the captains instrument P1 panel comes on
A level A EICAS message, CABIN ALTITUDE, appears
The aural warning siren sounds
NOTE:
The aural warning siren can be silenced by pressing the Master-Warning-and-Caution
switch/light.
Altitude Switch
The altitude switch is an aneroid type switch mounted on left miscellaneous electrical equipment P36 panel. As
cabin altitude increases the aneroid expands. At 10,000 feet cabin altitude the aneroid closes a warning switch
completing the warning circuit.
757 General Familiarization (7-2005)
ATA 21 - Page 95
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Differential Pressure Sensor
The differential pressure sensor is located on the left side of the forward equipment bay. A pitot static system
provides ambient pressure signals to the sensor. The cabin pressure is sensed directly by the sensor. The sensor
compares the ambient pressure to the cabin pressure, converts the pressure difference to an electrical signal and
sends it to the CABIN DIFF gage.
IF 21-48
757 General Familiarization (7-2005)
ATA 21 - Page 96
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pressurization and Indication Warning Module
The pressurization and indication warning module is located on the pilot’s overhead panel (P5). It consists of three
pressure gages which provide the cabin pressure indication. The CABIN DIFF gage receives signals from the
differential pressure sensor and displays the cabin-to-outside differential pressure. The CABIN ALT gage receives
cabin pressure signals through an open port in the indicator case and displays the pressure as the equivalent
altitude. The indicator shows positive cabin altitude from 0 to 25,000 ft. and airplane altitudes below sea level,
reaching a lower limit at the gage's OFF peg. The CABIN RATE gage receives cabin pressure signals through an
open port in the indicator case. The gage displays the rate of cabin pressure change as a function of time for
airplane altitudes of -24,000 ft. and above.
IF 21-49
757 General Familiarization (7-2005)
ATA 21 - Page 97
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 21 - Page 98
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EQUIPMENT COOLING SYSTEM
Two E/E (electrical/electronic) equipment cooling systems provide forced air cooling to the E/E equipment. The
forward system cools the forward equipment racks, main panels, overhead panels, and weather radar. The aft
system cools the aft equipment racks.
The system is designed for automatic operation requiring minimum crew attention.
The E/E cooling system utilizes part of the passenger cabin ventilation return air. This air is collected from the
cabin area and directed to the E/E compartments by the E/E supply fans. Equipment cooling is provided either by
directing supply air through the equipment ("blow thru") or by drawing E/E compartment or flight deck air through
the same equipment ("draw thru"). "Draw thru" is accomplished by using the left cabin air recirculation fan as the
forward E/E exhaust fan. The aft E/E system utilizes the aft equip/lav/galley ventilation fans as the aft E/E
exhaust fans.
Cooling air flows to the equipment rack electronics boxes through holes in metering plates located directly under
each unit. Some of the holes in the metering plates are blocked with rubber plugs depending on the cooling air
requirements of each unit. A decal is mounted on the equipment rack showing which holes are to remain open.
After cooling the equipment, the air is returned to the main E/E exhaust duct. In the forward system, the air is
ducted to the ECS mix manifold or exhausted overboard. The aft system exhausts into the aft equip/lav/galley
ventilation system.
Flow sensors are located in the supply and exhaust ducting.
Smoke sensors are located in the forward E/E cooling supply and exhaust ducting and aft E/E cooling exhaust
ducting with smoke clearance being achieved by exhausting the air overboard.
An automatic test circuit checks the proper operation of both the forward and aft equipment cooling systems. The
test is performed at the termination of every flight following engine shutdown.
Manual testing of both the forward and aft equipment cooling systems is also provided. The test is controlled by
an EQUIP COOL TEST switch located on the right sidewall panel P61.
The equipment cooling system utilizes two solid state printed circuit cards for system control. The cards are
located in the electrical equipment card file P50. The cards contain BITE which tests for operation of the fans,
current sensors, and overboard exhaust valve by illumination of 9 lights on the cards.
757 General Familiarization (7-2005)
ATA 21 - Page 99
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Functional Description – Equipment Cooling System Normal Operation
Return air from the passenger cabin vents through the return air grilles to below the passenger cabin floor. The
forward E/E equipment cooling supply fans (B10 and B11) draw in some of this air through the equipment cooling
cleaner. Forward E/E equipment cooling supply fan 2 (B11) normally operates when the airplane is airborne while
supply fan 1 (B10) provides backup in case of a failure in fan 2 (B11). Fan 1 (B10) normally operates while the
airplane is on the ground with the engines off.
Pushed by the operating forward supply fan, the cooling air passes through a check valve and is then routed
through the forward E/E equipment rack and flight deck instrument panels. The air directed to the flight deck
instruments passes through the overhead panels, main panels, center aisle stand, and then passes into the flight
deck compartment. Air is drawn from the flight deck through the main panels and joins with air drawn through the
airplane's weather radar. This air then combines with air from the flight deck and the forward E/E equipment
rack. The combined flow then passes through the left recirculation fan (B13), left recirculation filter, check valve,
and into the mix manifold.
In the aft system, the aft E/E equipment cooling supply fans (B10061 and B10062) draw in air exiting the aft
passenger cabin. Similar to the forward E/E cooling system, the air is drawn through an air cleaner by the supply
fan. Fan 2 (B10062) normally operates when the airplane is airborne, while fan 1 (B10061) provides backup. Fan 1
(B10061) normally operates on the ground with engines off.
757 General Familiarization (7-2005)
ATA 21 - Page 100
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 21-50
757 General Familiarization (7-2005)
ATA 21 - Page 101
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fan Check Valve
A check valve is installed in the cooling air supply duct downstream of each supply fan. Each valve contains two
spring-loaded closed flappers enclosed in a housing.
The valve opens in the normal flow direction at a maximum differential pressure of 0.55 inches of water. When the
differential pressure is less than 0.5 inches of water, the spring-loaded valve closes to prevent flow in the opposite
direction of normal flow.
IF 21-51
757 General Familiarization (7-2005)
ATA 21 - Page 102
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overboard Exhaust Valve
The overboard exhaust valve is located in the forward cargo compartment on the left side of the airplane about 3
feet (1 meter) aft of the forward end liner. The valve permits cooling air and smoke from the E/E cooling system
to be exhausted from the airplane. The valve consists of a spring-controlled, butterfly-type airfoil valve and a
three-position electrical actuator. Only two control positions, NORMAL and SMOKE, are used for E/E cooling
system control. The DITCH position is not used.
In operation, the valve actuator is commanded to either NORMAL or SMOKE. When the valve actuator is in
NORMAL, the valve is spring-loaded open. The flapper is aerodynamically shaped to "fly shut" when the differential
pressure between the interior and exterior surfaces of the overboard exhaust valve is greater than 2.8 PSI. When
the valve actuator is commanded to SMOKE, the valve flapper is driven partially open to exhaust equipment
cooling air overboard. Limit switches limit travel and provide position indication.
IF 21-52
757 General Familiarization (7-2005)
ATA 21 - Page 103
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Equipment Cooling Air Fan
Both forward E/E equipment cooling supply fans are located in the left forward sidewall of the forward cargo
compartment about 2 feet (0.7 meters) aft of the forward end liner. Both aft E/E equipment cooling supply fans
are located in the right sidewall aft of the aft cargo door.
The forward and aft equipment cooling fans are single stage vane-axial fans integrating the motor into the fan.
Three thermal switches are provided in the power windings of each fan to provide fan overheat protection. The
switches open the fan power relay circuit, interrupting power to the fan when temperature in a winding reaches
400°F (204°C) for the forward supply fans, and 160°F (71°C) for thee aft supply fans.
IF 21-53
757 General Familiarization (7-2005)
ATA 21 - Page 104
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Equipment Cooling Air Cleaner
The equipment cooling air cleaner is a centrifugal swirl device, using inertia to separate particulate matter from
the E/E cooling air. As the air flowing through the air cleaner is swirled, foreign material is thrown to the outside
of the flow and purged. The air cleaners are located upstream of the forward and aft equipment cooling supply
fans.
IF 21-54
757 General Familiarization (7-2005)
ATA 21 - Page 105
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fan Fail Sensor (Current Sensor)
A fan current sensor checks for current flow to each of the four supply fans. The P37 miscellaneous relay panel
houses the forward rack supply fan No. 2 current sensor and the aft supply fan No. 2 current sensor. The P36
miscellaneous relay panel houses the forward rack supply fan No. 1 current sensor and the aft supply fan No. 1
current sensor. Each current sensor consists of an induction coil and a switch. An electrical power wire to one
phase of the fan passes through the current sensor's induction coil. Current flow to the fan induces current in the
sensor's coil, actuating the current sensor's switch.
IF 21-55
757 General Familiarization (7-2005)
ATA 21 - Page 106
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Equipment Cooling Low Flow Sensor
Three low flow sensors: a forward equipment supply sensor (located in the input duct to the main panels), a
forward equipment exhaust sensor (if installed, located in exhaust duct from the main panels), and an aft
equipment cooling sensor (located in the input duct to the aft E/E equipment rack) sense low flow and provide
indication of inadequate cooling airflow. Each sensor contains a probe-like element with a self-heating thermistor.
Airflow through the ducts cools the heated probe. Low airflow in the duct cannot sufficiently cool the heated probe
and implies inadequate cooling of the E/E racks. When insufficient cooling airflow exists, the low flow sensor
closes, which provides a ground for low flow indication.
The low flow sensors contain a test circuit. When a ground signal is provided to the test terminal, a switch within
the sensor closes to provide an alternate means of activating the low flow alarm output.
IF 21-56
757 General Familiarization (7-2005)
ATA 21 - Page 107
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Equipment Cooling Smoke Sensor
There are two smoke sensors on the forward E/E cooling system supply duct (located downstream of the forward
supply fans), one sensor on the forward E/E cooling exhaust duct (located upstream of the left recirculation fan),
and one sensor on the aft E/E cooling exhaust duct (located downstream of the aft supply fans).
Each sensor consists of an inlet duct, an exhaust duct, a light source, a light trap, a photo diode, and a test LED.
The sensor's light source sends a concentrated beam of light into the light trap. The light beam passes through
airflow bled from the cooling ducts and channeled through the sensor. Smoke particles entering the sensor cause
the beam of light to scatter and hit the photo diode. The photo diode's electrical resistance decreases with light,
providing a ground for EICAS and the SMOKE light.
IF 21-57
757 General Familiarization (7-2005)
ATA 21 - Page 108
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Equipment Cooling Control Card
The equipment cooling control cards are located in the E/E equipment bay in the P50 card file. The card controls
the E/E cooling system logic for the overboard exhaust valve, smoke sensors, and fan failure.
IF 21-58
757 General Familiarization (7-2005)
ATA 21 - Page 109
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 24
ELECTRICAL POWER ............................................................................................................................. 4
AC Electrical Power ........................................................................................................................... 4
DC Electrical Power ........................................................................................................................... 4
Electrical Load Distribution................................................................................................................ 4
AC Power ............................................................................................................................................ 6
DC Power ............................................................................................................................................ 8
Primary DC System ............................................................................................................................ 8
Standby Power System ....................................................................................................................... 8
Battery/Battery Charger System ......................................................................................................... 8
ELECTRICAL POWER CONTROL ....................................................................................................... 10
Functional Description...................................................................................................................... 10
Generator Field Control Relay (GCR) .............................................................................................. 10
Generator Circuit Breaker (GCB) ..................................................................................................... 10
Bus Tie Breaker (BTB) ..................................................................................................................... 11
Auxiliary Power Breaker (APB) ....................................................................................................... 12
ELECTRICAL POWER CONTROL COMPONENTS ........................................................................... 14
Generator Control Unit (GCU) ......................................................................................................... 14
GCB/APB/BTB ................................................................................................................................ 15
Electrical Systems Control Panel & Generator Field and Hydraulic Control Panel ........................ 16
COCKPIT ELECTRICAL DISPLAY ...................................................................................................... 17
EICAS Maintenance Panel ............................................................................................................... 17
ELECTRIC/HYDRAULIC MAINTENANCE PAGE ............................................................................. 18
EICAS MESSAGE DISPLAY LOCATIONS .......................................................................................... 19
AC EXTERNAL POWER SYSTEM ....................................................................................................... 20
Functional Description...................................................................................................................... 20
EXTERNAL POWER COMPONENTS .................................................................................................. 22
Bus Power Control Unit.................................................................................................................... 22
External Power Panel ........................................................................................................................ 22
External Power Receptacle ............................................................................................................... 22
External Power Contactor ................................................................................................................. 22
Ground Power Current Transformer ................................................................................................. 22
APPLY EXTERNAL POWER TO THE AIRCRAFT ............................................................................. 24
Steps to remove External Power ....................................................................................................... 26
GENERATOR DRIVE SYSTEM ............................................................................................................ 27
Functional Description...................................................................................................................... 27
GENERATOR DRIVE COMPMONENTS ............................................................................................. 30
Integrated Drive Generator ............................................................................................................... 30
Scavenge Filter ................................................................................................................................. 31
IDG Air/Oil Heat Exchanger ............................................................................................................ 32
Quick Attach/Detach Coupling......................................................................................................... 33
IDG OIL SERVICING ............................................................................................................................. 34
757 General Familiarization (7-2005)
ATA 24 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU POWER GENERATION ................................................................................................................. 38
Functional Description...................................................................................................................... 38
APU COMPONENTS .............................................................................................................................. 39
DC POWER SYSTEM ............................................................................................................................. 40
Primary DC System .......................................................................................................................... 40
Transformer-Rectifier Units ............................................................................................................. 40
DC Tie Control Unit ......................................................................................................................... 41
DC Tie Relay .................................................................................................................................... 41
APU Start TRU Fan Control Relay .................................................................................................. 41
APU TRU Start Relay....................................................................................................................... 41
STANDBY POWER SYSTEM ................................................................................................................ 42
Functional Description...................................................................................................................... 42
STANDBY POWER SYSTEM COMPONENTS.................................................................................... 44
Static Inverter.................................................................................................................................... 44
Standby Power Control Panel ........................................................................................................... 44
AC Standby Power Relay ................................................................................................................. 45
Standby Power Relay ........................................................................................................................ 45
AC and DC Standby Bus OFF Relays .............................................................................................. 45
BATTERIES SYSTEM ............................................................................................................................ 46
Batteries ............................................................................................................................................ 46
Battery Chargers ............................................................................................................................... 48
Battery Shunts ................................................................................................................................... 50
Battery Current Monitor ................................................................................................................... 50
Main Battery Relay ........................................................................................................................... 52
Main Battery Transfer Relay ............................................................................................................ 52
DC Under Voltage Sensing Relay .................................................................................................... 52
Main Battery Charger Disable Relay ................................................................................................ 52
Main Battery Charger Detection Relay ............................................................................................ 52
APU Battery Charger Interlock Relay .............................................................................................. 52
Main/APU Battery Charger Detection Enable Relay ....................................................................... 52
757 General Familiarization (7-2005)
ATA 24- Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 24 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELECTRICAL POWER
The electrical power system consists of a 115-volts AC system, and a 28-volts DC system.
AC Electrical Power
The electrical power system generates and controls 115/200-volts, 3-phase, 400-Hz AC power. Two integrated
drive generators (IDGs), one mounted to and driven by each engine, supply main AC power. An auxiliary power unit
driven generator provides in-flight backup to the IDGs. System control is by three generator control units, and one
bus power control unit.
The system normally provides two, non-parallelable AC power channels. During a category III auto land, the static
inverter is powered from the hot battery bus to provide a third independent ac power channel.
DC Electrical Power
Two main 28-volts DC power channels are supplied by two transformer rectifier units (TRUs). The TRUs convert
main 115-volts AC power.
The main battery and its charger provide a backup source for the standby power system. During a category III auto
land, the main battery and its charger provide a third independent DC power channel.
Electrical Load Distribution
The main generators supply 115-volts, 3-phase, 400-Hz AC power to the left and right main AC buses. These buses
can also be supplied by the APU generator or external power. The main buses feed the following buses: utility
buses, galley buses, flight instrument transfer buses, AC center bus, AC standby bus, AC ground service bus, and
AC ground handling bus.
The left and right 28-volts AC buses are supplied with power by autotransformers tied to the main AC buses.
The two main 28-volts DC buses receive power from transformer-rectifier units, which convert main AC bus power.
The main DC buses supply power to the DC center bus, DC standby bus, and battery bus.
The ground service system consists of a 115-volt AC bus and a 28-volts AC bus. These buses supply power to
loads used on the ground and in-flight.
The ground handling system consists of a 115-volts AC bus and a 28-volts DC bus. These buses supply power to
ground handling equipment.
757 General Familiarization (7-2005)
ATA 24- Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-1
757 General Familiarization (7-2005)
ATA 24 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AC Power
The AC generation system consists of two integrated drive generators (IDGs), one auxiliary power unit (APU)
generator, and an external power source receptacle. All sources are rated to supply 115/200-volts, 3-phase, 400Hz, 90-kva AC power. System control and protection is provided by units in the electrical/electronics equipment
compartment.
Two IDGs, one mounted to the accessory gearbox of each engine, supply main power.
The APU generator is mounted to, and driven by, the APU. On the ground, the unit can supply ground service,
ground handling, and/or main power. In-flight, the APU generator is a substitute power source for either IDG.
External power may be connected to supply ground service, ground handling, and/or main power.
The power system includes three generator control units (GCUs). One GCU controls each IDG. The third GCU
controls the APU generator. The GCUs work with flight compartment switches to control breakers connecting power
to the main AC, and AC tie buses.
The bus power control unit (BPCU) oversees system operation. Each GCU is connected to the BPCU through a
serial data link. The BPCU sends commands and exchanges system status with the GCUs. The BPCU controls
external power, ground service, and ground handling operation.
Built-in-test equipment (BITE) in the GCUs and BPCU checks operation of the units and stores fault information.
BITE information is shown on the BPCU display.
757 General Familiarization (7-2005)
ATA 24- Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-2
757 General Familiarization (7-2005)
ATA 24 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DC Power
The DC generation system consists of three subsystems: primary dc system, standby system, and the
battery/battery charger system.
Primary DC System
The primary DC system consists of two transformer-rectifier (T-R) units. The T-R units convert 115-volts AC power
to 28-volts DC power for the DC buses.
The APU start TRU is the primary source of power for starting the APU. The TRU receives input power from the
right main AC bus, and converts it to DC power for the APU starter motor. Main system components are the APU
start TRU and the TRU APU start relay.
Standby Power System
The standby power system includes the main battery and static inverter. The standby power system provides 28volts DC and/or 115-volts, single phase, 400 Hz AC power to critical flight loads in case of a loss of main AC
and/or DC power. The system can supply these loads for 30 minutes. The standby system also serves as an
independent power source for the auto land system center channel during a category III auto land.
Battery/Battery Charger System
The main battery/charger system provides 28-volts DC power for loads on the hot battery bus, supplies power to
the APU controller during APU starting, and serves as the primary source for the standby power system.
The APU battery/charger system provides 28-volts DC power for APU starting and for the loads on the APU hot
battery bus.
757 General Familiarization (7-2005)
ATA 24- Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-3
757 General Familiarization (7-2005)
ATA 24 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELECTRICAL POWER CONTROL
Functional Description
Generator Field Control Relay (GCR)
Each generator control unit (GCU) contains a GCR. The GCR connects the GCU to the field winding of the generator.
Without the field current, no ac power is produced by the generator.
The GCUs work with the GEN CONT switches on the electrical systems control panel (P5) and the field switches on
the generator field and hydraulic control panel (P61) to control the GCRs.
The GEN CONT switches are push on/push off type, which latch in when on. When latched, an ON indicator in the
switch is visible. The field switches are momentary action type, which do not latch in. Pushing a field switch once
turns it on; pushing it again turns it off.
The right and left GEN CONT switches control their respective GCR and generator circuit breaker (GCB). When a
GEN CONT switch is off (out), its GCR and GCB are open. The GCB opens the circuit connecting generator output
power to its main ac bus. To close the GCR and GCB, the GEN CONT switch is pressed ON (in). The GCU first
energizes the close coil of the GCR. The GCR closes, allowing dc current to energize the exciter field windings of
the generator. The GCU monitors generator output power until the quality is acceptable. The GCU then sends 28volts DC to the close coil of the GCB. This allows the generator to supply ac power to its main bus. The OFF light
in the GEN CONT switch goes off when the GCB closes.
Generator Circuit Breaker (GCB)
The left and right GEN CONT switches on the electrical systems control panel (P5) control their respective GCBs.
The switches work with their generator control units (GCUs), to open and close the GCBs. In the closed position,
the GCB connects generator output to the generator main AC bus. With the GCB in the trip position, the circuit is
open. The amber OFF light in the GEN CONT switch is on when its GCB is open. Also, the EICAS display provides
the R (or L) GEN OFF level C EICAS message when the corresponding GCB is open, and the corresponding engine
is on. Each GCU contains separate GCB close and trip relays. To close a GCB, its GEN CONT switch is set to ON
(in). The generator field control relay (GCR) closes, allowing the generator to operate. The GCU checks power
quality. When the quality is acceptable, the GCU sends 28-volts DC to the GCB close coil. The GCB closes. The GEN
CONT switches are normally ON.
757 General Familiarization (7-2005)
ATA 24- Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Bus Tie Breaker (BTB)
The left and right BUS TIE switches on the electrical systems control panel (P5) control their respective BTBs.
Each switch is a push on/push off type, which latches in when on. When latched, an AUTO indicator in the switch
is visible. An amber ISLN light is also in the switch.
The object of the BTBs is to avoid a power loss on either main AC bus. If a main bus loses power, its BTB closes.
The main AC bus is then connected to the AC tie bus. This allows the main bus to receive alternate power from:
1. An external power source
2. The APU generator
3. The other main generator
The numbers correspond to alternate source priority. Each main bus generator always has highest priority.
Each generator control unit (GCU) contains separate BTB close and trip relays. The GCU in the APU position does
not control either BTB. The left and right GCU control their respective BTB. To close a BTB, the GCU close relay
sends 28-volts DC to the BTB close coil. To open a BTB, the GCU trip relay sends 28-volts DC to the BTB trip coil.
Normally, both BUS TIE switches are set to AUTO (in). The position of the BTBs is then determined by GCU logic
and wiring between the power breakers. If each main AC bus is supplied by its generator the BTBs are open. The
main AC buses are isolated from each other, and the AC tie bus. The ISLN lights in the BUS TIE switches will not
be on. An ISLN light only comes on due to a system fault, or the corresponding BUS TIE switch being in the out
position.
If the main AC buses are not energized, the battery switch must be ON for automatic BTB control. The BAT switch
is on the standby power panel. (This panel is on P5.)
Both BTBs close if one generator circuit breaker (GCB) opens, and no auxiliary or external power is available. (The
auxiliary power breaker (APB), and the external power contactor (EPC), are open.) The one operating generator
supplies both main AC buses. Both BTBs also close if both GCBs open. Any available source then supplies the
main buses.
When a BUS TIE switch is isolated (ISLN), its BTB cannot close. The amber ISLN light in the switch is on. Also,
the EICAS display provides the L (or R) BUS ISOLATED level C EICAS message when the corresponding BUS TIE
switch is set to ISLN. The main AC bus will remain isolated from the other power sources even if the bus loses
power.
757 General Familiarization (7-2005)
ATA 24 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Auxiliary Power Breaker (APB)
The APU GEN switch on electrical systems control panel (P5) controls the auxiliary power breaker (APB). The
switch is a push on/push off type, which latches in when on. When latched, an ON indicator in the switch is
visible. The switch works with the APU generator control unit (GCU) to open and close the APB. In the closed
position, the APB connects APU generator output power to the AC tie bus. With the APB in the trip position, the
circuit is open. The amber OFF light in the switch is on when the APB is open and:
1. The APU GEN switch is OFF (out position); or
2. The external power contractor (EPC) is open, and the APU is operating above 95 percent full speed
Along with the amber OFF light in the APU GEN switch, an APU GEN OFF message appears on the EICAS display.
The APU must be running, with the failure existing, for the level C EICAS message to appear.
When APU power quality is acceptable, the APU GCU allows the APB to close. The APB closes when external power
is not connected to the AC tie bus (external power contactor is open).
An open GCB results in its bus tie breaker closing. (The BUS TIE switches are assumed to be in their normal
(AUTO) position.) The APU generator supplies main AC bus power. The OFF light in the APU GEN switch should
not come on as long as the APU GEN switch is ON (latched-in position).
757 General Familiarization (7-2005)
ATA 24- Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-4
757 General Familiarization (7-2005)
ATA 24 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELECTRICAL POWER CONTROL COMPONENTS
Generator Control Unit (GCU)
The electrical power system contains three GCUs. Each GCU regulates the output voltage of its generator, and
coordinates system protection with the bus power control unit (BPCU).
The GCUs isolate faults and show failure information on the BPCU BITE display. (The GCUs are connected to the
BPCU by a serial data link.) The left and right GCUs work with flight compartment switches to control one relay
and two contactors associated with each generator: The generator field control relay, generator circuit breaker, and
bus tie breaker. The APU GCU works with flight compartment switches to control the APU generator field control
relay, and the auxiliary power breaker.
The left GCU is on shelf 1 of equipment center rack E5. The right GCU is on shelf 2 of E5. The APU GCU is on shelf
3 of E5. The front of each GCU has a nameplate, and two circuit breakers. An electrical connector is mounted on
the rear of the unit.
IF 24-5
757 General Familiarization (7-2005)
ATA 24- Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
GCB/APB/BTB
The generator circuit breakers (GCBs), auxiliary power breaker (APB), and bus tie breakers (BTBs) are identical.
Each breaker contains three main contacts along with seven normally open and sevenn normally closed auxiliary
contacts. The main contacts are kept open by spring force. Only a momentary voltage is needed to the breaker
coils to change position. No holding current is necessary.
The left BTB, and left GCB are in left generator power panel P31. The right BTB, and right GCB are in right
generator power panel P32. The APB is in APU/external power panel P34.
IF 24-6
757 General Familiarization (7-2005)
ATA 24 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Electrical Systems Control Panel & Generator Field and Hydraulic Control Panel
Electrical systems control panel M10063 is on pilots' overhead panel P5. The M10063 panel contains switches
controlling the AC power system: Two main generator control switches, one auxiliary generator control switch, two
guarded generator drive disconnect switches, two bus tie breaker switches, two utility bus switches, and one
external power switch. The panel also contains amber left and right BUS OFF lights for main AC bus status.
Generator field and hydraulic control panel M10191 is on right side panel P61. Along with hydraulic switches, the
panel contains three field switches; one for each of the left, right and APU generators. A white FIELD OFF light is
located in each switch.
IF 24-7
757 General Familiarization (7-2005)
ATA 24- Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COCKPIT ELECTRICAL DISPLAY
EICAS Maintenance Panel
The EICAS maintenance panel provides for selection of ground maintenance displays. It has momentary
pushbuttons for display selection, memory control, and self-test. The maintenance panel functions only on the
ground. The panel is located on the right side panel (P61).
IF 24-8
The DISPLAY SELECT switches control which maintenance page is to be displayed. These switches are the
ECS/MSG, ELEC/HYD, PERF/APU, CONF/MCDP, and ENG EXCD (for EXC format). When a DISPLAY SELECT
switch is pressed, its respective maintenance data format is displayed. In addition, the ECS/MSG switch is used to
cycle through additional maintenance messages, if necessary.
Five different maintenance pages can be displayed on the bottom display unit. These pages are controlled using
the EICAS maintenance panel.
The ELEC/HYD page contains both Electrical (ELEC) and Hydraulic (HYD) System parameters.
757 General Familiarization (7-2005)
ATA 24 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELECTRIC/HYDRAULIC MAINTENANCE PAGE
IF 24-9
757 General Familiarization (7-2005)
ATA 24- Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EICAS MESSAGE DISPLAY LOCATIONS
IF 24-10
757 General Familiarization (7-2005)
ATA 24 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AC EXTERNAL POWER SYSTEM
AC external power can be used on the ground to supply the main AC buses, the ground handling bus, and ground
service bus. System components include bus power control unit, external power panel, external power receptacle,
external power contactor, and ground power current transformer.
Functional Description
Connecting external power to the external power receptacle (EPR) turns on the red EXT PWR HOT BUS warning
light on P34 panel. The DC interlock circuit to the BPCU is completed, and the BPCU checks external power quality.
If external power voltage, frequency, and phase sequence are correct, the white AC PWR CONNECTED light on P30
panel will come on. The white EXT PWR AVAIL light on electrical system control panel (P5) will be on. If external
power is not selected, the clear PWR NOT IN USE light on P30 panel will be on.
The momentary action EXT PWR switch on the electrical systems control panel (P5) controls the system. Pressing
the switch once turns on the system. The BPCU causes the auxiliary power breaker and generator circuit breakers
to open, if previously closed. The BPCU energizes the EPC. The EXT PWR AVAIL and ON lights on electrical system
control panel (P5) will be on. The AC PWR CONNECTED light on P30 panel will be on. The PWR NOT IN USE light
on P30 panel will be off. Pressing the EXT PWR switch again removes external power.
If the engines are started, the EPC will open when engine speed is sufficient to produce correct quality power. The
generator circuit breakers close, providing power from the main generators. Pressing the EXT PWR switch will
switch external power to the buses.
If external power current exceeds certain limits, the BPCU will disconnect external power.
IF 24-11
757 General Familiarization (7-2005)
ATA 24- Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 24 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EXTERNAL POWER COMPONENTS
Bus Power Control Unit
The bus power control unit (BPCU) monitors the external power system. Protection circuits in the BPCU isolate
faults and disconnect poor quality power. The BPCU shares status information with the generator and auxiliary
generator control units. The BPCU controls all AC electrical buses, and controls load shedding. The BPCU is
located in forward equipment center rack E5.
External Power Panel
The external power panel P30 is located on the lower right side of the fuselage, aft of the nose gear wheel well.
The panel contains the external power receptacle, a white AC PWR CONNECTED light, and a clear PWR NOT IN
USE light.
External Power Receptacle
The external power receptacle (EPR) connects 115-volts, 3-phase, 400-Hz AC power to the airplane's electrical
system. The receptacle has six pins. Four pins are used to transfer the AC power, and two complete a DC interlock
with the BPCU. The external power receptacle is located in the P30 panel.
External Power Contactor
The external power contactor (EPC) is an electrically held, 115-volts, 3 phase, 400-Hz unit. When energized, the
EPC connects external power to the AC tie bus. The EPC is energized by the BPCU, from a flight compartment
switch command. Twenty-eight volts DC energize the contactor. The EPC is rated to operate continuously with a
current of 275 amps/phase. Protective functions in the BPCU can de-energize the EPC automatically. Control is
provided by a separate electrical connector. The EPC is in APU/external power panel P34.
Ground Power Current Transformer
The ground power current transformer (GPCT) senses current flow in each of the external power feeders. The
external power current is monitored by the BPCU. The GPCT is located in P34 panel.
757 General Familiarization (7-2005)
ATA 24- Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-12
757 General Familiarization (7-2005)
ATA 24 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APPLY EXTERNAL POWER TO THE AIRCRAFT
Turn the STBY POWER switch on the pilots' overhead panel, P5, to the AUTO position.
Push the BAT switch on the pilots' overhead panel, P5, to the ON position.
Open the access cover for the external power panel, P30.
WARNING:
757 General Familiarization (7-2005)
REMOVE ELECTRICAL POWER FROM THE POWER CABLE BEFORE YOU CONNECT THE
CABLE TO THE AIRPLANE. AN ELECTRICAL ARC CAN OCCUR WHICH CAN CAUSE INJURY
TO PERSONS AND DAMAGE TO EQUIPMENT.
ATA 24- Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Apply External Power to the Aircraft (Continued):
Connect the power cable to the external power receptacle.
Supply electrical power to the external power cable.
Make sure the CONNECTED and NOT IN USE lights on the external power panel, P30, come on.
Do these steps at the electrical system control panel on the pilots' overhead panel, P5:
1. Make sure the AVAIL light in the EXT PWR switch is on
2. Push the EXT PWR switch
3. Make sure the AVAIL and ON lights in the EXT PWR switch are on
Make sure the NOT IN USE light on the external power panel, P30, is off. Make sure the CONNECTED light on the
external power panel, P30, is on.
Push the EXT PWR switch on the P5 panel.
Make sure the white ON light in the switch comes on.
Make sure the NOT IN USE light in the P30 panel is off.
Make sure the BUS TIE switches on the P5 panel are in the AUTO position. Make sure the L BUS and the R BUS
off lights are off.
757 General Familiarization (7-2005)
ATA 24 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Apply External Power to the Aircraft (Continued):
Perform these steps to supply power to the utility/galley buses:
1. Push the applicable L or R UTILITY BUS switch on the P5 panel to the ON position
2. Make sure the yellow OFF light in the switch goes off
Steps to remove External Power
Push the EXT PWR switch on P5 panel and make sure the white ON light in the switch goes off.
Make sure the yellow (L BUS) BUS OFF light on P5 is on.
Make sure the yellow (R BUS) BUS OFF light on P5 is on.
Push the BAT switch on P5 to the off position.
Make sure the white NOT IN USE light on P30 is on.
Remove the power from the external power cable.
757 General Familiarization (7-2005)
ATA 24- Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
GENERATOR DRIVE SYSTEM
The purpose of the generator drive system is to convert mechanical power into electrical power. The major
components are two integrated drive generators (IDGs), one mounted to each engine. Each IDG has a
corresponding air/oil heat exchanger.
Functional Description
The integrated drive generator (IDG) consists of an axial gear differential/hydraulic speed control section, and a
generator section, mounted side-by-side in a single housing.
The axial gear differential/hydraulic speed control section in the IDG converts a varying engine input speed of
4,500-9,075 RPM, to a constant output speed of 12,000|150 RPM. The constant output speed allows the generator
section of the IDG to produce AC power with a frequency of 400 |5-Hz.
The power generation section in the IDG converts mechanical power (from the IDG axial gear differential/hydraulic
speed control section), into electrical power. The generator outputs 115/200-volt, 3-phase, 90-Kva, 400-Hz AC
power. This AC power is fed to the electrical system through four terminal studs on the IDG housing.
Each IDG is automatically controlled by a separate generator control unit (GCU). The GCU also works with flight
compartment switches for manual IDG control. The GCU regulates IDG output voltage, and provides IDG protection.
System status signals, shown on flight compartment indicators, are from the GCU.
To avoid damage to the IDG, the guarded GEN DRIVE DISC switch may be pressed. This activates the input
disconnect mechanism. The input disconnect mechanism disconnects the IDG input shaft from the drive section of
the IDG. This removes mechanical input power from the IDG. The mechanism is controlled by two guarded GEN
DRIVE DISC switches on the electrical control panel (P5). These switches are momentary action type, pushing a
switch once activates the disconnect.
The GEN DRIVE DISC switch should be pushed when the amber DRIVE light in the switch is on. The light is on
when charge pressure falls below 140 PSI, or IDG oil-out temperature reaches 365 | 9°F (185°|5°C). An EICAS
level C alert also appears (if the corresponding engine is running).
L (or R) GEN DRIVE. Failure to disconnect the IDG can further damage it. (The EICAS message results in an ELEC
AUTO EVENT in the EICAS system).
After the input disconnect mechanism has been activated, the IDG input shaft and the CSD section of the IDG can
only be reconnected on the ground. This is done by pulling the input drive disconnect reset ring fully out. This ring
is on the IDG housing.
A disconnected IDG that remains mounted to an engine for about 50 flight hours can receive damage to the ball
bearing assembly for the IDG input shaft.
757 General Familiarization (7-2005)
ATA 24 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-13
757 General Familiarization (7-2005)
ATA 24- Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Functional Description (Continued):
Oil exiting the IDG flows through the IDG scavenge filter. The oil then goes through the air/oil heat exchanger. The
heat exchanger provides continuous oil cooling by low pressure compressor air. A delta-pressure indicator, is used
to indicate a clogged scavenge filter. The delta-pressure indicator is located next to the filter housing. The
indicator causes a red button to pop out, if the pressure across the filter reaches 60 PSIG. This shows that the
filter is clogged. The indicator is prevented from operating when oil temperature is below 145°F (63°C).
After exiting the air/oil heat exchanger, oil returns to the IDG.
An oil level indicator is on the bottom of the IDG. The indicator consists of a display prism with two white marks
for viewing alignment. The indicator shows whether the IDG oil level is low. The indicator does not show a high
level, nor does it indicate the system oil quantity numerically. A dark prism indicates the IDG oil level is adequate.
A shiny bright prism indicates the IDG oil level is low.
IF 24-14
757 General Familiarization (7-2005)
ATA 24 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
GENERATOR DRIVE COMPMONENTS
Integrated Drive Generator
An integrated drive generator (IDG) is located on the left aft side of each engine, and is mounted on the engine
accessory gearbox. The IDG converts the varying input speed of the engine into 115/200-volts, 3-phase, 90-KVA,
400|4-Hz power. The IDG weighs 118 pounds.
unds.
IF 24-15
757 General Familiarization (7-2005)
ATA 24- Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Scavenge Filter
The scavenge filter is a replaceable component,
mponent, which filters oil flowing out of the integrated drive generator (IDG).
The oil then goes to the air/oil heat exchanger. The scavenge filter is located in the aft end of the IDG housing.
IF 24-16
757 General Familiarization (7-2005)
ATA 24 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IDG Air/Oil Heat Exchanger
The IDG air/oil heat exchanger is located on the lower left side of each
ch engine, just above each IDG. The heat
exchanger comprises an air inlet duct, cooler element, pressure relief valve, cooler outlet, and oil transfer tubes.
IF 24-17
757 General Familiarization (7-2005)
ATA 24- Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Quick Attach/Detach Coupling
The quick attach detach (QAD) kit allows rapid installation and removal of the IDG to and from the engine
accessory gearbox. The kit includes a QAD ring, and an adapter plate. The ring mates along the outside edge of the
adapter plate. The adapter plate is mounted to the accessory gearbox. An IDG mounting flange (on the IDG) mates
the IDG with the QAD. The QAD ring secures the IDG in a breech-type lockup.
IF 24-18
757 General Familiarization (7-2005)
ATA 24 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IDG OIL SERVICING
Open the IDG oil service access door on the left core cowl panel.
Access Panels:
413CL IDG Access Panel (Left Engine)
423CL IDG Access Panel (Right Engine)
Release the pressure on the IDG as shown in the graphic on the next page:
Remove the cover from the overflow drain coupling on the IDG.
Put the free end of the outlet hose into a container
WARNING:
757 General Familiarization (7-2005)
BE CAREFUL WHEN YOU CONNECT THE OUTLET ADAPTER TO THE OVERFLOW DRAIN
COUPLING. USE A RAG AROUND THE FITTING TO PREVENT A SPRAY CAUSED BY
PRESSURE IN THE IDG CASE. HOT OIL CAN CAUSE INJURIES.
ATA 24- Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IDG Oil Servicing (Continued):
Connect the adapter and the oil outlet hose to the overflow drain coupling on the IDG.
NOTE:
It is usual for some oil to drain when you connect the adapter to the coupling.
Permit the oil that drains from the hose, to flow into the container.
Keep the free end of the hose below the level of the IDG.
CAUTION:
DO NOT MIX TYPES OR BRAND NAMES OF OIL IN THE IDG. INCORRECT OILS CAN
CAUSE DAMAGE TO THE IDG.
Connect the pressure fill hose of the service cart to the pressure fill coupling on the IDG.
NOTE:
757 General Familiarization (7-2005)
Make sure the service cart pressure fill hose has only oil and does not have air mixed
with oil in the hose.
ATA 24 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Apply External Power to the Aircraft (Continued):
CAUTION:
MAKE SURE THAT THE IDG OVERFLOW DRAIN ADAPTER IS CORRECTLY INSTALLED.
MAKE SURE THAT THE FREE END OF THE OUTLET HOSE IS NOT BELOW THE SURFACE
OF THE OIL IN THE CONTAINER. MAKE SURE YOU DO NOT CAUSE KINKS IN THE
OUTLET HOSE. THE IDG CAN BE DAMAGED BY INCORRECT OIL FILL PROCEDURES.
Use the service cart to fill the IDG with oil.
CAUTION:
CONTINUE TO PUT OIL INTO THE IDG UNTIL A MINIMUM OF 1 QUART (1 LITER) OF
CLEAR OIL FLOWS INTO THE CONTAINER. THE IDG CAN BE DAMAGED BY INCORRECT
OIL FILL PROCEDURES.
The IDG is correctly filled after a minimum of 1 quart (1 liter) of oil flows into the container and no air is mixed
with the overflow oil.
NOTE:
The 1 quart (1 liter) of oil does not include the oil that drained when the hose was
connected to the overflow drain coupling or air mixed with the overflow oil.
Disconnect the service cart from the pressure fill coupling on the IDG.
Install the cover on the pressure fill coupling on the IDG.
CAUTION:
DO NOT REMOVE THE OUTLET HOSE FROM OVERFLOW DRAIN COUPLING UNTIL THE
FLOW OF OIL STOPS. TOO MUCH OIL IN THE IDG CAN CAUSE DAMAGE.
When the flow of oil stops, remove the outlet hose from the overflow drain coupling. Install the cover on the
overflow drain coupling on the IDG.
757 General Familiarization (7-2005)
ATA 24- Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 24 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU POWER GENERATION
The auxiliary power unit (APU) generator supplies electrical power on the ground when:
1. External power is not available
2. The integrated drive generators are not operating
The APU generator can supply electrical power in-flight after an integrated drive generator failure.
Functional Description
The APU generator is mounted to the APU by a keyhole flange. The APU drives the APU generator input shaft at a
constant speed of 12,000-RPM. The generator converts this mechanical input into 115/200-volts, 3-phase, 90-KVA,
400 |5-Hz AC electrical power (full output power is available at altitudes up to 40,000 ft.). The electrical output
power is available at four terminal studs on the APU generator. Leads run from the terminal studs to the auxiliary
power breaker (APB). The power is then distributed to the system.
Voltage on the generator side of the APB is regulated by the APU generator control unit (GCU). The GCU maintains
this point at 115 |1-volt AC, line-to-neutral, at 400-Hz, for loads up to 90-KVA.
The APU GCUs voltage regulator varies DC current to the generator field, to control generator output voltage. The
electrical connector, on the generator housing, allows the GCU to control generator output and provide generator
power to the APU GCU. The APU GCU is located on shelf 3 of equipment center rack E5.
The APU supplies oil to cool and lubricate the generator. Oil enters and exits through holes in the mounting flange.
Exiting oil is cooled and filtered by the APU before it is returned to the APU generator.
757 General Familiarization (7-2005)
ATA 24- Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU COMPONENTS
IF 24-19
757 General Familiarization (7-2005)
ATA 24 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DC POWER SYSTEM
Primary DC System
The primary DC system consists of two transformer-rectifier (T-R) units. The T-R units convert 115-volts AC power
to 28-volts DC power for the DC buses.
The APU start TRU is the primary source of power for starting the APU. The TRU receives input power from the
right main AC bus, and converts it to DC power for the APU starter motor. Main system components are the APU
start TRU and the TRU APU start relay.
The transformer-rectifier system supplies primary DC power and APU starting power by converting main AC power.
System components include transformer-rectifier units (TRUs), a DC tie control unit, a DC tie relay, an APU start
TRU fan control relay, and APU TRU start relay.
Transformer-Rectifier Units
Two identical TRUs convert nominal 115-volts,
volts, 3-phase AC power into unregulated 28-volts DC power. Each TRU is
capable of providing 120 amps continuous current with forced-air cooling. Internal protection circuits prevent TRU
damage due to a shorted output. An internal meter shunt allows measurement of TRU output current by the EICAS
computers. The main TRU's are located in left forward equipment center rack E1.
A third TRU is used for APU starting. This TRU contains an integral fan and overheat switches for thermal
protection. The APU start TRU is located aft of aft equipment rack E6.
IF 24-20
757 General Familiarization (7-2005)
ATA 24- Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DC Tie Control Unit
The DC tie control unit controls the operation of the DC tie relay when an auto land signal is not present. The DC
tie control unit is located in main power distribution panel P6.
DC Tie Relay
The DC tie relay is located in P6 panel. It is used to tie the main DC buses together if one loses power. The relay
receives 28-volts DC from the 28-volts DC battery bus, and is actuated by the DC tie control unit when the airplane
is not in the auto land mode.
IF 24-21
APU Start TRU Fan Control Relay
The APU start TRU fan control relay turns on the fan in the TRU when starting the APU. The fan will then keep
running until TRU temperature drops below the TRU fan control thermal switch closing temperature. The relay is in
the E6 panel.
APU TRU Start Relay
The APU TRU start relay connects 3-phase AC power from the right main AC bus to the APU start TRU. The relay
is in E6 panel.
IF 24-21
757 General Familiarization (7-2005)
ATA 24 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STANDBY POWER SYSTEM
The standby power system supplies power to critical flight loads if main power is lost.
There are two standby buses: 115-volts AC and 28-volts DC. When the left AC and left DC buses are powered the
AC standby bus receives power from the left AC bus and the DC standby bus receives power from the left DC bus.
If the left AC bus loses power, the AC standby bus receives power from the static inverter. If the left DC bus loses
power, the DC standby bus receives power from the hot battery bus.
Functional Description
Normal operation occurs when the STBY POWER switch on the standby power control panel is in AUTO position.
Standby bus power comes from the main AC and DC buses. The main battery transfer relay is energized. The AC
standby power relay is energized, sending main AC power to the standby power relay. The standby power relay is
de-energized, switching power to the standby buses. The bus off relays energize, breaking the ground connection
to the standby bus off light and EICAS computers. The static inverter receives input power, but supplies no load. If
the left main AC bus loses power or an auto land signal is received, the main battery transfer relay de-energizes.
The main battery relay energizes, connecting the hot battery bus to the standby bus, through the standby power
relay. The AC standby power relay is de-energized, connecting the static inverter to the AC standby bus.
If only the left main DC bus loses power, the main battery transfer relay de-energizes. The main battery relay
closes, supplying the DC standby bus from the hot battery bus. The AC standby bus is powered as normal.
Turning the STBY POWER switch to BAT position causes the system to operate as would occur if the left main AC
bus lost power.
Turning the STBY POWER switch to OFF position removes power from the standby bus. The amber OFF light on
the control panel comes on. The EICAS computers display a standby bus off message.
The standby buses will be off if the BAT switch on the standby power control panel is OFF and the main buses
loses power. The standby buses will also be off if the BAT switch is OFF and an auto land signal is received.
EICAS computers monitor inverter output voltage and frequency, and standby bus status.
Static inverter output voltage and frequency are displayed by the EICAS computer in the ELEC/HYD format. The
static inverter output voltage and frequency are in the STBY/BAT column in the rows labeled AC-V and FREQ.
Static inverter output voltage below 106-volts AC or above 124-volts DC will cause the EICAS status/maintenance
message STBY INVERTER to be displayed and stored in EICAS non-volatile memory.
757 General Familiarization (7-2005)
ATA 24- Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-22
757 General Familiarization (7-2005)
ATA 24 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STANDBY POWER SYSTEM COMPONENTS
Static Inverter
The static inverter converts 28-volts DC power to nominal 115-volts, single phase, 400 Hz AC power. The inverter
can supply rated power of 1 KVA continuously when forced-air cooled. The inverter is located in the forward
equipment area.
IF 24-23
Standby Power Control Panel
The standby power control panel, M10062 is located on the pilots' overhead panel P5. The M10062 panel has
switches to control the source of standby bus power. Lights on the control panel monitor power on the standby
buses and main battery charge status. On some airplanes, the STBY POWER switch must be pushed in before the
switch can be moved to OFF.
IF 24-23
757 General Familiarization (7-2005)
ATA 24- Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AC Standby Power Relay
The AC standby power relay K105 selects either the static inverter or the left main bus to power the AC standby
bus. The relay is located in the main power distribution panel, P6.
Standby Power Relay
The standby power-relay K109 controls AC and DC power to the AC and DC standby buses. The standby power
relay is in the P6 panel.
AC and DC Standby Bus OFF Relays
The AC and DC standby bus off relays K138 and K110 provide an unlatched grounded output if bus power is lost.
The relays are located in the P6 panel.
IF 24-24
757 General Familiarization (7-2005)
ATA 24 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
BATTERIES SYSTEM
The battery system is composed of the main battery system and the APU battery system. System components
include batteries, battery chargers, battery shunts, a battery current monitor, main battery charger disable relay,
APU battery charger interlock relay, main battery relay, main
in battery transfer relay, DC under voltage sensing relay,
main battery charger detection relay, and main/APU battery charger detection enable relay.
Batteries
The main and APU batteries are identical. They are nominal 24-volts DC, 20 cell, nickel-cadmium units that are
rated at 40 amp-hours. Battery weight is 96 pounds. The main battery is located in the forward equipment center.
The APU battery is located in aft equipment center rack E6.
IF 24-25
757 General Familiarization (7-2005)
ATA 24- Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 24 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Battery Chargers
The main and APU battery chargers are identical. Each charger operates from 115-volts, 3-phase, 400 Hz AC power
supplied by the AC ground service bus. Each charger can recharge a completely discharged battery in 75 minutes.
The battery charger operates in the charge mode or in the transformer-rectifier (TR) mode.
In the charge mode, the battery charger delivers a constant current of 38 amps to the battery. Charging stops when
battery voltage reaches a threshold voltage. The threshold voltage is automatically temperature adjusted by a
temperature sensor in the battery.
In the TR mode, the charger provides 28-volts DC to maintain battery charge, and meets load demands on the hot
battery bus up to 60 amps. Internal current limiting circuitry reduces the output voltage to limit current to 60
amps.
Each charger works with its battery's temperature sensor and interlock to prevent battery damage. Charging stops
if battery temperature is outside the range of 0°F - 140°F (-18°C - 60°C), or if the battery is disconnected.
The main battery charger is located in the fwd equipment center, and the APU battery charger is located in the E6
rack.
757 General Familiarization (7-2005)
ATA 24- Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-26
757 General Familiarization (7-2005)
ATA 24 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Battery Shunts
A battery shunt is installed in the ground lead of the main battery to provide for battery current measurement. This
measurement is sent to the EICAS computers.
A battery shunt is installed in the ground lead of the APU battery charger to provide APU battery charging current
measurement. This measurement is sent to the EICAS computers.
Battery Current Monitor
A battery current monitor is installed in the ground lead of the main battery. The monitor provides a signal to the
EICAS computers if the battery is discharging at greater than four to six amps.
IF 24-27
757 General Familiarization (7-2005)
ATA 24- Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 24 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Battery Relay
The main battery relay K104 connects the hot battery bus to the battery bus. Switching occurs if the left main DC
bus loses power or an auto land signal is present. Contacts on the relay provide mode control for the main battery
charger. The relay is in main power distribution panel P6.
Main Battery Transfer Relay
(1) The main battery transfer relay K106 connects the left main DC bus to the battery bus. The relay is in the P6
panel.
DC Under Voltage Sensing Relay
The DC under voltage sensing relay K113 de-energizes the main battery transfer relay if left main DC bus voltage
is low. The relay will open if voltage drops below 21.5 |.5 volts DC for 200 - 400 milliseconds, or below 15-volts
for 50 milliseconds. The relay will close when bus voltage exceeds 24.5 |.5 volts DC for 100 milliseconds. The
relay is in the P6 panel.
Main Battery Charger Disable Relay
The main battery charger disable relay (MBCDR) K115 switches AC power from the ground service bus to the
battery charger inputs. The MBCDR is located in main power distribution panel P6.
Main Battery Charger Detection Relay
The main battery charger detection relay K10425 is a time delay relay and is controlled by the main battery current
monitor and the main/APU battery charger detection enable relay K10424. When relay K10425 is energized, the
main battery charger shutdown detection circuit is connected to EICAS. Relay K10425 has a dropout time delay of
750 milliseconds and remains latched until electrical power is removed from the airplane or until the STBY POWER
switch on the standby power control panel is set to BAT. Relay K10425 is in the P6 panel.
APU Battery Charger Interlock Relay
The APU battery charger interlock relay (ABCIR) K116 connects AC power from the ground service bus to the
battery charger inputs. The ABCIR is located in E6 panel.
Main/APU Battery Charger Detection Enable Relay
The main/APU battery charger detection enable relay K10424 switches electrical power to the main battery current
monitor and the main battery charger detection relay K10425 from the battery bus to the ground handling bus
whenever the ground handling bus is energized. This ensures that the battery charging detection circuit is powered
whenever the ground service bus, which powers the battery charger, is powered. Relay K10424 is in the P6 panel.
757 General Familiarization (7-2005)
ATA 24- Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 24-28
757 General Familiarization (7-2005)
ATA 24 - Page 53
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 26
FIRE DETECTION AND PROTECTION SYSTEM ................................................................................ 4
Engine Fire Detection ......................................................................................................................... 4
Strut Overheat Detection .................................................................................................................... 4
Engine Turbine Cooling Overheat Detection ..................................................................................... 4
APU Fire Detection ............................................................................................................................ 4
Duct Leak Detection ........................................................................................................................... 4
Wheel Well Fire Detection ................................................................................................................. 4
Lower Cargo Compartment Smoke Detection.................................................................................... 4
ENGINE FIRE DETECTION SYSTEM.................................................................................................... 5
Overview............................................................................................................................................. 5
Dual Loop Faults ................................................................................................................................ 6
ENGINE FIRE INDICATION (COCKPIT)............................................................................................... 8
NACELLE OVERHEAT INDICATION (COCKPIT) ............................................................................ 10
SYSTEM COMPONENTS....................................................................................................................... 10
Engine Fire and Overheat Detectors ................................................................................................. 10
Engine Detector Locations................................................................................................................ 11
Fire and Overheat Detection Cards................................................................................................... 12
Fire/Overheat Logic/Test Cards........................................................................................................ 12
Fire/Overheat Test Panel .................................................................................................................. 13
ENGINE FIRE PROTECTION ................................................................................................................ 14
Overview........................................................................................................................................... 14
ENGINE FIRE SYSTEM COMPONENTS ............................................................................................. 16
Engine Fire Extinguisher Bottle ....................................................................................................... 16
Engine Fire Switch ........................................................................................................................... 18
Squib Test Panel ............................................................................................................................... 19
STRUT OVERHEAT DETECTION SYSTEM ....................................................................................... 20
Overview........................................................................................................................................... 20
STRUT OVERHEAT DETECTION SYSTEM COMPONENTS ........................................................... 21
Strut Overheat Detectors................................................................................................................... 21
Fire/Overheat Logic/Test Cards........................................................................................................ 22
Fire/Overheat Test Panel .................................................................................................................. 23
ENGINE TURBINE COOLING OVERHEAT DETECTION SYSTEM ............................................... 24
Overview........................................................................................................................................... 24
ENGINE TURBINE COOLING OVERHEAT DETECTION SYSTEM COMPONENTS ................... 25
Overheat Detectors ........................................................................................................................... 25
Fire/Overheat Logic/Test Cards........................................................................................................ 26
Fire/Overheat Test Panel .................................................................................................................. 27
APU FIRE DETECTION SYSTEM......................................................................................................... 28
Overview........................................................................................................................................... 28
APU FIRE DETECTION SYSTEM COMPONENTS ............................................................................ 29
APU Fire Detectors........................................................................................................................... 29
757 General Familiarization (7-2005)
ATA 26 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire and Overheat Detection Cards................................................................................................... 30
Fire/Overheat Test Panel .................................................................................................................. 31
LTS/APU/INTERPHONE Panel ...................................................................................................... 32
APU FIRE EXTINGUISHING SYSTEM ................................................................................................ 33
Overview........................................................................................................................................... 33
APU FIRE EXTINGUISHER SYSTEM COMPONENTS...................................................................... 34
APU Fire Extinguisher Bottle ........................................................................................................... 34
APU Fire Switch ............................................................................................................................... 35
APU Shutdown Remote Control Panel............................................................................................. 36
Squib Test Panel ............................................................................................................................... 37
LOWER CARGO COMPARTMENT SMOKE DETECTION SYSTEM .............................................. 39
Overview........................................................................................................................................... 39
LOWER CARGO COMPARTMENT SMOKE DETECTION SYSTEM COMPONENTS .................. 40
Cargo Smoke Detectors .................................................................................................................... 40
Smoke Detector Blowers .................................................................................................................. 40
Smoke Detector Plenum Pressure Switch......................................................................................... 40
Fire/Overheat Logic/Test Card ......................................................................................................... 42
Fire/Overheat Test Panel .................................................................................................................. 43
CARGO COMPARTMENT FIRE EXTINGUISHING SYSTEM .......................................................... 45
Overview........................................................................................................................................... 45
CARGO COMPARTMENT FIRE EXTINGUISHING SYSTEM COMPONENTS .............................. 46
Cargo Compartment Fire Extinguisher Bottle .................................................................................. 46
APU/CARGO Fire Control Panel ..................................................................................................... 48
Squib Test Panel ............................................................................................................................... 49
WHEEL WELL FIRE DETECTION SYSTEM ...................................................................................... 51
Overview........................................................................................................................................... 51
WHEEL WELL FIRE DETECTION SYSTEM COMPONENTS .......................................................... 52
Wheel Well Overheat Detectors ....................................................................................................... 52
Duct Leak and Wheel Well Fire Circuit Cards ................................................................................. 54
Fire/Overheat Test Panel .................................................................................................................. 55
Wheel Well Fire Light ...................................................................................................................... 56
WING AND BODY DUCT LEAK DETECTION SYSTEM .................................................................. 57
Overview........................................................................................................................................... 57
WING AND BODY DUCT LEAK DETECTION SYSTEM COMPONENTS ...................................... 58
Duct Leak Detectors ......................................................................................................................... 58
Duct Leak and Wheel Well Fire Circuit Cards ................................................................................. 60
Bleed Air Control Panel.................................................................................................................... 61
Miscellaneous Test Panel.................................................................................................................. 62
LAVATORY SMOKE DETECTOR SYSTEM ....................................................................................... 63
Overview........................................................................................................................................... 63
Functional Test ................................................................................................................................. 63
LAVATORY WASTE COMPARTMENT AUTOMATIC FIRE EXTINGUISHING SYSTEM .......... 64
Overview........................................................................................................................................... 64
Fire Extinguisher Bottle .................................................................................................................... 64
Temperature Indicator....................................................................................................................... 64
757 General Familiarization (7-2005)
ATA 26- Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PORTABLE FIRE EXTINGUISHING SYSTEM ................................................................................... 65
Overview........................................................................................................................................... 65
PORTABLE FIRE EXTINGUISHING SYSTEM COMPONENTS ....................................................... 66
Halon Extinguishers.......................................................................................................................... 66
Pressurized Water Extinguishers ...................................................................................................... 67
757 General Familiarization (7-2005)
ATA 26 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FIRE DETECTION AND PROTECTION SYSTEM
Engine Fire Detection
Dual-loop fire and overheat detection systems are installed in each engine area. Both loops of a system must
sense a fire/overheat condition in order for an alarm to be given. If one loop is inoperative, then the system
operates on the remaining loop.
Strut Overheat Detection
A dual-loop overheat detection system is installed on each engine strut. Both loops must sense an overheat
condition in order for an alarm to be given. If one loop is inoperative, then the system operates on the remaining
loop.
Engine Turbine Cooling Overheat Detection
A dual-loop overheat detection system is installed on each engine's turbine casing. Both loops must sense an
overheat condition in order for an alarm to be given. If one loop is inoperative, then the system operates on the
remaining loop.
APU Fire Detection
A dual-loop fire detection system is installed in the APU compartment. Both loops must sense a fire condition in
order for an alarm to be given. If one loop is inoperative, then the system operates on the remaining loop. When a
fire is detected, the APU is automatically shutdown by the detection system.
Duct Leak Detection
Two dual-loop duct leak detection systems are installed near all pneumatic ducting from the wing leading edges,
through the air conditioning bays, the wheel well, the aft cargo compartment, and the aft pressure bulkhead back
to the APU. Both loops of a system must sense a duct leak condition in order for an alarm to be given. If one loop
is inoperative, then the system operates on the remaining detector.
Wheel Well Fire Detection
A single-loop fire detection system is installed, and is continuous through both main landing gear wheel wells.
Lower Cargo Compartment Smoke Detection
Dual smoke detection systems are installed in the forward and aft cargo compartments. Both detectors in either
cargo compartment must sense smoke in order for a fire alarm to be given. If one detector is inoperative, then the
system operates on the remaining detector.
757 General Familiarization (7-2005)
ATA 26- Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FIRE DETECTION SYSTEM
Overview
Dual-loop fire and overheat detection systems are installed in each engine area. These systems monitor engine and
nacelle temperatures, and provide alerts for both fire and overheat conditions.
The four systems (left and right engine fire and nacelle overheat detection) use power from the 28-volts DC battery
bus, with alternate power from the right main DC bus. The circuit breakers for each system are on overhead circuit
breaker panel P11. The systems contain no on/off switches.
The fire detection systems are each composed of four dual-loop detectors. Thus, each loop of a fire detection
system has four detector elements, one from each detector. The nacelle overheat detection systems are each
composed of one dual-loop detector.
Each detection loop has a control card, and each engine has an automatic/fire/overheat logic/test/system
(AFOLTS) card. All four systems have one fire/overheat test panel.
The AFOLTS card configures the signals from each system in AND logic, and determines the appropriate output to
the flight crew as shown in table 1. Table 1 illustrates the normal operating mode.
757 General Familiarization (7-2005)
ATA 26 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Dual Loop Faults
Dual loop engine fire detection system faults, which render the system inoperable, will produce FAIL outputs from
AFOLTS on completion of a system test. Table 2 illustrates the test mode.
Dual loop engine fire detection faults, detected during normal operation, will show a fire condition. An engine fire
indication will be displayed on the flight deck, along with the appropriate EICAS messages.
An engine fire indication is given by the following:
1. The appropriate red (LEFT or RIGHT) fire switch handle, on the P8 control stand, lights up and unlocks
2. The appropriate red (L or R) FUEL CONTROL switch, on quadrant stand P10, lights up
3. The red FIRE light, on captain's instrument panel P1-3, comes on
4. The appropriate fire warning message - L or R ENGINE FIRE, is displayed on EICAS (Engine Indication and
Crew Alerting System - P2 panel)
5. The red master WARNING lights, on glare shield panels P7, come on
6. The fire bell sounds, on the flight deck aural warning speakers
757 General Familiarization (7-2005)
ATA 26- Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-1
757 General Familiarization (7-2005)
ATA 26 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FIRE INDICATION (COCKPIT)
A nacelle overheat indication is given by the following:
1. The appropriate amber engine overheat light - L or R ENG OVHT, on the P8 control stand, comes on
2. The appropriate overheat caution message - L or R ENG OVHT, is displayed on EICAS (P2)
3. The amber master CAUTION lights, on glare shield panels P7, come on (inhibited when both fuel control
switches are in cutoff)
4. The caution annunciation owl tone sounds, on the flight deck aural warning speakers (inhibited when both
fuel control switches are in cutoff)
757 General Familiarization (7-2005)
ATA 26- Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-2
757 General Familiarization (7-2005)
ATA 26 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NACELLE OVERHEAT INDICATION (COCKPIT)
SYSTEM COMPONENTS
Engine Fire and Overheat Detectors
A detector consists of two sensing elements which are attached to a support tube by quick-release mounting
clamps. Each sensing element contains an inert gas, a gas-emitting core material, and has a responder on one
end. The responder contains two pressure switches, and provides the electrical interface with the airplane wiring.
The inert gas in a sensing element expands as a function of average gas temperature. The gas-emitting core
material expels gas due to high localized temperatures. Both actions cause an increase in pressure in the element,
which causes an alarm pressure switch (in the responder) to activate. Both actions are completely reversible - as
the temperature decreases, the pressure decreases, and the alarm switch deactivates. When a sensing element is
damaged allowing the inert gas to leak out, an integrity pressure switch deactivates. The pressure switches provide
alarm and integrity signals to the appropriate detector control card.
The fire and overheat detectors are all identical except for the normal resistance value (at the electrical interface),
the length, and a factory-set alarm point (temperature) for the alarm switch to activate.
IF 26-3
757 General Familiarization (7-2005)
ATA 26- Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Detector Locations
The engine fire detectors are installed: on the right and left sides of the low-pressure compressor case lower
sector (zone 1); on the underside of the engine strut, above the combustion case (zone 3); and around the turbine
case lower sector (zone 3). The nacelle overheat detectors are installed behind the pre-cooler air inlets, in the
pneumatic ducting area.
IF 26-4
757 General Familiarization (7-2005)
ATA 26 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire and Overheat Detection Cards
Each detection loop has a corresponding detection control card. These cards constantly monitor and process
signals produced by each loop. The cards output alarm or fault signals to the fire/overheat logic/test card for the
appropriate system. The detection control cards are in the P54 fire detection card file. The card file is in the
electrical/electronic equipment compartment, along the right side of the nose gear wheel well.
IF 26-5
Fire/Overheat Logic/Test Cards
Two AFOLTS (Automatic Fire/Overheat Logic/Test System) cards interpret detection control card signals, provide
system functional tests, and output fire/overheat warning and fault indication signals. Each card handles the fire
and overheat detection systems for one engine. The AFOLTS cards are located in the P54 card file.
IF 26-5
757 General Familiarization (7-2005)
ATA 26- Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Test Panel
The fire/overheat test panel contains a momentary pushbutton switch - ENG/APU/CARGO. This switch initiates a
test of the fire, overheat, and smoke detection systems for the engines,
nes, APU, and cargo compartments. The test
results in cockpit fire and overheat indications.
The test panel also has an amber SYS FAIL - FAIL P-RESET switch light. This switch light indicates a fire,
overheat, or smoke detection system failure. The switch light is pressed to reset. The fire/overheat test panel is
on aft pilots' control stand P8.
IF 26-6
757 General Familiarization (7-2005)
ATA 26 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FIRE PROTECTION
Overview
The engine fire extinguishing system has controls that release one or two applications of fire extinguishing agent to
a fire in either engine compartment. The system has test capability.
The engine fire extinguishing system includes the following: Two engine fire extinguisher bottles, engine fire
control panel, and the squib test control panel.
The engine fire extinguishing system receives power from the 28-volts DC hot battery bus, through the circuit
breakers on main power distribution panel P6.
IF 26-7
757 General Familiarization (7-2005)
ATA 26- Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 26 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FIRE SYSTEM COMPONENTS
Engine Fire Extinguisher Bottle
Engine fire extinguisher bottles are located forward of the aft cargo compartment. The extinguisher bottle includes
two squib cartridges, a pressure switch, and the combined safety relief and filler port.
Two squib cartridges are on the discharge valves of each extinguisher bottle. When detonated the cartridge
ruptures a retaining disc in the valve releasing the extinguishing agent.
The pressure switch detects a decrease in bottle pressure and activates the bottle discharge lights. The pressure
switch can be manually tested.
The safety relief valve is a thermal expansion overpressure rupture disc. If bottle pressure is too high, the safety
relief ruptures, allowing the bottle to discharge. The filler port is for introducing the extinguishing agent and
pressurizing gas into the bottle.
The extinguishing agent is bromotrifluoromethane (Halon), and the pressurizing gas dry nitrogen. The agent leaves
no residue when discharged.
757 General Familiarization (7-2005)
ATA 26- Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-8
757 General Familiarization (7-2005)
ATA 26 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Fire Switch
The left and right (L/R) engine fire switch handles are located on the engine fire controll panel at pilots' aft control
stand P8. When an engine fire is detected, the engine fire switch handle red warning light comes on. A solenoid
energizes releasing a mechanical interlock on the fire switch handle shaft. When the mechanical interlock is
released, the fire switch handle can be operated by pulling the handle out and rotating it. Rotating the handle
releases the extinguishing agent. After rotation, the fire handle automatically returns to an off-center position. To
push the handle back in it must be in the center horizontal position. The fire switch handle can be manually
unlocked by pressing the button behind the handle.
IF 26-9
757 General Familiarization (7-2005)
ATA 26- Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Squib Test Panel
The squib test panel is located on the P61 right side panel. The switches on the panel are used to check
extinguisher bottle squib cartridges. When pressed, the test switch on the panel checks circuit continuity through
the squib cartridges. A green test light comes on for a successful test.
IF 26-10
757 General Familiarization (7-2005)
ATA 26 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STRUT OVERHEAT DETECTION SYSTEM
Overview
Dual-loop overheat detection systems are installed in each engine area. These systems monitor strut temperatures,
and provide alerts for overheat conditions.
Both systems (left and right strut overheat detection) use power from the 28-volts DC battery bus, with alternate
power from the right main DC bus. The circuit breakers for each system are on overhead circuit breaker panel P11.
The system contains no on/off switches.
The overheat detection systems are each composed of a dual-loop detector. Thus, each loop of an overheat
detection system has three detector switches. These switches are electrically connected in parallel, so that one
switch can trigger an alarm signal for that loop.
Each engine has a fire/overheat logic/test card, and both systems have one fire/overheat test panel.
Both systems operate in AND logic. Thus, both loops of a system must sense an overheat condition in order for an
alarm to be given.
757 General Familiarization (7-2005)
ATA 26- Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STRUT OVERHEAT DETECTION SYSTEM COMPONENTS
Strut Overheat Detectors
An overheat detector is a bi-metallic switch. When temperature exceeds the detectors trip temperature, the switch
closes a path to ground. Each detector switch is identical except for trip temperature.
Six detector switches are located in each engine strut. Four switches are mounted through the firewall, two forward
and two aft of the pre-cooler area. The remaining two switches are located in the pre-cooler area.
IF 26-11
757 General Familiarization (7-2005)
ATA 26 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Logic/Test Cards
Two AFOLTS (Automatic Fire/Overheat Logic/Test System) cards interpret detector switch signals, provide
system functional tests, and output overheat warning and fault indication signals. Each card handles the overheat
detection system for one engine. The AFOLTS cards are located in the P54 card file.
IF 26-12
757 General Familiarization (7-2005)
ATA 26- Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Test Panel
The fire/overheat test panel contains a momentary pushbuttonn switch - ENG. This switch initiates a test of the
strut overheat and in cockpit fire and overheat indications.
The test panel also has an amber SYS FAIL - FAIL P-RESET switch light. This switch light indicates a fire,
overheat, or smoke detection system failure. The switch light is pressed to reset. The fire/overheat test panel is
on aft pilots' control stand P8.
IF 26-13
757 General Familiarization (7-2005)
ATA 26 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE TURBINE COOLING OVERHEAT DETECTION SYSTEM
Overview
Dual-loop overheat detection systems are installed in each engine turbine area. These systems monitor engine
turbine cooling air temperatures, and provide alerts for overheat conditions.
The systems use power from the 28-volts DC battery bus, with alternate power from the right main DC bus. The
circuit breakers for each system are on overhead circuit breaker panel P11. The system contains no ON/OFF
switches.
The overheat detection systems are each composed of a dual-loop detector. Thus, each loop of a fire detection
system has one detector switch.
Each engine has a fire/overheat logic/test card, and both systems have one fire/overheat test panel.
Both systems operate in AND logic. Thus, both loops of a system must sense overheat condition in order for an
alarm to be given.
757 General Familiarization (7-2005)
ATA 26- Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE TURBINE COOLING OVERHEAT DETECTION SYSTEM COMPONENTS
Overheat Detectors
The overheat detectors are temperature-sensitive, normally-closed switches. The switches are of bi-metallic
construction, and function on the differential expansion properties of the two metals.
Both switches are mounted on a mixer block, located at the nine o'clock position on the turbine case. The mixer
block is fed with turbine cooling air by an air sampling tube. When turbine cooling air temperature exceeds the trip
point of the switch, the switch opens to provide an overheat signal.
IF 26-14
757 General Familiarization (7-2005)
ATA 26 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Logic/Test Cards
Two AFOLTS (Automatic Fire/Overheat Logic/Test System) cards interpret detection switch signals, provide
system functional tests, and output fire warning and fault indication signals. Each card handles the overheat
detection systems for one engine. The AFOLTS cards are located in the P54 card file.
IF 26-15
757 General Familiarization (7-2005)
ATA 26- Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Test Panel
The fire/overheat test panel contains a momentary pushbuttonn switch - ENG. This switch initiates a test of the
engine turbine cooling overheat and strut overheat detection systems. The test results in cockpit fire and overheat
indications.
The test panel also has an amber SYS FAIL - FAIL P-RESET switch light. This switch light indicates a fire,
overheat, or smoke detection system failure. The switch light is pressed to reset. The fire/overheat test panel is
on aft pilots' control stand P8.
IF 26-16
757 General Familiarization (7-2005)
ATA 26 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU FIRE DETECTION SYSTEM
Overview
A dual-loop fire detection system is installed in the APU compartment. This system monitors temperatures in the
compartment, and provides alerts for a fire condition.
The APU fire detection system uses power from the 28-volts DC battery bus, with alternate power from the right
main DC bus. The system circuit breakers are on overhead circuit breaker panel P11. The system contains no
on/off switches.
The system is composed of two dual-loop detectors. Thus, each loop of the system has two detector elements, one
from each detector.
Each detection loop has a control card. The system has one fire/overheat logic/test card, and a fire/overheat test
panel.
When the airplane is on the ground, an APU fire indication is given at the nose landing gear - on the
LTS/APU/INTERPHONE panel, as well as in the cockpit.
The detection system operates in AND logic on the ground and in the air. Thus, both loops of the system must
sense a fire or fault condition in order for an alarm to be given.
757 General Familiarization (7-2005)
ATA 26- Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU FIRE DETECTION SYSTEM COMPONENTS
APU Fire Detectors
A detector consists of two sensing elements which are attached to a support tube by quick-release mounting
clamps. Each sensing element contains an inert gas, a gas-emitting core material, and has a responder on one
end. The responder contains two pressure switches and provides the electrical interface with the airplane wiring.
The inert gas in a sensing element expands as a function of average as temperature. The gas-emitting core
material expels gas due to high localized temperatures. Both actions cause an increase in pressure in the element,
which causes an alarm pressure switch (in the responder) to activate. Both actions are completely reversible - as
the temperature decreases, the pressure decreases, and the alarm switch deactivates. When a sensing element is
damaged allowing the inert gas to leak out, an integrity pressure switch deactivates. The pressure switches provide
alarm and integrity signals to the appropriate detector control card.
The fire detectors are installed above and below the APU. One detector forms a 180 degree arc around the air inlet
above the APU and the other detector is below the APU, on an access door.
IF 26-17
757 General Familiarization (7-2005)
ATA 26 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire and Overheat Detection Cards
Each detection loop has a corresponding detection control card. These cards constantly monitor and process
signals produced by each loop. The cards output alarm or fault signals to the fire/overheat logic/test card.
The detection control cards are in the P54 fire detection card file. The card file is in the electrical/electronic
equipment compartment, along the right side of the fuselage.
The AFOLTS (Automatic Fire/Overheat Logic/Test System) card interprets detection control card signals, provides
a system functional test, and outputs fire warning and fault indication signals. The AFOLTS card is located in the
P54 card file.
IF 26-18
757 General Familiarization (7-2005)
ATA 26- Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Test Panel
The fire/overheat test panel contains a momentary pushbutton switch - ENG/APU/CARGO. This switch initiates a
test of the fire, overheat, and smoke detection systems for the engines, APU, and cargo compartments. The test
results in cockpit fire and overheat indications.
The test panel also has an amber SYS FAIL - FAIL P-RESET switch light. This switch light indicates a fire,
overheat, or smoke detection system failure. The switch light is pressed to reset. The fire/overheat test panel is
on aft pilots' control stand P8.
IF 26-19
757 General Familiarization (7-2005)
ATA 26 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LTS/APU/INTERPHONE Panel
LTS/APU/INTERPHONE panel P62 contains the APU FIRE light, and the APU fire warning horn. In addition, the
panel contains switches that allow the APU to be shutdown (when a fire indication occurs), and the APU fire
extinguisher to be discharged from the nose landing gear area. P62 is located on the right side of the nose landing
gear.
IF 26-20
757 General Familiarization (7-2005)
ATA 26- Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU FIRE EXTINGUISHING SYSTEM
Overview
The APU fire extinguishing system has controls in two locations that release one application of fire extinguishing
agent to a fire in the APU compartment. The system has test capability.
The APU fire extinguishing system includes the APU fire extinguisher bottle located forward of the APU firewall,
the APU/CARGO fire control panel on pilots' aft control stand P8, the P62 APU shutdown remote control panel on
the nose gear, and the squib test panel located on the P61 right side panel.
The APU fire extinguishing system receives power from the 28-volts DC hot battery bus, through a circuit breaker
on main power distribution panel P6.
The APU shutdown remote control panel P62 is located on the nose landing gear. P62 contains a red APU FIRE
LIGHT, APU SHT DN/C.O. SWITCH, and APU BTL DISCH SWITCH. When a fire is detected in the APU
compartment, the APU FIRE LIGHT comes on, and a warning horn sounds. Pressing APU SHT DN/CO shuts down
the APU and arms the remote APU bottle discharge. Pressing APU BTL DISCH SWITCH discharges the APU fire
bottle. If APU SHT DN/C.O. SWITCH is pressed, the APU remote fire indication power must be recycled OFF and
ON to reset the system, otherwise fire horn is disabled.
757 General Familiarization (7-2005)
ATA 26 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU FIRE EXTINGUISHER SYSTEM COMPONENTS
APU Fire Extinguisher Bottle
One APU fire extinguisher bottle is found forward of the APU fire wall on the lower right side. The extinguisher
bottle includes a squib cartridge, pressure switch, and the combined safety relief and filler port.
The squib cartridge located on the discharge valve, when detonated, ruptures a retaining disc in the valve releasing
the extinguishing agent.
The pressure switch detects a decrease in bottle pressure and activates the bottle discharge lights. The pressure
switch can be tested using the manual override test hex key or the ground test pushbutton.
The safety relief is a thermal expansion overpressure rupture disc. If bottle pressure is too high, the safety relief
ruptures, allowing the bottle to discharge. The filler port is for introducing the extinguishing agent and pressurizing
gas into the bottle.
The extinguishing agent is bromotrifluoromethane (Halon), and the pressurizing gas dry nitrogen. The agent leaves
no residue when discharged.
IF 26-21
757 General Familiarization (7-2005)
ATA 26- Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Fire Switch
The APU fire switch is located on the APU fire control panel at pilot's aft control stand P8. When a fire is detected
in the APU compartment, the APU fire switch handle red warning light comes on. A solenoid energizes releasing a
mechanical interlock on the fire switch handle shaft. When the mechanical interlock is released, the fire switch
handle can be operated by pulling the handle out and rotating
ting it. Rotating the handle releases the extinguishing
agent. After rotation, the fire handle automatically returns to an off-center position. To push the handle back in it
must be in the center horizontal position. The fire switch handle can be manually unlocked by pressing the button
behind the handle.
IF 26-22
757 General Familiarization (7-2005)
ATA 26 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Shutdown Remote Control Panel
The APU shutdown remote control panel P62 is located on the nose landing gear. P62 contains a red APU FIRE
LIGHT, APU SHT DN/C. O. SWITCH, and APU BTL DISCH SWITCH. When a fire is detected in the APU
compartment, the APU FIRE LIGHT comes on, and a warning horn sounds. Pressing APU SHT DN/CO switch shuts
down the APU and arms the remote APU bottle discharge. Pressing APU BTL DISCH SWTICH discharges the APU
fire bottle. If APU SHT DN/C. O. SWITCH is pressed, the APU remote fire indication power must be recycled OFF
and ON to reset the system, otherwise fire horn is disabled.
IF 26-23
757 General Familiarization (7-2005)
ATA 26- Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Squib Test Panel
The squib test panel is located on the P61 right side panel. The switches on the panel are used to check
extinguisher bottle squib cartridges. When pressed, the TEST switch on the panel checks continuity of the squib
cartridges. A green test light comes on for a successful test.
IF 26-24
757 General Familiarization (7-2005)
ATA 26 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 26- Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LOWER CARGO COMPARTMENT SMOKE DETECTION SYSTEM
Overview
Each of the forward and aft cargo compartments smoke detection systems are dual systems. The smoke detector
systems monitor the air in the cargo compartments and provide alerts for smoke conditions.
The dual smoke detection systems consist of a forward and aft component which is powered by the 28-volts DC
battery bus and 115-volts AC ground service bus respectively. The system circuit breakers are on the overhead
panel P11. There are no smoke detection system ON/OFF switches.
The main components of each smoke detection system are: a pair of smoke detectors, a pair of air-flow blowers, a
plenum pressure switch, a fire/overheat logic/test card, and a fire/overheat test panel.
The operation of both forward and aft systems is the same. Both systems operate in AND logic. Thus, both
detectors of a system must sense smoke in order for a fire alarm to be given.
757 General Familiarization (7-2005)
ATA 26 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LOWER CARGO COMPARTMENT SMOKE DETECTION SYSTEM COMPONENTS
The forward cargo compartment smoke detector assembly (detectors, blowers, and plenum pressure switch) is
located in the forward cargo compartment ceiling just aft of No. 1 cargo door. The aft cargo compartment smoke
detector assembly is mounted behind the aft cargo compartment sidewall just forward of No. 2 cargo door.
Cargo Smoke Detectors
The smoke detectors operate on an optical principle. The detector consists of a photocell and pilot light in a lightproof chamber plus associated amplifier, output, and test circuits on a printed circuit board. The detector has inlet
and outlet tubes to provide access to air and smoke.
On SMOKE DETECTORS WITH LAMP PLACARDS: Smoke entering the light-proof chamber reflects the light
from the pilot lamp onto a photocell. The photocell provides a signal to an amplifier where the signal is processed
and routed to a relay. The relay provides an alarm output signal to the fire/overheat logic/test card. Because the
smoke detector is very sensitive to the amount of light in the detector, the correct lamp must be used in the
detector. If the incorrect lamp is used, then the detector will not be able to sense a smoke condition.
On SMOKE DETECTORS WITHOUT LAMP PLACARDS: Smoke entering the light-proof chamber reflects the light
from the LED onto a photo diode. There are two photo diodes in the chamber. One controls the intensity of the
LED. The other photo diode measures the reflected light from the smoke particles. The reflected light increases the
voltage output across the second photo diode. When the smoke concentration gets above 10%, the voltage will
increase to above the reference voltage monitored by a comparator and relay driver. The relay provides an alarm
output signal to the fire/overheat logic/test card. These detectors use Light Emitting Diodes (LEDs). The LEDs are
not replaceable parts.
Smoke Detector Blowers
One smoke detector blower is used to pull air into sampling tubes distributed throughout the cargo compartment,
and then through the two detectors. A second blower serves as a back-up in case the operating blower fails.
Smoke Detector Plenum Pressure Switch
An air-flow pressure switch is installed in the plenum of each cargo compartment smoke detection system. The
switch detects a failed blower, and turns on the standby blower and respective FWD or AFT DET FAN EICAS
maintenance message. If the standby blower has also failed, EICAS status message CARGO DET AIR will be
displayed.
757 General Familiarization (7-2005)
ATA 26- Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-25
757 General Familiarization (7-2005)
ATA 26 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Logic/Test Card
The AFOLTS (Automatic Fire/Overheat Logic/Test System) card interprets smoke detector signals, provides
system functional tests, and outputs the fire or fail indication signals for both systems.
The AFOLTS card has a function to reconfigure the alarm logic from the normally "AND" (dual loop operation) logic
to the "OR" (single loop operation) logic when a detector fails during a power-up or manual test.
AIRPLANES WITH -136 AFOLTS CARDS (PRR 54790); The AFOLTS card has a function to reconfigure the alarm
logic when the detector fails a routine test. It initiates a test when one detector in a zone is in alarm and the other
is not for more than 5 seconds, and periodically tests the second detector once every hour during flights
dispatched with one detector inoperative.
The AFOLTS card is in the P54 fire detection card file. The card file is in the electrical/electronic equipment
compartment.
IF 26-26
757 General Familiarization (7-2005)
ATA 26- Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Test Panel
The fire/overheat test panel contains a momentary push button switch - ENG/APU/CARGO. This switch initiates a
test of the fire, overheat, and smoke detection systems for the engines, APU, and cargo compartments. The test
results in cockpit fire and overheat indications.
The test panel also has an amber SYS FAIL - FAIL P-RESET switch light. This switch light indicates a fire,
overheat, or smoke detection system failure. The switch light is pressed to reset. The fire/overheat test panel is
on aft pilots' electrical control panel P8.
IF 26-27
757 General Familiarization (7-2005)
ATA 26 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 26- Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
CARGO COMPARTMENT FIRE EXTINGUISHING SYSTEM
Overview
The cargo compartment fire extinguishing system has controls that release one or two applications of fire
extinguishing agent to a fire in either cargo compartment. The system has test capability.
The cargo fire extinguishing system includes the following: two cargo fire extinguisher bottles, APU/CARGO fire
control panel, and the squib test control panel.
The cargo fire extinguishing system receives power from the 28-volts DC hot battery bus, through a circuit breaker
on main power distribution panel P6. The FIRE EXTINGUISHING CARGO BTL 1 AND BTL 2 circuit breakers on main
power distribution panel P6 controls power to the system.
757 General Familiarization (7-2005)
ATA 26 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
CARGO COMPARTMENT FIRE EXTINGUISHING SYSTEM COMPONENTS
Cargo Compartment Fire Extinguisher Bottle
Two cargo compartment fire extinguisher bottles are located forward of the aft cargo compartment. The
extinguisher bottle includes two squib cartridges, a pressure switch, and the combined safety relief and filler port.
Two squib cartridges are on the discharge valves of each extinguisher bottle. When detonated, the cartridge
ruptures a retaining disc in the valve releasing the extinguishing agent.
The pressure switch detects a decrease in bottle pressure and activates the bottle discharge lights. The pressure
switch can be manually tested by the push of the button or the turn of the hex key located on the switch.
The safety relief valve is a thermal expansion over-pressure rupture disc. If bottle pressure is too high, the safety
relief ruptures, allowing the bottle to discharge. The filler port is for introducing the extinguishing agent to a
pressurizing gas into the bottle.
The extinguishing agent is bromotrifluoromethane (Halon), and the pressurizing gas is dry nitrogen. The agent
leaves no residue when discharged.
757 General Familiarization (7-2005)
ATA 26- Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-28
757 General Familiarization (7-2005)
ATA 26 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU/CARGO Fire Control Panel
The FWD/AFT ARMED and BTL 1/2 DISCH switches are located on the APU/CARGO fire control panel at pilots'
aft control stand P8. When a cargo compartment fire is detected, the red FWD or AFT warning light comes on. The
ARMED switch on the panel, when pressed, arms the extinguisher bottle for discharge to the compartment desired.
When the bottle is armed, a white (ARMED) light appears on the switch. To discharge the extinguishing agent,
press the BTL DISCH 1 or 2 switch. Cargo bottle 1 or 2 will be discharged depending on the switch selected. The
amber DISCH light comes on when the bottle discharges.
IF 26-29
757 General Familiarization (7-2005)
ATA 26- Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Squib Test Panel
The squib test panel is located on the P61 right side panel. The TEST switch on the panel is used to check
extinguisher bottle squib cartridges. When pressed, the TEST switch on the panel checks continuity through the
squib cartridges. A green test light for each squib comes on for a successful test.
IF 26-30
757 General Familiarization (7-2005)
ATA 26 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 26- Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WHEEL WELL FIRE DETECTION SYSTEM
Overview
A continuous single loop sensor system is installed in the left and right main wheel wells for detecting overheat
conditions in those areas. When an overheat temperature is reached, the resistance of the sensor decreases such
that a warning alert is provided to the flight crew.
An overheat condition causes the master warning lights (red) on the glare shield (P7) to come on, the fire bell to
sound, the WHEEL WELL FIRE (red) light to come on, and the EICAS indicator (P2) to display WHEEL WELL FIRE.
The fire bell and master warning lights are reset by pressing one of the lights. The WHEEL WELL FIRE light and
the EICAS display will remain on as long as the overheat condition exists.
A test switch is provided on the fire/overheat test module at the rear of the pilots' control stand (P8). When
operated, the test switch simulates a ground on the continuous loop sensor. The duct leak and wheel well overheat
control card detects the ground as being an overheat condition and causes a wheel well fire warning.
The wheel well fire detection system consists of four sensor elements, two circuit cards, a warning light and test
switch. The system utilizes 115-volts AC electrical power from the standby bus, and 28-volts DC from the battery
bus.
757 General Familiarization (7-2005)
ATA 26 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WHEEL WELL FIRE DETECTION SYSTEM COMPONENTS
Wheel Well Overheat Detectors
The continuous loop sensor is composed of four elements, two in each wheel well. The sensor loop in each wheel
well is mounted around the overhead by silicone bushings. The two elements begin on the inboard bulkhead, and
are connected together at the outboard side via the forward and aft bulkheads. The elements are connected via
wiring to the other wheel well elements and to the control card (P50).
The overheat sensors contain a thermistor material. The resistance of this material decreases as the temperature
of the material increases. At a temperature greater than 400°F | 20°F, the sensor provides an alarm signal
(ground) to the control card in the P50 card file. When the temperature of material cools sufficiently, the alarm
signal is automatically cancelled.
757 General Familiarization (7-2005)
ATA 26- Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-31
757 General Familiarization (7-2005)
ATA 26 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Duct Leak and Wheel Well Fire Circuit Cards
A power supply card, and control card located in P50 serve the functions of the wheel well fire detection system.
These same cards, also serve the functions of the duct leak detection system.
The power supply card receives 28-volts DC from the battery bus, and provides regulated 5-volts DC, 19-volts DC,
and 20-volts DC to the control card.
The control card contains a microprocessor and memory which performs fault detection and BITE functions. If a
fault is detected, a record is made in memory, for recall as required. A cover plate is installed on the card. The
he
cover plate contains a LED display and four BITE pushbutton controls, identified as DISP TEST, LOC TEST, MEM
CLR and MEM READ.
IF 26-32
757 General Familiarization (7-2005)
ATA 26- Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fire/Overheat Test Panel
The fire/overheat test panel is on the pilots' control stand P8. The test panel contains the wheel well test switch
and switches and relays associated with other fire and overheat detection systems.
IF 26-33
757 General Familiarization (7-2005)
ATA 26 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Wheel Well Fire Light
The WHEEL WELL FIRE light is on the first officer's instrument
rument panel P3. The light provides a red annunciation
when on.
IF 26-34
757 General Familiarization (7-2005)
ATA 26- Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WING AND BODY DUCT LEAK DETECTION SYSTEM
Overview
The wing and body duct leak detection system monitors the temperature along the left and right wing bleed air
duct zones, the air conditioning bays and the APU duct zone. When an overheat condition is detected in any of
these zones, a caution alert is provided in the flight compartment.
The system is divided into left and right zones, each containing dual detection elements set to trigger an alarm at
255°F. The left zone has five elements that are wired in series and extend, along the left wing leading edge, the
left air conditioning bay, the wheel well, the aft cargo compartment, and the pressure bulkhead aft to the APU. The
right zone has two elements that are wired in series and extend along the right wing leading edge, and the right air
conditioning bay.
The two zones are wired to the duct leak and wheel well fire control card in P50 electrical systems card file. If an
overheat is sensed, the appropriate duct leak caution light on pneumatic control panel (P5) lights, and a level B
caution appears on EICAS. A system test switch is provided on the miscellaneous test panel on P61 sidewall panel.
The switch is labeled "DUCT LEAK". A successful test, lights the two duct leak lights on pneumatic control panel
(P5). BITE is provided on the control card at P50 to assist with trouble shooting, and a DUCT LEAK BITE message
displays on the EICAS maintenance page if a fault exists.
The wing and body duct leak overheat detection system uses 115-volts AC, and 28-volts DC power.
757 General Familiarization (7-2005)
ATA 26 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WING AND BODY DUCT LEAK DETECTION SYSTEM COMPONENTS
Duct Leak Detectors
Detector elements consist of a fine nickel wire surrounded by a heat sensitive insulation, enclosed in a steel tube.
Electrical connectors are attached to each end. The resistance of the heat sensitive insulation decreases as the
temperature of the material increases. At a pre-selected temperature (resistance value), the detector provides an
alarm signal (ground) to the detector card in the P50 card file. When the temperature of material cools sufficiently,
the alarm signal is automatically cancelled.
IF 26-35
757 General Familiarization (7-2005)
ATA 26- Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 26-36
757 General Familiarization (7-2005)
ATA 26 - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Duct Leak and Wheel Well Fire Circuit Cards
A power supply card, and control card located in P50 serve the functions of the duct leak detection system. These
same cards, also serve the functions of the wheel well fire detection system.
The power supply card receives 28-volts DC from the R bus and provides regulated 5-volts DC, 19-volts DC, and
20-volts DC to the control card.
The control card consists of a microprocessor board, relay board, and display board. It performs monitor and
control functions of the system.
1. The microprocessor board contains a microprocessor and memory.
2. The relay board contains the sensor loop configuration relays.
3. The display board acts like a cover plate, and contains a LED display and four BITE pushbutton controls.
The pushbuttons are identified, DISP TEST, LOC TEST, MEM CLR, and MEM READ.
IF 26-37
757 General Familiarization (7-2005)
ATA 26- Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Bleed Air Control Panel
The M10259 bleed air control panel is on the overhead panel
nel P5. The panel contains the L WING DUCT LEAK and R
WING DUCT LEAK lights. The lights provide an amber annunciation when on. The panel also contains the controls
for the pneumatic system.
IF 26-38
757 General Familiarization (7-2005)
ATA 26 - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Miscellaneous Test Panel
The M10398 miscellaneous test panel is located on the right side panel P61. This panel contains a large number of
test switches for various systems. The DUCT LEAK system test switch is on this panel.
IF 26-39
757 General Familiarization (7-2005)
ATA 26- Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LAVATORY SMOKE DETECTOR SYSTEM
Overview
Smoke detectors are installed in each lavatory to provide early warning of the presence of smoke.
The units are 28-volts DC photoelectric type, which operate on the light scatter principle that utilizes a pulsing LED
light source and photodiode sensor in a fully screened sensing chamber. Every 4 to 5 seconds the pulsing LED
emits an infrared beam that bypasses the photodiode under normal conditions. When smoke enters the sensing
chamber, the infrared beam is deflected into the sensor by the smoke particles. The LED pulse rate is increased to
8 times the normal rate. When the photodiode confirms the smoke is present in the chamber for 2 consecutive
pulses, it will produce the signal necessary to trip the alarm.
NOTE:
The lavatory smoke detectors may not be triggered by cigarette smoke.
Functional Test
WARNING:
HORN IS RATED AT 90 db, USE HEARING PROTECTION.
Ensure power is on (LED on face of the detector blinks every 16-20 seconds). Remove slotted test cap (or use
needle nose to grip) turn test knob counterclockwise. Horn will sound in less than 6 seconds. Return knob to
normal, horn will silent within 6 seconds.
IF 26-40
757 General Familiarization (7-2005)
ATA 26 - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LAVATORY WASTE COMPARTMENT AUTOMATIC FIRE EXTINGUISHING SYSTEM
Overview
The lavatory waste compartment automatic fire extinguishing system provides fire extinguishing capability to the
lavatory waste compartment. Extinguishing occurs by flooding the lavatory waste compartment with an inert gas.
Fire Extinguisher Bottle
The fire extinguisher bottle includes an elongated spherical steel container with discharge tubes (fusible tips) and
mounting bracket. The container is about 2.5 inches in diameter with volume of approximately 10 cubic inches. The
container is filled with a Halon extinguishing agent, which leaves no residue when discharged. The fire extinguisher
bottle is mounted inside the lavatory cabinet assembly on the waste disposal chute. The two discharge tubes
extend into the waste container.
Temperature Indicator
The temperature indicator is a thin vinyl plate containing four heat sensitive patches. Each patch will change color
from grey to black when exposed to temperatures from 180° to 250°F. The temperature indicator, with selfadhesive backing, is located inside the lavatory cabinet assembly below the extinguisher bottle discharge tubes.
IF 26-41
757 General Familiarization (7-2005)
ATA 26- Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PORTABLE FIRE EXTINGUISHING SYSTEM
Overview
Portable fire extinguishers are installed in several general locations in the passenger compartment. If the
extinguisher is not easy to see, the location will be identified with a placard. The extinguishers are usually installed
in one or more of these areas: galley or lavatory stowage areas, closets, doghouses, or near attendant seats. A fire
extinguisher is also installed in the flight compartment.
The portable extinguishers are attached to wall-mounted brackets by quick-release mounting straps.
757 General Familiarization (7-2005)
ATA 26 - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PORTABLE FIRE EXTINGUISHING SYSTEM COMPONENTS
Halon Extinguishers
Halon extinguishers are used to extinguish electrical and flammable liquid fires.
The Halon extinguisher is rechargeable. A pressure gage shows when you must recharge or replace the fire
extinguisher.
To operate the extinguisher, pull the handle locking pin. Hold the extinguisher upright and squeeze the handle and
lever together. Point the nozzle flow at the base of the fire. The Halon extinguishing agent leaves no residue after
discharge.
IF 26-42
757 General Familiarization (7-2005)
ATA 26- Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pressurized Water Extinguishers
Pressurized water extinguishers are used to extinguish non-electrical fires.
The water extinguisher is rechargeable. Antifreeze is added to the water to prevent freezing.
A carbon dioxide cartridge is mounted on the handle of the extinguisher. To operate the extinguisher, turn the
cartridge. This punctures the cartridge and pressurizes the water container. Push the trigger and aim the nozzle
flow at the base of the fire.
IF 26-43
757 General Familiarization (7-2005)
ATA 26 - Page 67
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 27
FLIGHT CONTROLS ................................................................................................................................ 7
Introduction......................................................................................................................................... 7
Primary Flight Controls ...................................................................................................................... 7
Secondary Flight Controls .................................................................................................................. 7
Flight Control Actuators, Servos, and Electronics ............................................................................. 8
Flight Controls - Hydraulic Distribution .......................................................................................... 10
FLIGHT CONTROL SYSTEM ELECTRONICS UNIT (CSEU) ........................................................... 12
Overview........................................................................................................................................... 12
AILERON AND AILERON TRIM CONTROL SYSTEM ..................................................................... 14
Overview........................................................................................................................................... 14
OPERATION ............................................................................................................................................ 16
Manual Operation ............................................................................................................................. 16
Abnormal Operation ......................................................................................................................... 16
Autopilot Operation .......................................................................................................................... 16
AILERON AND AILERON TRIM CONTROL SYSTEM COMPONENTS ......................................... 18
Lateral Control Wheel ...................................................................................................................... 18
Aileron Control Drum and Bus Force Limiter.................................................................................. 19
Lateral Feel, Centering, and Trim Mechanism ................................................................................. 20
Aileron Trim Switches ...................................................................................................................... 21
Aileron Trim Actuator ...................................................................................................................... 21
Lateral Control Override Mechanism ............................................................................................... 22
Aileron Quadrant and Override Mechanism..................................................................................... 23
Power Control Actuators (PCAs) ..................................................................................................... 24
Aileron .............................................................................................................................................. 25
AILERON POSITION INDICATING SYSTEM .................................................................................... 26
Operation .......................................................................................................................................... 26
AILERON POSITION INDICATING SYSTEM COMPONENTS ........................................................ 27
Flight Control Surface Position Indicator ......................................................................................... 27
Aileron Position Transmitter ............................................................................................................ 28
RUDDER AND RUDDER TRIM CONTROL SYSTEM ....................................................................... 29
Overview........................................................................................................................................... 29
RUDDER AND RUDDER TRIM CONTROL SYSTEM COMPONENTS ........................................... 30
Rudder Pedals ................................................................................................................................... 30
Forward Quadrant and Jackshaft Assembly ..................................................................................... 30
Pedal Adjustment Crank ................................................................................................................... 30
Rudder Control Cables...................................................................................................................... 32
Aft Quadrant, Feel, Centering and Trim Mechanism ....................................................................... 34
Rudder Trim Actuator....................................................................................................................... 34
Rudder Thermal Compensating Linkage .......................................................................................... 36
Autopilot Rollout Guidance Servos .................................................................................................. 38
Rudder Ratio Changer Mechanism ................................................................................................... 40
757 General Familiarization (7-2005)
ATA 27 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Ratio Changer Actuator (RRCA) ......................................................................................... 40
Rudder Ratio Changer Bypass Valve ............................................................................................... 40
Linear Variable Differential Transformer (LVDT) .......................................................................... 40
Primary Control Linkage - Load Limiting........................................................................................ 40
Secondary Control Linkage .............................................................................................................. 40
Yaw Damper Summing Lever .......................................................................................................... 43
Yaw Damper Servos ......................................................................................................................... 43
Power Control Actuators (PCAs) ..................................................................................................... 44
PCA Reaction Link ........................................................................................................................... 44
PCA Linkages ................................................................................................................................... 44
Rudder Structure ............................................................................................................................... 46
Rudder Trim Control Switch and Indicator ...................................................................................... 47
RUDDER POSITION INDICATING SYSTEM ..................................................................................... 48
Operation .......................................................................................................................................... 48
RUDDER POSITION INDICATING SYSTEM COMPONENTS ......................................................... 49
EICAS Position Indicator ................................................................................................................. 49
Position Transmitter.......................................................................................................................... 50
ELEVATOR CONTROL SYSTEM ......................................................................................................... 52
Overview........................................................................................................................................... 52
ELEVATOR CONTROL SYSTEM COMPONENTS ............................................................................ 54
Control Column ................................................................................................................................ 54
Elevator Control Aft Quadrants ........................................................................................................ 56
Elevator Control Aft Mechanism Torque Box ................................................................................. 56
Elevator Feel Unit and Actuator ....................................................................................................... 56
Neutral Shift and Override Mechanism ............................................................................................ 56
Asymmetry Limiter Actuator............................................................................................................ 56
Elevator Power Control Actuator ..................................................................................................... 58
Elevator ............................................................................................................................................. 58
Elevator Feel Computer .................................................................................................................... 58
Elevator Pressure Reducer and Bypass Valves................................................................................. 60
STALL WARNING SYSTEM ................................................................................................................. 61
Overview........................................................................................................................................... 61
STALL WARNING COMPONENTS...................................................................................................... 62
Control Column Shakers................................................................................................................... 62
Stall Warning Module....................................................................................................................... 63
ELEVATOR POSITION INDICATING SYSTEM ................................................................................. 64
Overview........................................................................................................................................... 64
ELEVATOR POSITION INDICATING COMPONENTS ..................................................................... 65
EICAS Position Indicator ................................................................................................................. 65
Elevator Position Transmitter ........................................................................................................... 66
HORIZONTAL STABILIZER TRIM CONTROL SYSTEM ................................................................. 67
Overview........................................................................................................................................... 67
HORIZONTAL STABILIZER TRIM CONTROL SYSTEM COMPONENTS ..................................... 70
Stabilizer Trim Control Wheel Switches .......................................................................................... 70
757 General Familiarization (7-2005)
ATA 27- Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STAB TRIM Control Levers .................................................................................................................... 72
Stabilizer Trim Control Module (STCM) ......................................................................................... 72
Stabilizer Trim Ball screw Actuator ................................................................................................. 74
Stabilizer Trim Drive Hydraulic Motor ............................................................................................ 76
Stabilizer Trim Secondary Brake...................................................................................................... 77
Stabilizer Trim Limit Switch and Position Transmitter Modules (LSTMs)..................................... 78
Stabilizer Trim Cut-off Switches ...................................................................................................... 79
Stabilizer Trim Shutoff Switches...................................................................................................... 80
Stabilizer/Elevator Asymmetry Limiter Module (SAM) .................................................................. 81
STABILIZER TRIM POSITION INDICATING SYSTEM .................................................................... 82
Operation .......................................................................................................................................... 82
STABILIZER TRIM POSITION INDICATING SYSTEM COMPONENTS ........................................ 83
Stabilizer Position Indicators ............................................................................................................ 83
B. Stabilizer Trim Limit Switch and Position Transmitter Module (LSTM) ................................... 84
Stabilizer Position Modules (SPMs) ................................................................................................. 86
TRAILING EDGE FLAP SYSTEM ........................................................................................................ 87
Overview........................................................................................................................................... 87
Trailing Edge Flap Primary Control ................................................................................................. 89
Trailing Edge Flap Alternate Control ............................................................................................... 89
Primary Power .................................................................................................................................. 89
Alternate Power ................................................................................................................................ 89
Flap Drive System ............................................................................................................................ 90
Flap Load Relief ............................................................................................................................... 90
Depressurization ............................................................................................................................... 90
Asymmetry........................................................................................................................................ 90
TRAILING EDGE FLAP SYSTEM COMPONENTS ............................................................................ 94
Flap Control Lever ............................................................................................................................ 94
Alternate Control Panel .................................................................................................................... 95
Flap Power Drive Unit (PDU) .......................................................................................................... 96
Flap/Slat Depressurization Module .................................................................................................. 98
Flap Torque Tubes ............................................................................................................................ 99
Flap Transmissions ......................................................................................................................... 100
Flap Carriages ................................................................................................................................. 101
Fairings ........................................................................................................................................... 102
Flap/Slat Electronic Unit (FSEU) ................................................................................................... 104
TRAILING EDGE FLAP POSITION INDICATING SYSTEM........................................................... 106
Overview......................................................................................................................................... 106
TRAILING EDGE FLAP POSITION INDICATING SYSTEM COMPONENTS .............................. 108
Flap/Slat Position Indicator ............................................................................................................ 108
Flap/Slat Electronic Unit (FSEU) ................................................................................................... 109
Engine Indicating and Crew Alerting System (EICAS) ................................................................. 110
Resolvers......................................................................................................................................... 111
Synchros.......................................................................................................................................... 111
SPOILERS AND DRAG DEVICES ...................................................................................................... 112
Overview......................................................................................................................................... 112
757 General Familiarization (7-2005)
ATA 27 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
SPOILER/SPEEDBRAKE CONTROL SYSTEM................................................................................. 113
Overview......................................................................................................................................... 113
Operation ........................................................................................................................................ 114
Spoiler/Speed brake System ........................................................................................................... 114
SPOILER/SPEEDBRAKE SYSTEM COMPONENTS ........................................................................ 116
Rotary Variable Differential Transducers (RVDTs) ...................................................................... 116
Spoiler Control Modules (SCMs) ................................................................................................... 117
Power Control Actuators (PCAs) ................................................................................................... 118
Electro hydraulic Servo Valves (EHSVs) ....................................................................................... 119
Spoiler Panels ................................................................................................................................. 120
Speed brake Lever and LVDTs ...................................................................................................... 121
Flight Deck Indications................................................................................................................... 122
AUTO-SPEEDBRAKE CONTROL SYSTEM...................................................................................... 123
Overview......................................................................................................................................... 123
AUTO-SPEED-BRAKE SYSTEM COMPONENTS ............................................................................ 124
Auto-Speed-brake Switches............................................................................................................ 124
Auto-Speed-brake Actuator and No-Back Clutch .......................................................................... 125
Reverse Thrust Mechanism ............................................................................................................ 126
Forward Thrust Lever Position Switches........................................................................................ 127
LEADING EDGE SLAT SYSTEM ....................................................................................................... 128
Overview......................................................................................................................................... 128
L.E. Slat Primary Control ............................................................................................................... 129
L.E. Slat Alternate Control ............................................................................................................. 129
Primary Power ................................................................................................................................ 129
Alternate Power .............................................................................................................................. 129
Drive System .................................................................................................................................. 130
Autoslats ......................................................................................................................................... 130
Slat Loss Detection Switches.......................................................................................................... 130
FSEU-1, FSEU-2, and FSEU-3 ...................................................................................................... 130
LEADING EDGE SLAT SYSTEM COMPONENTS ........................................................................... 132
Slat PDU ......................................................................................................................................... 132
Hydraulic Control Valve Module ................................................................................................... 132
Hydraulic Motor ............................................................................................................................. 132
Electric Motor ................................................................................................................................. 132
Gearbox........................................................................................................................................... 134
Position Transmitter........................................................................................................................ 134
Shock Absorber............................................................................................................................... 134
Slat Bypass Valve ........................................................................................................................... 135
Flap/Slat Depressurization Module ................................................................................................ 136
Slat Drive Torque Tubes ................................................................................................................. 137
Angle Gearboxes............................................................................................................................. 138
Rotary Actuators and Pinion Gears ................................................................................................ 138
Slats................................................................................................................................................. 139
Slat Loss Detection ......................................................................................................................... 140
FSEU-1, FSEU-2, and FSEU-3 ...................................................................................................... 142
757 General Familiarization (7-2005)
ATA 27- Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LEADING EDGE SLAT POSITION INDICATING SYSTEM ........................................................... 145
Overview......................................................................................................................................... 145
Flap/Slat Position Indication on Primary Power............................................................................. 145
Flap/Slat Position Indication on Alternate Power........................................................................... 147
Takeoff Warning ............................................................................................................................. 147
Slat Position Transmitters ............................................................................................................... 148
757 General Familiarization (7-2005)
ATA 27 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27- Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FLIGHT CONTROLS
Introduction
Aerodynamic characteristics of the 757 have been tailored to reduce drag and provide high performance. Single
leading edge slats and trailing edge flaps are used for high aerodynamic efficiency, functional reliability, and low
maintenance requirements. Hydraulically powered conventional flight controls with redundant systems provide failsafe control throughout the flight envelope.
Primary Flight Controls
The primary flight controls of the airplane are the ailerons, elevators, and rudder. Each control surface is powered
by power control actuators (PCAs) that are cable operated. Each PCA receives hydraulic power from a separate
hydraulic system. The primary flight controls use all three hydraulic systems for redundancy. There is no reversion
to manual control of the surfaces if hydraulic power fails.
Secondary Flight Controls
The secondary flight controls are the spoiler/speed brakes, horizontal stabilizer, leading edge (LE) slats, and
trailing edge (TE) flaps.
Six spoiler/speed brakes are installed on each wing. They are hydraulically powered and electronically controlled.
The spoiler/speed brakes receive hydraulic power from all three hydraulic systems.
Five LE slats are installed on each wing. The slats are powered either hydraulically or electrically. Normally,
hydraulic motors rotate drive shafts that drive the rotary actuators. Electric motors power the drive shafts as a
backup. The slats receive hydraulic power from the left system.
Two TE flaps (inboard and outboard) are installed on each wing. The flaps are powered either hydraulically or
electrically. Normally, hydraulic motors rotate drive shafts that drive rotary actuators. Electric motors power the
drive shafts as a backup. The flaps receive hydraulic power from the left system.
The horizontal stabilizer is hydraulically powered and manually or electrically controlled. The stabilizer receives
hydraulic power from the right and center systems.
757 General Familiarization (7-2005)
ATA 27 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flight Control Actuators, Servos, and Electronics
Three autopilot lateral control servos (ALCS) provide automatic control. Each ALCS is powered by a separate
hydraulic system. The ALCSs have internal servovalves for autopilot inputs. All of the ALCSs have the same part
number.
A total of twenty-five power control actuators move the control surfaces. Each PCA uses one of the three hydraulic
systems.
Six autopilot servos provide electronic control inputs to the flight control systems. There are three elevator servos
and three directional servos. The elevator and directional servos use all three hydraulic systems. Each hydraulic
system supplies one elevator servo and one directional servo. The two pitch augmentation servos are also supplied
by separate hydraulic systems.
Two yaw damper servos provide gust damping inputs to the rudder system. Each servo uses one of two hydraulic
systems. The two yaw damper servos have the same part number.
The control system electronics unit (CSEU) is a group of power supply and electronic modules used in the flight
control system. There are two (left and right) CSEUs installed for redundancy.
757 General Familiarization (7-2005)
ATA 27- Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-1
757 General Familiarization (7-2005)
ATA 27 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flight Controls - Hydraulic Distribution
Three independent hydraulic systems provide power to the flight controls. Each system uses a combination of
pumps to supply power.
Flight control redundancy is provided by the use of multiple hydraulic systems. The aileron, elevator, rudder, and
spoiler PCA's use all three hydraulic systems. The elevator and directional servos and ALCS's each use one of the
three hydraulic systems. The slats, flaps, and rudder ratio changer use a single hydraulic system.
The ram air turbine (RAT) is an emergency hydraulic source. The RAT is used in the event of a double engine
failure. The RAT is used by the primary flight control system only.
The tail shutoff valves are used for system isolation. The valves are provided for ground use only. The valves are
normally open in flight.
Safety harness attachment receptacles are provided on the wing and horizontal stabilizer upper surfaces for use by
maintenance personnel working high above ground.
757 General Familiarization (7-2005)
ATA 27- Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-2
757 General Familiarization (7-2005)
ATA 27 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FLIGHT CONTROL SYSTEM ELECTRONICS UNIT (CSEU)
Overview
The Control System Electronics Unit (CSEU) is a collection of flight control modules and four power supply
modules. The CSEU is energized by its own power supply modules.
The CSEU interfaces with other primary airplane systems to provide the computation and control functions for yaw
damping, stabilizer trim, elevator travel limit, rudder control authority, and spoiler deployment.
The system consists of independent left and right CSEUs. All modules are line replaceable units (LRU). Equivalent
modules can be changed between equivalent shelf positions on the left and right CSEUs.
The left CSEU is on the top E3-1 shelf, and the right CSEU is on the top E4-1 shelf. These shelves are found in the
main electronics equipment center. Access is thru the electronics access door, 119BL, found on the bottom of the
fuselage, aft of the nose wheel well.
The CSEU consists of these modules (and equipment numbers):
1. Yaw damper module (YDM) (M522 LH; M523 RH)
2. Rudder ratio changer module (RRC) (M528 LH; M529 RH)
3. Stabilizer trim/elevator asymmetry limit module (SAM) (M524 LH; M525 RH)
4. Spoiler control module (SCM) (M530, M531, M532 LH; M533, M534, M535 RH)
5. Power supply module (PSM) (M536, M537 LH; M538, M539 RH)
The CSEU modules provide the electrical interface between the sensors and the signal sources. These include the
air/ground logic sensors, the hydraulic pressure sense signals, the pilot's and first officer's flight control
transducers, and the control surfaces transducers and actuators.
Built-In-Test Equipment (BITE) in the CSEU modules supply continuous monitor features specified for the monitor
circuits. When the CSEU is powered, the background is continuously monitored for faults, and the in-flight failure
data is stored.
757 General Familiarization (7-2005)
ATA 27- Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-3
757 General Familiarization (7-2005)
ATA 27 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AILERON AND AILERON TRIM CONTROL SYSTEM
Overview
The ailerons, assisted by the spoilers, provide lateral control of the airplane. There are two ailerons; one at the
outboard trailing edge of each wing. Each aileron is moved by two power control actuators (PCAs) which move the
aileron up and down.
Control wheels in the flight deck regulate aileron movement. The control wheels actuate cables which in turn
actuate the PCAs. The PCAs are powered by separate hydraulic systems to allow aileron control in the event of a
single hydraulic failure.
The autopilot system controls the ailerons electrically. Autopilot commands actuate the PCAs and back-drive the
control wheels.
The aileron trim system operates the ailerons and control wheels electrically to adjust the ailerons to a neutral or
trimmed (wings level) position. The trim indicator on the control wheels shows the degree of trim.
757 General Familiarization (7-2005)
ATA 27- Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-4
757 General Familiarization (7-2005)
ATA 27 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
OPERATION
Manual Operation
Control wheel rotation provides the initial manual command inputs to the lateral control system. Control cables
relay commands to lateral control mechanisms in the left and right main gear wheel wells.
The feel, centering, and trim mechanism and the lateral control override mechanism receive cable command inputs
from the control wheel. Linked by a control rod, the mechanisms act together through control cables commanding
left and right ailerons equally, but in opposite directions.
The aileron quadrant receives cable commands from the left or right wheel well mechanisms. As the quadrant
rotates, it moves mechanical linkage to the PCAs. This causes the PCA servo valve to port hydraulic power to the
actuators. Under hydraulic pressure, the actuator extends or retracts and the aileron is raised or lowered as
required.
Abnormal Operation
Each aileron component has a breakout or override so that the aileron system will remain operable should it or a
downstream component become jammed.
The captain's and first officer's control wheels are joined by a bus force limiter rod. Should either control wheel
become inoperable, the force limiter's two-way spring allows the other wheel to operate the remaining aileron
system independently.
The feel, centering, and trim mechanism and the lateral control override mechanism each have an override or
shear joint to allow continued operation of one should the other become jammed.
The wing aileron override mechanism can allow break out of a PCA jam or an aileron quadrant jam. These
overrides ensure continued use of the remaining aileron.
Autopilot Operation
The autopilot lateral control servos (ALCSs) receive commands directly from the autopilot system. The ALCSs are
linked directly to the wheel well mechanisms by cranks. Autopilot commands will cause the cranks to move which
drive the wheel well mechanisms. Actual aileron control remains the same as with manual operation.
757 General Familiarization (7-2005)
ATA 27- Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-5
757 General Familiarization (7-2005)
ATA 27 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AILERON AND AILERON TRIM CONTROL SYSTEM COMPONENTS
Lateral Control Wheel
The control wheel drives a drum at the base of the control column to operate the aileron system. The bus force
limiter links the captain's and first officer's control drums to ensure symmetrical input. Control wheel rotation
actuates cables which run aft to lateral control mechanisms in the left and right main gear wheel wells.
Control wheel travel is 82.5° to the left and right of the control wheel neutral position. Maximum corresponding
aileron travel is 21° up and 21° down for about 55° control wheel travel. The remaining 30° of travel drive the
spoiler RVDT units. An autopilot disconnect button is located on the control wheel. The aileron trim indicator is on
the top of the control wheel.
IF 27-6
757 General Familiarization (7-2005)
ATA 27- Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aileron Control Drum and Bus Force Limiter
The aileron control drum houses the aileron control quadrant. The quadrant rotates body cables which drive the
lateral control mechanisms in the wheel wells for the left and right main landing gear. Connected to the drum are
spoiler RVDT units.
The bus force limiter links both control drums so that they operate in unison. It also provides a two-way breakout
in the event a control wheel becomes inoperable. The force limiter is spring-loaded in both directions so that either
control wheel can operate independently of the other, if necessary.
IF 27-7
757 General Familiarization (7-2005)
ATA 27 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Lateral Feel, Centering, and Trim Mechanism
The lateral feel, centering, and trim mechanism, located in the wheel well for the left main landing gear,
mechanically drive the left aileron PCAs. A control rod connects the mechanism to the lateral control override
mechanism in the wheel well for the right main landing gear. The feel, centering, and trim mechanism also has an
override cam to allow operation of the right aileron if the left side becomes jammed (the left side has a shearout).
Two centering springs keep the ailerons in a neutral position when there are no commanded inputs. Command
inputs are either manual from the control wheel and cables or hydraulic from the autopilot system. The autopilot
lateral control servo (ALCS) cranks input autopilot commands directly to the feel, centering, and trim mechanism.
IF 27-8
757 General Familiarization (7-2005)
ATA 27- Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aileron Trim Switches
The aileron trim switches are located on the aft electronic control panel (P8). The trim switches must be operated
together to provide electrical input to the trim actuator.
Aileron Trim Actuator
The aileron trim actuator is mounted on the support box and connected to the feel, centering, and trim mechanism.
The actuator receives electrical signals from the trim switches. The trim actuator extends or retracts, adjusting the
neutral position.
on.
IF 27-9
757 General Familiarization (7-2005)
ATA 27 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Lateral Control Override Mechanism
The lateral control override mechanism is located in the wheel well of the right main landing gear. A control rod
connects the lateral control override mechanism to the feel, centering, and trim mechanism so that left and right
ailerons respond equally to command inputs. A crank connects the override mechanism to the right ALCS to
provide autopilot control. Override springs and a cam allow operation of the left aileron should the right side
become jammed.
IF 27-10
757 General Familiarization (7-2005)
ATA 27- Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aileron Quadrant and Override Mechanism
The aileron quadrant and override mechanism is located inboard of the PCAs in the wing trailing edge. A control
rod connects the quadrant and override to the PCA input crank.
The aileron quadrant transmits control wheel or autopilot commands to the PCAs. The override mechanism
provides a breakout to allow continued use of one aileron should the other become jammed.
IF 27-11
757 General Familiarization (7-2005)
ATA 27 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Power Control Actuators (PCAs)
The PCAs are mounted side-by-side on the wing rear spar. Each PCA consists of an actuator with internal control
and shuttle valves. The actuator rod end is attached to a bracket on the aileron. The summing levers connect each
PCA with the aileron quadrant input crank. Reaction links provide the mechanical reaction path forces to raise and
lower the ailerons.
Each PCA is powered by a separate hydraulic system. The left aileron outboard PCA receives pressure from the
right hydraulic system; the right aileron outboard PCA receives pressure from the left hydraulic system. The left
and right aileron inboard PCAs are powered by the center hydraulic system.
IF 27-12
757 General Familiarization (7-2005)
ATA 27- Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aileron
The aileron is mounted on the wing trailing edge at five hinge points. The aileron has full 40° travel; 20 ± 1°
deflection down, 20 ± 1° deflection up. The autopilot provides full aileron deflection. Aileron trim is limited to
11.6° down and 11.6° up. The aileron position transmitter is mounted on the aileron surface and wing rear spar.
IF 27-13
757 General Familiarization (7-2005)
ATA 27 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AILERON POSITION INDICATING SYSTEM
Operation
The aileron position indicating system provides the flight crew with visual indication of aileron angular
displacement.
When the aileron deflects, it drives the control rod which provides an electrical signal in the transmitter. The
voltage is proportional to the amount of aileron movement. The EICAS computer receives the signal and displays
aileron position when the status button is pressed. Two pointers indicate aileron position separately.
Aileron position indicating is automatic when power is on. Both left and right bus power supplies the system when
the AILERON POS circuit breakers L and R are closed.
757 General Familiarization (7-2005)
ATA 27- Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AILERON POSITION INDICATING SYSTEM COMPONENTS
Flight Control Surface Position Indicator
The flight control surface position indicator is displayed on the status page of the lower EICAS screen. The
indicator shows the position of the ailerons, rudder, and elevators.
To display aileron position indicator, press STATUS button on pilot's select display panel P9.
IF 27-14
757 General Familiarization (7-2005)
ATA 27 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aileron Position Transmitter
The position transmitters are accessible through panels on the wing lower surface. Adjustable control rods attach
the transmitters to the ailerons. The transmitter is rigged to electrical zero when the aileron is faired with wing
structure.
IF 27-15
757 General Familiarization (7-2005)
ATA 27- Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RUDDER AND RUDDER TRIM CONTROL SYSTEM
Overview
The rudder and rudder trim control system provides directional control and stability around the vertical axis.
The captain's and first officer's rudder pedals provide rudder control through a cable and linkage system to the
power control actuators (PCAs) at the rudder. An electric actuator which repositions the feel and centering springs
provides rudder trim control about system centered position. Trim switches on the aft electronic control panel P8
control the trim actuator. Three directional autopilot servos are also connected to the input controls. A rudder ratio
changer system in the input linkage reduces rudder control commands as a function of airspeed. Yaw damper
servo outputs, summed with other control inputs, provide a command signal to the PCAs.
The rudder is powered by the three hydraulic PCAs. Each PCA is independently supplied with pressure by one of
three separate hydraulic systems. Any one PCA is capable of providing the required rudder travel.
Rudder position is indicated on the EICAS status page from the rudder position transmitter. A trim position
indicator next to the rudder trim switch on aft electronic control panel P8 displays the amount of rudder trim. An
amber RUDDER RATIO caution light and EICAS message identifies an inoperative ratio changer system. A yaw
damper INOP light for each yaw damper indicates
ndicates yaw damper inoperative or OFF.
IF 27-16
757 General Familiarization (7-2005)
ATA 27 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RUDDER AND RUDDER TRIM CONTROL SYSTEM COMPONENTS
Rudder Pedals
Two sets of rudder pedals are located below the captain's and first officer's instrument panel. Each pair have rig
pin holes R1 located in the pedal arms below the flight deck floor. Access to this area is through access door
(113AL) forward of the nose gear wheel well. Pedal travel is limited to about 3.7 inches from neutral in each
direction by stops on the two forward quadrant and jackshaft assemblies.
Forward Quadrant and Jackshaft Assembly
An adjustable yoke is mounted on the aft side of each forward quadrant. The yoke is attached to the pedals by two
control/push rods and is operated by a cable-driven jackshaft to provide fore and aft pedal adjustment. A bus rod
ties both forward quadrants together, so either set of pedals provides input to both quadrants.
The quadrants are the forward terminal for the rudder cables. The left quadrant has stop bumpers that contact
structural stops to limit pedal movement. The left quadrant also contains rig pin hole R2 which is used to rig the
rudder pedals and set cable tension.
Pedal Adjustment Crank
The captain's and first officer's adjustment cranks are found on the pedal cover front panel. The cranks adjust
pedals independently of one another and are detented to retain position. Each crank turns a flex cable which drives
the jackshaft yoke.
757 General Familiarization (7-2005)
ATA 27- Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-17
757 General Familiarization (7-2005)
ATA 27 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Control Cables
Two control cables are attached at the left and right forward quadrants under the flight deck floor. The cables run
under the passenger compartment and through the aft pressure bulkhead, up to the aft quadrant.
Three turnbuckles per cable are used for adjusting cable tension with rig pin R2 and R3 installed. Access to the
forward set of turnbuckles is through the forward cargo compartment ceiling. The center set is accessed through
the cabin floor and the aft set through access door 311AL.
IF 27-18
757 General Familiarization (7-2005)
ATA 27- Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aft Quadrant, Feel, Centering and Trim Mechanism
The aft quadrant, feel, centering and trim mechanism is located at the base of the vertical fin, above the autopilot
servos. All mechanism inputs and outputs are connected to its offset torque tube. The aft quadrant transfers cable
inputs from the rudder pedals to the torque tube. The quadrant also back-drives the cable system in response to
rudder movement caused by autopilot and trim inputs. Other input sources are the trim actuator and three
directional autopilot servos. The torque tube output consists of two control rods/pogos which provide ratio changer
mechanism input.
Two aft quadrant stops limit rotation of the aft quadrant
Feel and centering functions are provided by a cam-roller-spring mechanism. It provides an increasing feel force to
rudder pedal input and a centering force to return pedals and rudder to neutral when pedal deflection is removed.
Rudder Trim Actuator
The housing of the rudder trim actuator is attached to the vertical fin rear spar and the rod end is connected to the
feel, centering and trim mechanism. Trim commands from the trim switch cause the actuator to extend or retract,
which rotates the feel, centering and trim mechanism. This provides a new zero force pedal position corresponding
to the trimmed rudder position. The actuator provides the ground point for the feel, centering and trim mechanism.
The trim indicator shows units of trim, since rudder trim in degrees is dependent on ratio changer position. The
maximum trim authority is ± 21.9 degrees (rudder travel rate of 1.1 degrees per second) and minimum trim
authority is ± 2.5 degrees (rudder travel rate of 0.13 degrees/sec). Actuator stroke is ±1.37 inches and needs
about 16.7 seconds for full travel in one direction.
757 General Familiarization (7-2005)
ATA 27- Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-19
757 General Familiarization (7-2005)
ATA 27 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Thermal Compensating Linkage
During climb, structure cools faster than rudder control linkage. This causes rudder to shift left due to different
expansion rates of the structure and rudder control linkage, requiring right rudder trim to compensate. As
temperatures equalize during cruise, rudder trim is gradually removed, leaving only the normal cruise trim
requirements. Opposite conditions occur during descent, as structure warms faster than rudder control linkage.
The rudder thermal compensating linkage eliminates the need for manual rudder trim in such situations. The lower
compensating rod expands/contracts at the same rate as the rudder control linkage. Therefore, when the structure
is expanding/contracting faster than the rudder linkage, the compensating linkage repositions the input crank,
providing input to the yaw damper summing lever to counteract rudder deflection due to different
expansion/contraction rates between structure and rudder control linkage.
IF 27-20
757 General Familiarization (7-2005)
ATA 27- Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-20A
757 General Familiarization (7-2005)
ATA 27 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Autopilot Rollout Guidance Servos
The three autopilot rollout guidance servos, located directly below the feel, centering and trim mechanism, are
driven by the flight control computers. These servos are normally disengaged and are used for autoland rollout
guidance only.
IF 27-21
757 General Familiarization (7-2005)
ATA 27- Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Ratio Changer Mechanism
The rudder ratio changer mechanism is mounted in the vertical fin directly above the aft quadrant feel, centering
and trim mechanism. The mechanism limits the rudder travel as a function of airspeed; as airspeed increases,
available rudder travel decreases. The ratio changer output is driven by an electro hydraulic actuator and
controlled by signals from one of two rudder ratio changer modules (RRCM).
The rudder ratio changer modules (RRCM) are located in the main equipment center on shelves E3 and E4. The
modules receive 28-volts DC from bus L and are controlled by the RUD RATIO circuit breaker located on overhead
circuit breaker panel P11.
Rudder Ratio Changer Actuator (RRCA)
The RRCA is located on top and to the left of the ratio changer. The RRCA is attached to the ratio changer support
bracket mounted on the vertical fin aft spar. Inputs to the RRCA consist of hydraulic pressure from the left
hydraulic system and electrical signals to the solenoid valve and EHSV. Outputs from the RRCA are hydraulic
pressure to the middle power control actuator (PCA) and mechanical drive to the left ratio changer crank.
Rudder Ratio Changer Bypass Valve
Loss of the rudder ratio changer actuator (RRCA) normally causes left hydraulic system power to be cut off from
the middle power control actuator (PCA). Loss of the center and right hydraulic systems causes the loss of the
lower and upper PCAs, respectively. The bypass valve ensures rudder control by bypassing the RRCA and
supplying left hydraulic system power to the middle PCA in the event of ratio changer shutdown and loss of right
hydraulic system.
Linear Variable Differential Transformer (LVDT)
An LVDT is mounted on the ratio changer and driven by the right crank. Dual channels provide position feedback
signals to the two rudder ratio changer modules (RRCM). The left LVDT channel supplies signals to the left RRCM
and the right channel supplies the right RRCM. The LVDT has an adjustment nut for rigging purposes. The LVDT is
powered by 26-volts AC from PSM 1R and 1L.
Primary Control Linkage - Load Limiting
A shearout is located on the right ratio changer crank. The shearout protects the primary control linkage between
the ratio changer and yaw damper summing lever in the event of a jam.
Secondary Control Linkage
The secondary control linkage includes a pogo between the left ratio changer crank and the yaw damper summing
lever. The ratio changer output pogo limits the force fight in the two control paths due to component tolerances
and is slightly biased to reduce lost motion by preloading the linkage joints. Rod end bearings are called out on the
pogo rod decals
757 General Familiarization (7-2005)
ATA 27- Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-22
757 General Familiarization (7-2005)
ATA 27 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27- Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Yaw Damper Summing Lever
The summing levers are located directly above the ratio changer mechanism. The levers sum control inputs from
the ratio changer together with yaw damper servo inputs and outputs to the PCA control rods. The torque tube has
four connecting lugs for the primary and secondary control path summing levers and two crank arms for the yaw
damper servo input linkage and the yaw damper bias pogo.
The yaw damper pogo is parallel with the yaw damper servo inputs. It provides a ground point for the control
linkage in case of yaw damper servo input linkage failure and is slightly biased to remove linkage backlash.
Yaw Damper Servos
Two servos are mounted side-by-side on the vertical fin aft spar. Their purpose is to provide rudder input
commands for dutch roll damping and gust load relief. Left and center hydraulic systems power the servos.
IF 27-23
757 General Familiarization (7-2005)
ATA 27 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Power Control Actuators (PCAs)
The three rudder PCAs are interchangeable. Each PCA is controlled by primary control path inputs and the middle
and lower PCAs receive secondary control path inputs. The upper PCA is powered by the right hydraulic system,
the middle PCA by the left system (through the RRCA) and the lower PCA is powered by the center system.
The PCAs are mounted parallel in a vertical row above the yaw damper summing lever, between the vertical fin
rear spar and rudder front spar. The PCA cylinder base is mounted on a trunnion block connected to the vertical
fin rear spar; the rod end is mounted to the rudder front spar.
PCA Reaction Link
The reaction link provides mechanical feedback from rudder surface movement to reposition the actuator body and
null out the valve input. One end is connected to the rudder front spar on the right side of the rudder pivot. The
other end is connected to the trunnion assembly. Curved slots on top and bottom of the trunnion assembly provide
retention for the PCA in the event of a reaction link or hanger link failure. The trunnion right side is connected to
and supported by the bearing mounted hanger on vertical fin rear spar.
PCA Linkages
The PCA linkages transfer inputs from the yaw damper summing lever to the actuators through control rods and
torque tube levers. The primary control path (right side) provides input to the middle PCA torque tube lever.
Control input is continued through the primary control path to the lower and upper PCAs via non-adjustable rods.
The secondary control path (left side) connects the summing lever to the lower PCA torque tube lever and is nonadjustable. Shearout rivets are provided at each PCA input valve lever to protect against a jam in the PCA control
valve or linkage.
757 General Familiarization (7-2005)
ATA 27- Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-24
757 General Familiarization (7-2005)
ATA 27 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Structure
The rudder is constructed of graphite epoxy over a nomex honeycomb. The rudder weighs about 334 pounds and is
about 30 feet tall.
Eight hinge fittings are mounted between the vertical fin rear spar and the rudder front spar. Two thrust hinges
provide rudder positioning. Six hinges are expansion link hinges needed to account for the difference in thermal
expansion between the graphite rudder and aluminum hinges. A seal between the rudder front spar and fin rear
spar prevents airflow across the hinge line at all deflection angles. Hoisting points are located at the rudder front
spar between the bottom two hinges and between the sixth and seventh hinges from the bottom.
IF 27-25
757 General Familiarization (7-2005)
ATA 27- Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Trim Control Switch and Indicator
The rudder trim switch is located on aft electronic control panel P8 and activates arming and control switches
which direct electrical input to the rudder trim actuator motor. The switch is spring-loaded to return to neutral.
The RUD TRIM circuit breaker is located on overhead circuit breaker panel P11 and receives 28-volts DC from the
dc standby bus.
The trim position indicator is driven electrically by a transmitter in the rudder trim actuator on the feel, centering
and trim mechanism. The indicator shows up to 17 units of left or right trim. The indicator is located on P8 next to
the rudder trim knob and is powered by 28-volts AC from left bus. The RUDDER TRIM POS circuit breaker is
located on P11.
IF 27-26
757 General Familiarization (7-2005)
ATA 27 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RUDDER POSITION INDICATING SYSTEM
Operation
The rudder position indicating system shows the amount of deflection of the rudder surface. This system does not
include the rudder trim indication system.
The left power bus supplies 28-volts AC to the rotor in the transmitter, which is inductively coupled to the stator.
The rotor is mechanically connected to the airplane control surface and rotates with the surface at the same time.
Movement of the rudder changes the rotor position which in turn changes the stator output. This output signal is
interpreted by the EICAS computers and displayed on the flight deck. The control surface position is displayed on
the left corner of the lower EICAS screen. Press the EICAS status button on panel P9 for display.
757 General Familiarization (7-2005)
ATA 27- Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RUDDER POSITION INDICATING SYSTEM COMPONENTS
EICAS Position Indicator
The EICAS rudder position indicator appears on the lower EICAS display on the pilot's center instrument panel P2.
The indicator is driven by the left or right EICAS computer
puter on equipment shelf E4. The indicator scale in each
direction is equal to 34 degrees.
IF 27-27
757 General Familiarization (7-2005)
ATA 27 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Position Transmitter
The rudder position transmitter is mounted below the lower
wer power control actuator. The transmitter's crank is
attached to the rudder front spar by an adjustable control rod. The position transmitter is clamped to a bracket
mounted to a vertical fin rib. Access to transmitter is through panel 324BL. Electrical power to the transmitter is
controlled by RUDDER POS circuit breaker on overhead circuit breaker panel P11.
IF 27-28
757 General Familiarization (7-2005)
ATA 27- Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELEVATOR CONTROL SYSTEM
Overview
The elevator control system controls airplane pitch attitude by movement of two elevator surfaces which are hinged
on the rear spar of the horizontal stabilizer. The elevator is used for primary pitch control. Pitch trim is provided
by the horizontal stabilizer.
Primary control of the elevators is through pilot or autopilot inputs. Additional elevator positioning is provided by
stabilizer motion which changes the elevator neutral position. The elevator control system is fully powered with no
manual reversion capability.
Each elevator surface is positioned by three hydro-mechanical power control actuators (PCAs). Each actuator is
powered by an independent hydraulic system. The actuators provide surface restraint for flutter suppression.
Control system feel forces are provided by a dual hydro-mechanical feel system.
Elevator positioning is achieved by manual inputs from the captain's or first officer's control column. The two
columns are mechanically slaved together through a spring-loaded override mechanism. Dual cable systems with
tension regulators installed in the forward quadrants are used to transmit pilot inputs to the dual aft quadrants.
The aft quadrants are also mechanically slaved together through a spring-loaded override mechanism. The override
mechanisms allow separation of the column inputs in the event of a jam in one system. A force of approximately
65 pounds is required to separate the systems.
An elevator asymmetry limiter actuator limits the asymmetric motion between the captain's and first officer's aft
quadrants. The amount of asymmetric motion allowed is less at high than at low speeds. This assures that the
pilots cannot command enough asymmetric motion to cause excessive loads.
An elevator feel and centering unit is connected to both aft quadrants, to provide the desired control column feel
and centering characteristics. The feel force is generated by a combination of mechanical and hydraulic devices.
The feel and centering unit is positioned by a neutral shift and override mechanism which allows the stabilizer
trim to introduce a shift of the neutral position in the airplane nose down direction. An override in the neutral shift
mechanism allows the control system to override a jam in the feel and centering mechanism.
Three autopilot servos are connected in parallel to the aft quadrant. Autopilot operation causes the entire control
system to move, including both control columns.
A stick shaker is provided on each control column to warn the pilots of an approaching airplane stall condition.
The stick shakers, which are activated by signals from the stall warning system, consist of an electric motor
driving an unbalanced mass.
Elevator positions are displayed on the lower engine indicating and crew alerting system (EICAS) screen when the
STATUS button is pressed. The EICAS provides airplane status, maintenance, caution, and warning messages.
These messages appear on the EICAS screens.
757 General Familiarization (7-2005)
ATA 27- Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-29
757 General Familiarization (7-2005)
ATA 27 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELEVATOR CONTROL SYSTEM COMPONENTS
Control Column
Two control columns are rigidly mounted on concentric torque
rque tubes beneath the cabin floor. The torque tubes are
slaved together through a cam override. Crank assemblies on each end of the torque tube transmit motion aft by
control rods to forward quadrants/tension regulators. The control columns transmit pilot commands to control the
airplane about the pitch axis. Forward and aft motion of the control columns actuates the elevators. Movement of
the elevators controls the airplane about the pitch axis. Access to the torque tube and control columns is through
the forward compartment access door.
Two forward quadrants are located beneath the cabin floor just aft of the control column torque tube. A control rod
connects each quadrant to a crank assembly installed to the outboard end of each torque tube. The forward
quadrants are connected to the aft quadrants by separate pairs of cables.
A tension regulator in a forward quadrant maintains correct cable tension during temperature variation and
structural deflections. A scale on the bottom of the tension regulator is used to adjust for cable tension.
NOTE:
SOME AIRPLANES; a tension regulator is not installed on the first officer's forward
quadrant.
IF 27-30
757 General Familiarization (7-2005)
ATA 27- Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-31
757 General Familiarization (7-2005)
ATA 27 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Elevator Control Aft Quadrants
Dual aft quadrants are located just aft of the horizontal stabilizer. The aft quadrants receive control cable inputs
from the forward quadrants. The aft quadrants transmit feel and centering forces to the cable system and elevator
control inputs to the power control actuators. The aft quadrants are equipped with a cam override to allow
operation of one elevator should the other jam.
Elevator Control Aft Mechanism Torque Box
The aft mechanism torque box is located just behind the aft quadrants. It supports the aft quadrants, feel unit,
autopilot servos, and neutral shift mechanism.
Elevator Feel Unit and Actuator
The elevator feel unit and actuator are installed on the aft mechanism torque box. The feel unit generates column
force by two means:
1. Dual mechanical centering springs.
2. A hydraulic feel actuator. The feel actuator is a dual-tandem, dual-load-path, floating-body actuator. The
force output of the actuator is the force determined by the highest of the two feel pressures. The actuator
imparts a pull force on the linkage of each output lever through the dual load path link and tie straps. The
tie straps are arranged to give a high force gradient around neutral and lower force gradient for large
displacements.
Neutral Shift and Override Mechanism
The neutral shift and override mechanism is installed on the aft mechanism torque box. The neutral shift
mechanism is connected by control rods to the feel unit and stabilizer. The feel unit is reacted to ground on the
horizontal stabilizer through an override in the neutral shift mechanism. This allows the stabilizer trim to introduce
elevator neutral shift in the airplane nose down direction when the stabilizer is trimmed in the 0.0 to 4.0 unit
range. The override in the neutral shift mechanism allows the control system to override the feel unit in the event
of a jam in the feel actuator.
Asymmetry Limiter Actuator
The linear actuator consists of a 28-volts DC reversible motor that drives an acme screw to extend and retract a
ram. The motor assembly has a magnetic coil-operated disc brake that engages when the motor is de-energized.
Extend and retract limit positions of the ram are controlled by electrical switches located under the switch cover.
The asymmetry limiter actuator limits the maximum possible commanded elevator motion by limiting the travel
between aft quadrants. The system is operational full time but is used only following a left or right elevator control
jam. Asymmetry is limited to 24 degrees of elevator at low speeds and 8 degrees at high speeds. The actuator is
installed on the first officer's aft quadrant.
757 General Familiarization (7-2005)
ATA 27- Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-32
757 General Familiarization (7-2005)
ATA 27 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Elevator Power Control Actuator
Each elevator is positioned by three hydro-mechanical servo actuators. The PCAs are mounted on the horizontal
stabilizer rear spar. The PCAs are each powered by a separate hydraulic system. Each PCA consists of a balanced
piston, a slide/sleeve four-way servo valve, and mechanical summing linkage. Each PCA includes an over-travel
mechanism downstream of the mechanical summing lever. This ensures that full control system inputs can be
applied in either direction with the piston in the full opposite direction without transmitting excessive loads to the
control linkage. The PCAs receive control inputs from the aft quadrants.
Elevator
Two elevator control surfaces are hinged on the rear spar of the stabilizer. Each elevator is driven by three power
control actuators. The elevators are used primarily for maneuvering with pitch trim provided by the horizontal
stabilizer.
Elevator Feel Computer
The feel computer is a dual hydro-mechanical unit that generates two independent feel pressures. The pressures
are a function of airspeed impact pressure and horizontal stabilizer position. The computer consists of two
separate units, one for each hydraulic system (right and center).
Each unit is a force balance valve that develops regulated hydraulic feel pressure to balance the forces developed
from the pressure differential (total pressure minus static pressure) across a diaphragm. The two feel pressures
are monitored by two identical differential pressure sensing mechanisms. When one feel pressure differs
significantly from the other feel pressure, an electrical switch generates a fault message to the pilot on the EICAS
status page.
757 General Familiarization (7-2005)
ATA 27- Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-33
757 General Familiarization (7-2005)
ATA 27 - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Elevator Pressure Reducer and Bypass Valves
One elevator pressure reducing valve is connected to each hydraulic system (L, R, C) and to the flight quadrant.
The valves reduce hydraulic inlet pressure to a preset specified outlet pressure. A pressure bypass valve is
operated by pressure from the right hydraulic system and is connected to the left hydraulic system. If the right
system pressure is lost, the bypass valve will bypass the left pressure reducer and provide full left system
pressure to the left system PCU. The pressure reducers and bypass valves are located on the forward side of the
APU firewall.
IF 27-34
757 General Familiarization (7-2005)
ATA 27- Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STALL WARNING SYSTEM
Overview
The purpose of the stall warning system is to advise the pilots when the airplane nears a stall condition. A stall
condition is indicated by the shaking of both the captain's and first officer's control columns. Two independent,
identical stall warning modules are located in the warning electronics unit (WEU). Each module activates a stick
shaker. The modules use flap and slat information to determine a trip angle of attack (AOA) from a stored table of
values. When the airplane AOA equals or exceeds the computed trip value, the shakers are activated and remain
activated until AOA is reduced below the trip value.
IF 27-35
757 General Familiarization (7-2005)
ATA 27 - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STALL WARNING COMPONENTS
Control Column Shakers
The shakers are DC motor driven vibrating devices. They cause the elevator control column to vibrate when the
shakers are supplied with power. The power is supplied to the shakers by the stall warning modules in the WEU.
IF 27-36
757 General Familiarization (7-2005)
ATA 27- Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stall Warning Module
There are two stall warning modules. The modules are located in thee WEU. The left module activates the captain's
shaker. The right module activates the first officer's shaker. The inputs to both stall warning modules come from
common sources with the exception of air ground logic and flap position. The stall warning modules also provide a
discrete output to activate the auto-slat system.
IF 27-37
757 General Familiarization (7-2005)
ATA 27 - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELEVATOR POSITION INDICATING SYSTEM
Overview
The elevator position indicating system is powered by a 28-volts, 400 Hz AC Bus. The 400 Hz output from the
aircraft power supply is received by the synchro rotor input, and is inductively coupled to the stator of the
transmitter. The rotor is mechanically connected to the aircraft control surface, so that the rotor turns
simultaneously with the control surface. When the rotor changes position, the output from the stator also changes.
The output from the transmitters is sent to the left and right EICAS computers. The surfaces position indicator
appears on the EICAS screen when the STATUS button is pushed, showing elevator position. The indicator will
disappear when the STATUS button is pushed a second time.
757 General Familiarization (7-2005)
ATA 27- Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ELEVATOR POSITION INDICATING COMPONENTS
EICAS Position Indicator
Elevator position indication is provided by a surface position indicator on the EICAS screen. Elevator position
transmitters send left and right elevator position data to the left and right EICAS computers. The EICAS computers
process the signals and display the elevator position on the surfaces position indicator. The indicator appears on
the lower EICAS screen when the STATUS button is pressed.
IF 27-38
757 General Familiarization (7-2005)
ATA 27 - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Elevator Position Transmitter
The elevator position transmitter is a synchro transmitter. The transmitter consists of a stator, rotor, housing
assembly, and a shaft assembly. The elevator position transmitter rotor is connected by a crank arm and pushrod
to the elevator. Each transmitter is connected to the stabilizer rear spar near the fuselage. The transmitters send
position data to the left and right EICAS computers.
IF 27-39
757 General Familiarization (7-2005)
ATA 27- Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HORIZONTAL STABILIZER TRIM CONTROL SYSTEM
Overview
The overall function of the stabilizer trim system is to maintain the airplane in a trimmed condition and to provide
automatic mach and speed stability.
The system consists of :
1. Several alternative sources of trim command signals
2. Two independent actuation devices which position the stabilizer in response to source commands
3. An electronic module which selects a particular trim signal source and also energizes one or the other
actuation device.
The three sources of electric trim command signals are:
1. The flight control computer (FCC)
2. The mach/speed trim function within the stabilizer/elevator asymmetry module (SAM)
3. The column-mounted electric trim switches.
The selection of the active trim source is done by logic circuits in the SAMs.
The SAMs accept input signals from the FCC, the mach/speed trim function in the SAM, and the control wheel
switches. The output signals from the SAM are sent to two STCMs.
There are two independent channels of trim command and actuation. Each channel consists of a SAM, an STCM,
and a hydraulic motor and brake on the stabilizer ball screw actuator. The two channels are known as trim channel
left and trim channel right. When an FCC or mach/speed function is the active trim source, only one trim channel
is activated, providing half-rate trim. When the electric trim switches are energized, both trim channels are
activated simultaneously, and the stabilizer is driven at full-rate trim. The trim rate varies with airplane speed and
horizontal stabilizer position as determined by hydraulic pressure output from the elevator feel computer. The
variable hydraulic pressure output from the elevator feel computer is fed to each STCM. In each STCM a rate
control valve regulates actual hydraulic pressure available to drive the horizontal stabilizer.
In addition to the three electric trim signal sources, stabilizer position can be controlled by manual trim levers.
The levers operate hydraulic valves inside each STCM by mechanical cables. They can override all three electric
trim signal sources.
Travel limiting devices control the range of stabilizer movement in all operating modes. In the electrical control
modes, limit switches within the stabilizer trim limit switch and position transmitter modules (LSTMs) prevent
stabilizer travel beyond that required by the normal flight envelope. In the mechanical control mode, a cable
system repositions valves in the STCMs to shutoff power to the drive hydraulic motors before stops are reached.
757 General Familiarization (7-2005)
ATA 27 - Page 67
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overview (Continued):
Stabilizer position data is generated by the three stabilizer trim limit switch and position transmitter modules
(LSTMs). The modules are driven by mechanical cables attached to the stabilizer. The modules transmit position
data to the SAMs, the FCC's, and a pair of position indicators on the flight deck.
System redundancy and protective devices prevent stabilizer un-commanded motion. Each of the trim channels is
dualized, requiring separate electrical and hydraulic ARM and CONTROL signals to port fluid to the hydraulic
motors. All electrical trim commands can be overridden by use of the manual trim cable system. Hydraulic power
to each control module can be removed by a shutoff switch located on the quadrant stand. The switch controls a
motor-operated shutoff valve in the STCM. Cutoff switches operated by control column motion in opposition to
stabilizer motion will interrupt electrical trim commands. Failure detection/annunciation is provided for stabilizer
trim motion without a valid trim command and for failures resulting in half rate trim operation.
757 General Familiarization (7-2005)
ATA 27- Page 68
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-40
757 General Familiarization (7-2005)
ATA 27 - Page 69
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HORIZONTAL STABILIZER TRIM CONTROL SYSTEM COMPONENTS
Stabilizer Trim Control Wheel Switches
Two sets of dual switches, located in the captain's and first officer's control wheels, actuate the stabilizer trim
system in the manual-electric control mode. The two switches in each set are mounted side-by-side on the control
wheel such that both can be actuated simultaneously with one thumb. Each switch set consists of two single pole,
three position, center off, switches.
The stabilizer trim control wheel switches enable the captain or first officer to input manual trim commands to
both SAMs. Actuating both switches in airplane nose up direction inputs simultaneous airplane nose up (stab LE
down) arm and control commands. Airplane nose down actuation inputs airplane nose down (stab LE up) arm and
control commands.
IF 27-41
757 General Familiarization (7-2005)
ATA 27- Page 70
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27 - Page 71
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STAB TRIM Control Levers
The control stand is equipped with two levers which control the stabilizer in the manual mechanical mode. The
levers are connected to the STCMs by cables. One lever activates the arm hydraulic valve in both STCMs and the
other lever activates the control hydraulic valve. Movement of the STAB TRIM control levers overrides all electric
trim commands.
The arm lever is equipped with a stab trim standby switch which transmits an electrical signal to the SAM when
the STAB TRIM control levers are engaged. This synchronizes the mach/speed trim function to the stabilizer
position commanded by the manual control system. It also prevents mistaken annunciation of unscheduled
stabilizer movement.
Stabilizer Trim Control Module (STCM)
The stabilizer trim system uses two identical hydraulic control modules. They provide direction and rate control to
the stabilizer trim ball screw actuator. Each module consists of an arming valve, control valve, two manually
controlled valves, four solenoid-operated valves, rate controller, motor-operated shutoff valve, manual brake
bypass valve, and a secondary brake pressure switch used in fault monitoring within the SAM. The simultaneous
operation of both STCMs will produce twice the trim rate of just one STCM. The modules are installed just forward
of the stabilizer in Section 48.
A manual brake bypass valve is located in each module and will release the hydraulic pressure on the secondary
brake when depressed. This action applies the brake. The valve is designed to return to its neutral position when
the external force is removed.
757 General Familiarization (7-2005)
ATA 27- Page 72
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-42
757 General Familiarization (7-2005)
ATA 27 - Page 73
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Trim Ball screw Actuator
The ball screw actuator consists of two sets of drive hydraulic motor/secondary brake combinations which singly
or simultaneously drive the ball screw actuator. The ball screw actuator consists of a differential gearbox,
jackscrew and ball nut assembly, and a primary brake.
Each STCM controls a drive hydraulic motor and a secondary brake. The STCM ports fluid to the brake and its
associated drive motor. The brake is mounted in-line with the motor and acts upon the motor output shaft. When
hydraulic pressure reaches the brake-release pressure level, the spring-loaded brake disc disengages and permits
motor operation. The brake is sized to hold the maximum output torque of the drive motor.
The differential gearbox drives a jackscrew and ball nut assembly. The two motors sum their outputs through the
differential. This permits full torque capability with one or both motors driving. When only one side of the
differential is driven (one motor operation) the opposite side is held stationary. This results in one-half the trim
rate. Full trim rate is accomplished with both motors operating. The gearbox utilizes dual load path design to
assure that the braking function of either the hydraulic motor or brake is retained after any single failure.
The jackscrew is dual load path by virtue of a safety rod inside the jackscrew housing which will hold it intact in
the event of screw fracture. The inner rod is torsionally constrained to the primary brake at the base and to the
jackscrew at the top. The ball nut consists of four ball circuits all of which are loaded. If one fails, only 25 percent
of the load carrying capability of the ball nut will be lost. In addition to the ball nut, there is an acme nut which
will ride in the ball races on the jackscrew if all ball nut circuits fail.
The primary brake consists of a disc of friction material and a ratchet plate on both sides of the jackscrew flange.
The ratchet plates and their associated pawls permit ratcheting when the jackscrew is driven in a direction to bring
the airplane into trim, and resist stabilizer loads back driving actuator in the APL out of trim direction.
The ball screw actuator is located in the empennage. The actuator is held in place by two gimbal pin attachments.
The upper gimbal pin attaches the ball nut to the stabilizer. The lower gimbal pin connects the actuator to
bulkhead structure below the stabilizer.
In the alternate-electric-stabilizer trim mode, there are mechanical stops on the upper and lower gimbals to limit
ball screw movement. The stabilizer leading edge up limit is approximately + 3.85 degrees or 24.86 - 25.46 inches
of the B-Dimension on ball screw actuator or 0.20 +/- 0.25 units of trim on position indicator. The stabilizer
leading edge down limit is approximately -11.36 degrees or 0.80 - 1.40 inches of the B-Dimension or 15.30 +/0.25 units of trim.
For all stabilizer trim modes, the stabilizer trim limits are same as the alternate trim mode, except when you trim
the stabilizer leading edge up with the flaps are not in the fully retracted position. In this case, the B-Dimension is
20.43 - 21.03 inches or approximately 3 units of trim.
757 General Familiarization (7-2005)
ATA 27- Page 74
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-43
757 General Familiarization (7-2005)
ATA 27 - Page 75
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Trim Drive Hydraulic Motor
Two identical hydraulic motors power the stabilizer trim drive. The motors are constant-displacement, bent-axis,
piston-type units. The direction of motor rotation is controlled by fluid through the unit. By reversing fluid flow, the
direction of motor rotation can be reversed. The motors may be operated continuously, intermittently, or stalled
without damage at rated pressures and in a system having proper overload relief. There are no controls or
instruments on the hydraulic motor. Operation is controlled entirely by the STCMs.
IF 27-44
757 General Familiarization (7-2005)
ATA 27- Page 76
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Trim Secondary Brake
The secondary brake is a completely enclosed sealed unit. The brake is normally engaged, under which condition
the brake springs are compressing the stator and rotor discs together. On introducing hydraulic pressure into the
brake pressure port, the piston is forced back, compressing the pressure springs and pulling the pressure plate
away from the disc pack. The brake release springs cause a separation of each layer of rotor and stator discs as
the pressure plate is retracted by the piston. On removal of the axial force of the brake pressure springs, the rotor
discs and shaft are free to rotate.
IF 27-45
757 General Familiarization (7-2005)
ATA 27 - Page 77
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Trim Limit Switch and Position Transmitter Modules (LSTMs)
The stabilizer trim control system uses three LSTMs. The LSTMs are connected by cables to the leading edge of
the stabilizer torque box. Each LSTM contains two position transmitters (one synchro and one rotary variable
differential transducer (RVDT)), three travel limit switches, and a green band switch. The synchro
hro transmitter
sends position data to the control stand stabilizer position indicators. The RVDT sends position data to the
stabilizer position module (SPM). The travel limit switches in the left and right LSTMs limit the electrical trim
range in every mode except for manual lever operation. The center LSTM green band switch provides signals to
the warning electronics unit (WEU) for takeoff warning.
IF 27-46
757 General Familiarization (7-2005)
ATA 27- Page 78
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Trim Cut-off Switches
The column operated stabilizer trim cutoff system interrupts electrical trim commands when the control columns
are moved in opposition to stabilizer motion. There are four pairs of switches located on the structure beneath the
control columns and actuated by cams attached to the control columns. Four switches are operated by control
column movement in the elevator-up direction, and four by movement in the elevator-down direction.
The cutoff switches interrupt stabilizer trim motion in opposite direction to elevator commands if the control
columns are displaced a few degrees forward or a few degrees aft of neutral in a direction opposed to the
stabilizer trim. Four switches control the left SAM operation and the other four control the right SAM. The switches
are wired such that if the columns are operated asymmetrically, as would occur after an abnormal system
condition, both trim modules will remain operative in both directions. This allows full rate trim capability after a
jam. The STAB TRIM control levers have control authority over the stabilizer trim cutoff switches.
IF 27-47
757 General Familiarization (7-2005)
ATA 27 - Page 79
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Trim Shutoff Switches
Two guarded stabilizer trim shutoff switches allow system hydraulic pressure to be removed from the STCMs. The
switches control motor-operated shutoff valves in the STCMs. The right system shutoff switch controls the right
STCM. The center system shutoff switch controls the left STCM. The switches are mounted on the control stand.
IF 27-48
757 General Familiarization (7-2005)
ATA 27- Page 80
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer/Elevator Asymmetry Limiter Module (SAM)
The SAM is part of the integrated package of flight control electronics, the control system electronics unit (CSEU).
Each of the two (left and right) CSEUs contains a SAM. The SAMs provide electric stabilizer trim control and
associated failure monitoring.
Each of the two SAMs provides:
1.
2.
3.
4.
5.
6.
7.
8.
Stabilizer trim mode priority logic for manual electric trim, autopilot trim, and mach/speed trim
Stabilizer trim engage/disengage and auto transfer logic
Unscheduled trim fault detection
Manual-electric trim monitoring
Fault messages for the caution and warning system as well as for maintenance
Mach/speed trim function
Rudder ratio changer command computation
Elevator asymmetry limiter control
Fault balls on the SAMs display line replaceable unit (LRU) status. A manual RESET button on the SAM clears
intermittent failures.
IF 27-49
757 General Familiarization (7-2005)
ATA 27 - Page 81
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STABILIZER TRIM POSITION INDICATING SYSTEM
Operation
Stabilizer trim position is provided by two independent indicators. The indicators are located on the left and right
sides of the control stand. They are equipped with greenband range denoting safe takeoff stabilizer trim setting.
The stabilizer trim position indicating system provides position information to the flight crew as well as other
systems.
Stabilizer trim position is displayed by indicators on the control stand. When the stabilizer trim setting is outside
of the takeoff greenband, the takeoff aural warning circuit is armed. If the indicator loses power, an "OFF" flag will
appear at the bottom of the indicator. If there is a loss of the synchro signal to the indicator, the indicating tape
will disappear from view.
Stabilizer trim position is sensed by three stabilizer trim limit switch and position transmitter modules
(LSTMs).The LSTMs are driven by cables attached to the stabilizer. The left and right LSTMs transmit position
data to the stabilizer position indicators.
The LSTMs also provide:
1. Electrical travel limits to Horizontal Stabilizer Control System
2. Data for takeoff configuration warning and annunciation
3. Stabilizer position data to Stabilizer Position Modules (SPMs)
The SPMs provide stabilizer position data to:
1. Autopilot Flight Director System (AFDS)
2. Stabilizer Trim/Elevator Asymmetry Limiter Modules (SAMs)
757 General Familiarization (7-2005)
ATA 27- Page 82
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
STABILIZER TRIM POSITION INDICATING SYSTEM COMPONENTS
Stabilizer Position Indicators
Stabilizer trim position is displayed by two indicators on the control stand. The indicators show stabilizer position
from 0 to 15.5 units of trim. The indicators receive their inputs from synchro transmitters
itters in the LSTMs. A
greenband on the face of each indicator shows the range of stabilizer settings for safe takeoff. The greenband
range is from 2.0 to 7.0 units of trim. If a power failure occurs, the indicators will show an OFF flag at the bottom
of the indicator. If a signal loss occurs, the indicator's tape will disappear out of view.
IF 27-50
757 General Familiarization (7-2005)
ATA 27 - Page 83
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
B. Stabilizer Trim Limit Switch and Position Transmitter Module (LSTM)
The stabilizer control system uses three interchangeable LSTMs, each connected via a cable system to the leading
edge of the stabilizer torque box. The LSTMs are mounted just forward of and below the stabilizer in Section 48.
Each LSTM contains two position transmitters (one synchro and one rotary variable differential transducer
(RVDT)), three travel limit switches, and a greenband switch.
The two stabilizer position indicators are electrically driven by the synchro transmitters in the left and right
LSTMs. The synchro in the center LSTM is used for flight recorder information.
The RVDTs in the three LSTMs signal stabilizer trim position to corresponding SPMs.
The travel limit switches are only functional on the left and right LSTMs. They interrupt the trim signal from the
SAMs to the Stabilizer Trim Control Modules (STCMs) at the stabilizer up and down travel limits. The up and
down travel limit switches for normal operation are only functional on the left and right LSTMs. The up and down
limit switches for standby trim are only functional on the left and right LSTM. They interrupt the trim signal from
the SAMs to the Stabilizer Trim Control Modules (STCMs) at the stabilizer up and down travel limits.
The greenband switch is only functional on the center LSTM. The switch signals the takeoff configuration card to
issue a takeoff configuration warning and annunciation if the stabilizer trim position is outside of the greenband
range and the airplane is in a takeoff configuration.
757 General Familiarization (7-2005)
ATA 27- Page 84
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-51
757 General Familiarization (7-2005)
ATA 27 - Page 85
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Stabilizer Position Modules (SPMs)
Each of the SPMs receives a stabilizer trim position input signal from the corresponding LSTM.
The left and right SPMs signal stabilizer trim position to the SAMs.
All SPMs signal stabilizer trim position to the AFDS.
The SPMs are located in the electrical system card file P50.
IF 27-52
757 General Familiarization (7-2005)
ATA 27- Page 86
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TRAILING EDGE FLAP SYSTEM
Overview
The trailing edge flaps provide extra aircraft lift during takeoff and landing. When extended the flaps operate in
conjunction with leading edge slats to effectively increase camber and area of the wings.
Increased wing camber and area decreases stall speeds and allows the aircraft to be maneuvered at lower
airspeeds. Lower aircraft stall speeds allow lower takeoff and landing speeds. This reduces ground roll, increases
creases
wheel and brake life, increases climb-out performance, and creates safer operating conditions.
There are two flaps on each wing, an inboard flap and an outboard flap. Each flap consists of a main flap and an
aft flap. The flaps are faired with the wing for cruise conditions and are extended aft and downward at takeoff and
landing.
During normal operation, the flaps are powered by the left hydraulic system. Normal flap operation is controlled by
the flap control lever mounted on the pilot's control stand.
During alternate operation the flaps are electrically powered. Control of the flaps is provided by a guarded
alternate drive arming switch and a rotary alternate flaps/slats position selector switch located on the center
panel.
A flap position indicating system is provided to show flap position during extension and retraction. A combined
flaps and slats dial indicator is located on the center instrument panel.
IF 27-53
757 General Familiarization (7-2005)
ATA 27 - Page 87
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27- Page 88
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Trailing Edge Flap Primary Control
The flaps are normally controlled with the flap control lever on the center aisle stand. Movement of the flap control
lever rotates a cable drum attached to the flap control cables. The flap control cables are connected to the input
drum on the flap power drive unit (PDU) and cause rotation of the input drum.
The input drum is connected to a cam which rotates with the drum. This cam is linked to the control valve and
positions the control valve spool. The control valve will pass hydraulic fluid to the retract or extend port of the PDU
hydraulic motor depending upon flap control lever input.
As the flaps move towards the commanded setting a feedback mechanism repositions the control valve spool to a
neutral position, stopping the hydraulic motor.
Trailing Edge Flap Alternate Control
Back up operation of the flap system is done electrically. The flap system is controlled electrically by a rotary
position switch and the trailing edge arming switch.
The arming switch moves both flap and slat bypass valves to the bypass position and closes the flap/slat
depressurization valve to remove hydraulic pressure from the high lift hydraulic system.
The rotary selector switch commands a closed loop electrical system that selects the same seven flap positions as
the hydraulic control system. Rotary selector switch position is monitored by FSEU-2 and FSEU-3. FSEU-3
generates signals to close extend or retract relays which power the PDU electric motor and energize the electric
motor clutch.
A position transmitter on the PDU sends the flap position data to FSEU-3. The FSEU uses this data to control
alternate motor operation. When the flaps reach the commanded position the drive relays are opened, the clutch is
released, and the alternate drive motor stops.
Primary Power
The left hydraulic system provides primary power to the flap drive system. In the event of a left engine failure, the
flaps and slats, as well as the landing gear, are powered by the power transfer unit (PTU). The PTU consists of a
hydraulic motor which drives a hydraulic pump. The PTU motor is powered by the right hydraulic system and the
PTU pump provides left hydraulic system pressure.
Alternate Power
The flaps are driven by an electric motor in the alternate control mode. The electric motor runs on 3-phase, 400
Hz, 115-volts AC current provided by the right AC bus.
757 General Familiarization (7-2005)
ATA 27 - Page 89
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap Drive System
The flaps are positioned by two ball screw actuators attached to each flap. The ball screws are driven through a
system of transmissions and gearboxes which attach to the torque tube system. The torque tubes extend out into
each wing and are powered by the flap power drive unit (PDU).
Each flap rides on two flap tracks which determine the angle that the flap makes with the wing and which transfer
the lifting forces of the flaps to the wing structure.
The aft flaps ride on tracks within the flap structure and are driven to the extended position by bell cranks and
connecting rods.
The outboard flap track for the inboard flap is attached to the forward support fitting of the wing by a fuse pin.
Both the inboard and outboard flap tracks for the outboard flap are attached to the wing by fuse bolts. These fuse
pins and bolts shear under extreme load to prevent major damage to the wing structure.
Flap Load Relief
The flap load relief system provides automatic retraction of the flaps from the 30-unit position to the 25-unit
position to protect the flaps from excessive air loads. The flap load relief system operates only when the flap lever
is positioned in the 30-unit detent. FSEU-1 receives airspeed data from the digital air data computer. If this speed
exceeds 170 knots, the FSEU sends a signal to the flap load relief solenoid on the flap control valve. The solenoid
moves the control valve sleeve allowing hydraulic fluid to pass to the hydraulic motor and drive the flaps to the 25unit position. The flaps will return to the 30-unit position when airspeed is reduced to 165 knots. The flap control
lever will not change position during load relief.
The flap load relief system is inoperative when the alternate drive system is armed.
Depressurization
The flap/slat depressurization module shuts off hydraulic power to the flaps and slats when the flaps have been
retracted for 25 seconds or when the alternate drive systems are armed. Asymmetry or un-commanded motion
causes hydraulic shutdown of the flap/slat systems.
Asymmetry
A difference of 4% in right flap travel compared to left flap travel is considered an asymmetry condition.
FSEU-1 detects flap asymmetry by comparing resolver output from each wing.
In normal control mode an asymmetry condition causes the FSEU to command hydraulic shutdown of the flap/slat
systems. The amber TRAILING EDGE light comes on and the EICAS display screen will show a TE FLAP ASYM
message.
In alternate mode FSEU-2 detects an asymmetry condition resulting in the TE FLAP ASYM message on the EICAS
display and illumination of the amber TRAILING EDGE lights.
757 General Familiarization (7-2005)
ATA 27- Page 90
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pg. 1 of 2
IF 27-54
757 General Familiarization (7-2005)
ATA 27 - Page 91
For Training Purposes Only
Pg 2 of 2
IF 27-54A
757 General Familiarization (7-2005)
ATA 27- Page 92
AeroEd LLC – Aviation Education Resource
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-55
757 General Familiarization (7-2005)
ATA 27 - Page 93
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TRAILING EDGE FLAP SYSTEM COMPONENTS
Flap Control Lever
The flap control lever, located on the center aisle stand, controls operation of the flaps and slats during normal
operation.
The control stand has detents at the 0-, 1-, 15-, 20-, 25-, and 30-unit positions. The flap lever is spring-loaded to
lock into each detent. Lifting the lever releases the handle to allow rotation.
If the flap lever is not in any detent or not properly seated in a detent for 25 seconds, a condition exists that
results in FSEU commands to depressurize the flap/slat hydraulic systems and to move both flap and slat bypass
valves to the bypass condition. The amber TRAILING EDGE light comes on and the EICAS display screen will show
a TE FLAP DISAGREE message.
Gates are provided at detents 1 and 20 to prevent inadvertent movement past these positions.
Flap control lever position is electrically monitored by a resolver attached to the flap lever by linkages.
Two control cables, WFA and WFB, are attached to the cable drum on the flap control lever. Control movements of
the flap control lever are transferred to the flap drive system by these cables. Cable WFA is the flap retract cable
and cable WFB is the flap extend cable. The control cables run from the cable drum under the cabin floor and back
to the PDU at the forward end of the left main wheel well. The control cables are attached to the flap PDU input
drum.
757 General Familiarization (7-2005)
ATA 27- Page 94
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Alternate Control Panel
The flap/slat systems are controlled electrically by a rotary
tary position switch and the leading and trailing edge
arming switches. The arming switches operate the hydraulic motor by-pass valves and depressurize the high lift
system hydraulics.
The rotary switch energizes the electric motor clutches and commands a closed loop electrical system that selects
the same seven flap positions as the hydraulic control system (except that slats are fully extended for the 20-unit
position selection). Slats are commanded to the fully extended position for the 20-, 25-, and 30-unit positions, and
to the intermediate position for the 1-, 5-, and 15-unit positions. Flaps extend to the same positions as under
normal hydraulic control.
IF 27-56
757 General Familiarization (7-2005)
ATA 27 - Page 95
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap Power Drive Unit (PDU)
The flap power drive unit is the source of mechanical power for the flap drive system.
The flap PDU is a gearbox housing. Mounted on the housing are a hydraulic motor, alternate electric motor,
hydraulic bypass valve, control valve assembly, mechanical controls and a rotary position transmitter. The
hydraulic and electric motors drive the common gear train which drives the torque tubes.
The hydraulic motor powers the flap drive system during normal operation. The motor is designed for reversible
operation at a supply system pressure of 3000 PSI.
The alternate electric motor is mounted on the PDU housing and provides power to the flaps if hydraulic power is
not available. The electric motor is a 400-Hz, 3-phase, 115-volts AC motor.
The motor is controlled by FSEU-3 and relays for extension or retraction. The rotary position selector switch
commands motor actuation through the FSEU.
Maximum duty cycle for the electric motor is 4 minutes (3 minutes off) to avoid overheating.
The hydraulic control valve directs hydraulic fluid to the retract or extend ports of the hydraulic motor to power the
flap drive system.
The solenoid on the control valve is powered by FSEU-1 when a flap load relief condition exists, the solenoid
moves directing the hydraulic motor to extend or retract the flaps. A cam follower is linked to the input crank on
the control valve. As the flaps approach their commanded position the follow-up cam moves the linkages to return
the control valve to its null position, shutting off hydraulic power to the motor.
The bypass valve removes hydraulic power from the PDU during alternate flap operation or during protective flap
drive system shutdown.
When in the bypass position the bypass valve connects the extend and retract ports of the hydraulic motor,
preventing the hydraulic motor from jamming the flap drive system.
The resolver on the PDU transmits information on the angular position of the PDU gear train to FSEU-3. The FSEU
uses this position information to control the alternate electric motor.
The PDU must be rigged to ensure that the PDU will move the flaps to the commanded position.
757 General Familiarization (7-2005)
ATA 27- Page 96
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-57
757 General Familiarization (7-2005)
ATA 27 - Page 97
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap/Slat Depressurization Module
The flap/slat depressurization module is installed in the left main gear wheel well directly forward of the flap
PDU. The hydraulic flow to the flap and slat hydraulic motors passes through the depressurization module.
The primary function of the depressurization module is to remove hydraulic power from the flap and slat drive
systems during cruise or in the event of drive system failures. Depressurization at cruise reduces valve erosion
and minimizes the possibility of un-commanded flap/slat motion. In the event of a flap asymmetry condition or
un-commanded motion problem, the flap/slat systems are shut down hydraulically.
This module also gives flap and slat operation priority over landing gear operation. If pressure in the left hydraulic
system drops below 1800 PSI, the priority valve closes, leaving a 3-gallon per minute orifice for the landing gear.
Within the module the flap/slat sequencing valve serves to reduce total flow to the flaps and slats when both
systems are operating at the same time. The valve normally serves as a flow limiter to the flap system. When slat
motion is commanded, flow to the flap system is reduced to allow adequate flow to the slat system. Total flow to
the flaps and slats is limited to assure adequate flow to other hydraulic systems.
IF 27-58
757 General Familiarization (7-2005)
ATA 27- Page 98
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap Torque Tubes
Output from the PDU is transmitted by a single torque tube system. The torque tubes are aluminum tubes
connected with steel sleeve and spline couplings. Safety straps
raps are installed to prevent damage to surrounding
airplane systems in the event of a torque tube failure.
Torque tubes attach to each other through intermediate support bearings. Torque tubes attach to gearboxes to
change direction and to transmit power to flap actuation transmissions.
IF 27-59
757 General Familiarization (7-2005)
ATA 27 - Page 99
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap Transmissions
Each trailing edge flap is driven by two transmission assemblies. Each transmission contains a 15-to-1 gear
reduction, a no-back brake assembly and a torque limiter. The torque limiter is a ball ramp, disk brake device that
protects the system gearing from damage due to jamming.
The no-back brake assembly consists of a brake disk that is pushed against a brake plate by flap air loads
transferred through the drive system. The flaps are kept from retracting by ratchet pawls which keep the brake
plate from turning in the retract direction. During flap extension the pawls ratchet freely and the brake plate
rotates with the brake disk. During flap retraction the brake plate is locked in place and the brake disk rubs
against it. This causes a large friction torque which the drive system must overcome.
Output is from the flap transmissions drive ball screw actuators, which convert rotary motion to linear motion. The
ball screws are attached to the transmissions through universal joints which allow the ball screws to change angle
as the flaps are extended. Ball nuts on the ball screws are attached through gimbals to flap carriages to drive the
flaps.
IF 27-60
757 General Familiarization (7-2005)
ATA 27- Page 100
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap Carriages
Each main flap is attached to two carriage roller assemblies which ride on flap tracks. The carriages are steel
forgings to which roller bearings are attached. The rollers support the flap on the flap track and transfer the lift
generated by the flaps to the wing structure. The carriage is moved along the flap track by the ball nut. Flap angle
is determined by the shape of the flap tracks and the position of the carriage along the track.
IF 27-61
757 General Familiarization (7-2005)
ATA 27 - Page 101
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fairings
The flap tracks and the flap drive mechanisms are mounted below the wing surface so they are housed in
aerodynamic fairings.
IF 27-62
757 General Familiarization (7-2005)
ATA 27- Page 102
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27 - Page 103
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap/Slat Electronic Unit (FSEU)
Monitoring and control of the flap and slat drive systems is done electrically by the flap/slat electronics unit.
The FSEU consists of 3 interchangeable modules; FSEU-1, FSEU-2 and FSEU-3. The 3 modules are physically and
functionally identical but they perform different functions based upon where they are installed in the equipment
rack. Each FSEU module has a microprocessor programmed to perform all FSEU functions. When the module is
installed in one of three locations in the equipment rack the wiring from the electrical connectors enables only a
specific channel, defining the module's function.
FSEU-1 performs all primary flap and slat drive system protection, indication and automatic operation.
Specifically, FSEU-1 provides:
1.
2.
3.
4.
5.
6.
7.
Un-commanded motion detection and protection of the flaps during hydraulic operation
Asymmetry detection and protection during hydraulic operation
Flap load relief functions
Auto-slat controls
Flaps up depressurization
Slat position indication signals
Signals to other aircraft systems
FSEU-2 performs alternate flap and slat drive protection and annunciation.
Specific functions are:
1.
2.
3.
4.
5.
Un-commanded motion detection under alternate control
Asymmetry detection under alternate control
Auto-slat control
Position indication
Signals to other aircraft systems
FSEU-3 performs alternate flap and slat drive control.
Specific functions are:
1. Alternate electric drive motor control
2. Signals to other aircraft systems
Built in Test Equipment (BITE) electrically tests the FSEU and interfacing electronic components. BITE software is
designed to isolate faults in the electronic system down to the line replaceable unit level.
The BITE test is activated by pressing a switch on the front of each FSEU box. The BITE test is self terminating
and the results of the test are shown on the FSEU front panel.
757 General Familiarization (7-2005)
ATA 27- Page 104
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-63
757 General Familiarization (7-2005)
ATA 27 - Page 105
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TRAILING EDGE FLAP POSITION INDICATING SYSTEM
Overview
The trailing edge flap position indicating system provides flap position and flap system failure information to the
flight deck.
Flap surface and flap lever position sensing is performed with resolvers for both alternate and normal flap control
separate. Sensors are used for each mode.
Position data is used by the Flap Slat Electronics Unit (FSEU) for asymmetry detection, un-commanded motion
detection, fault annunciation, position indication, control protection, and flap/slat position information to other
systems.
Position indication is identical for both normal and alternate control modes. Flap position is indicated by a
synchro-driven dial indicator.
A TRAILING EDGE amber light will illuminate to indicate flap system faults. A concurrent Engine Indicating and
Crew Alerting System (EICAS) message will show the nature of the fault.
757 General Familiarization (7-2005)
ATA 27- Page 106
For Training Purposes Only
ATA 27 - Page 107
AeroEd LLC – Aviation Education Resource
IF 27-64
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TRAILING EDGE FLAP POSITION INDICATING SYSTEM COMPONENTS
Flap/Slat Position Indicator
The primary source of flap position information is the flap/slat position indicator located on the First Officer's
main instrument panel (P3).
The position indicator uses two needles (one for each wing) to indicate flap position between the UP-unit and 30unit positions.
The position indicator needles are driven directly from the flap synchros except between flap positions UP and 1.
During extension, if either the flaps or the slats have left the retracted position, and the flaps are not past the 1unit position, both needles will simultaneously move to a position half way between UP and 1. The needles will
jump to 1 when the slats reach the sealed position and the flaps reach the 1-unit position.
When the flaps move beyond the 1-unit position, flap position will be provided directly to the needles by the
synchros at the ends of the flap drive torque tubes. During retraction, both needles will move from the 1-unit
indication to the 1/2-unit indication if the flaps or slats retract from the 1-unit position.
When both flaps and slats are retracted, both needles will jump from the 1/2-unit indication to the UP indication.
The amber TRAILING EDGE light illuminates whenever an un-commanded motion or asymmetry condition is
detected. If the condition is corrected the light will stay on until the alternate drive system is armed, then unarmed
for un-commanded motion, or the power to FSEU-1 is interrupted for an asymmetry condition.
IF 27-65
757 General Familiarization (7-2005)
ATA 27- Page 108
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap/Slat Electronic Unit (FSEU)
The FSEU processes information from the high lift system and provides indications and data for display and
warning when crew alerting or warning is required.
The FSEU monitors the output of the flap system resolvers to obtain flap system position information. Position
data is used by the FSEU for asymmetry detection, un-commanded motion detection, fault annunciation, position
indication, control in the alternate mode, protection in the normal mode, and flap/slat position information to other
systems.
The FSEU has a built-in test equipment system which performs a self test of the FSEU and interfacing hardware.
IF 27-66
757 General Familiarization (7-2005)
ATA 27 - Page 109
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Indicating and Crew Alerting System (EICAS)
Failures of the flap system are displayed on the EICAS screen. Message levels are according to following table:
EICAS supplies following specific TE flap messages:
1.
2.
3.
4.
Level A: FLAPS
Level B: TE FLAP DISAGREE and TE FLAP ASYM
Level C: FLAP LD RELIEF
Level S:
a)
b)
FLAP/SLAT ELEC
FLAP ISLN VAL
5. Level M: FLAP/SLAT ELEC, FLAP ISLN VAL, and FLAP/SLAT BITE
757 General Familiarization (7-2005)
ATA 27- Page 110
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Resolvers
Position information used by the FSEU is provided by resolvers. The resolvers compare the phase shift of a rotary
transformer with respect to a reference signal. The resolver output is in degrees of rotation. The analog resolver
output is converted to digital form which the FSEU can use.
There are six resolvers in the flap position indication system:
1.
2.
3.
4.
Two at the end of the right torque tube system: RFR1 and RFR2
Two at the end of the left torque tube system: RFL1 and RFL2
One on the flap power drive unit: RFP3
One in the pilot's control stand linked to the flap control lever: RFH1
Synchros
There are two flap position synchros in the flap position indicating system. They are located at the end of the
torque tube in each wing.
The synchro in the left wing provides input to the "L" position needle and the synchro in the right wing provides "R"
needle input.
Flap position indication is provided directly from the flap synchros except between up and 1 positions. Between up
and 1, a fixed output synchro provides input to the position indicator needles.
IF 27-67
757 General Familiarization (7-2005)
ATA 27 - Page 111
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
SPOILERS AND DRAG DEVICES
Overview
The twelve spoiler control surfaces are numbered 1 thru 12, consecutively, for reference. Spoiler surfaces 1 thru 6
are on the left wing, numbered from outboard to inboard. Spoiler surfaces 7 thru 12 are on the right wing,
numbered from inboard to outboard. Spoilers 1, 2, 3, 5, 6, 7, 8, 10, 11 and 12 are used to assist aileron lateral
control. These panels also respond to speed brake lever commands in flight. Spoiler surfaces 4 and 9 are ground
speed brakes only. All twelve surfaces respond to speed brake lever commands through the auto-speed brake
system.
IF 27-68
757 General Familiarization (7-2005)
ATA 27- Page 112
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
SPOILER/SPEEDBRAKE CONTROL SYSTEM
Overview
The spoiler/speed brake system uses the same flight control surfaces to perform two functions. The system
deploys the flight spoilers to assist the ailerons in lateral control of the airplane. The system also operates the
same surfaces as speed brakes to increase drag and reduce lift in flight.
The spoiler/speed brake system is electrically controlled and hydraulically powered. Rotary variable differential
transducers (RVDTs) translate aileron control wheel inputs into electrical signals. The spoiler control modules
(SCMs) receive these signals and command the power control actuators (PCAs) to raise spoilers. Placing the speed
brake lever in the UP position will deploy all flight speed brakes.
IF 27-69
757 General Familiarization (7-2005)
ATA 27 - Page 113
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Operation
Spoiler assist in lateral control begins when the control wheel rotates. Either the flight crew or the autopilot system
can operate the control wheels. The bus force limiter linking the control drums transfers input from one control
wheel to the other. All the RVDTs move anytime either control wheel is rotated.
RVDT signals received by the SCMs are processed and mixed with other inputs. SCM logic then determines which
spoilers, if any, should be deployed. The SCM sends an appropriate command signal to the left or right flight
spoiler PCAs. As the spoiler moves to the commanded position, the PCA RVDTs feed signals back to the SCM. The
feedback signal is compared with input commands. When the two signals agree, spoiler movement is stopped.
Manual operation of the speed brakes in flight begins when the speed brake lever is moved from the DOWN
position through ARMED to the UP position. There is no speed brake movement when the lever is moved from the
down and locked position to the ARMED detent.
Inputs to the SCMs for the speed brake system come from the speed brake LVDT pairs. Speed brake lever
movement is transferred to the LVDTs through a linkage and the speed brake mechanism. One LVDT of a pair
sends a signal to each SCM in the left CSEU and the other LVDT sends a signal to each SCM in the right CSEU.
The SCMs process and mix LVDT signals to produce a command signal. All flight spoilers respond to speed brake
commands during flight.
When there is input from both control drum RVDTs and speed brake LVDTs, the SCMs combine the two signals.
The resulting SCM command is a summation of the two signals.
Spoiler/Speed brake System
Each of the SCMs receives input from three spoiler RVDTs. The three SCMs in the left CSEU receive inputs from
the captain's control wheel RVDTs. The three SCMs in the right CSEU receive inputs from the first officer's control
wheel RVDTs. Each SCM also receives inputs from three speed brake LVDTs and two PCA RVDTs. The resultant
output is sent to the PCAs to position spoiler panels.
The ten flight spoilers respond to control wheel inputs. These flight spoilers extend 45° with full control wheel
movement except for 6 and 7 which extend 15°. As flight speed brakes, spoilers 1, 2, 3, 10, 11, and 12 extend 30°;
spoilers 5 and 8 extend 20°; and spoilers 6 and 7 extend 15°. Spoilers 4 and 9 are ground speed brakes only.
757 General Familiarization (7-2005)
ATA 27- Page 114
For Training Purposes Only
ATA 27 - Page 115
AeroEd LLC – Aviation Education Resource
IF 27-70
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
SPOILER/SPEEDBRAKE SYSTEM COMPONENTS
Rotary Variable Differential Transducers (RVDTs)
Three RVDTs are grouped together in a can-like unit. The unit is mounted on the bottom of each control drum
assembly. The units are accessible through an access panel forward of the nose gear wheel well.
The six spoiler RVDTs translate aileron control wheel rotation into electrical signals.
s. The signal voltage is
proportional to the amount of control wheel movement. The SCMs mix RVDT inputs with other inputs according to
preprogrammed logic.
IF 27-71
757 General Familiarization (7-2005)
ATA 27- Page 116
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Spoiler Control Modules (SCMs)
Six SCMs control the twelve spoiler surfaces. Each SCM operates a symmetrical pair of spoilers. The modules are
part of the flight control system electronics units (CSEU). Three SCMs are part of the left CSEU and three are part
of the right CSEU. The SCMs will show you when failures in the spoiler/speed brake system occur.
The 100 series SCMs have small windows (called fault balls) on the front of them that show yellow when a failure
occurs. The yellow fault ball identifies the failure. You can push the reset button on the SCM to make the fault ball
black.
The 200 series SCMs have a display to show failures. The failures are shown as words on the display. The 200
series SCM is better for trouble-shooting than the older 100 series SCM.
IF 27-72
757 General Familiarization (7-2005)
ATA 27 - Page 117
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Power Control Actuators (PCAs)
The inboard and outboard spoiler PCAs are similar. Each PCA consists of an actuator, an electro hydraulic servo
valve (EHSV), and a rotary variable differential transformer (RVDT). The outboard PCAs are trunnion-mounted on
the wing rear spar. The inboard PCAs are trunnion-mounted on the landing gear support beam. The PCA rod ends
are attached to spoiler panel surfaces. Each spoiler has one PCA, powered by one of the three hydraulic systems.
The PCAs are accessible when the trailing edge flaps are extended.
The PCA extends or retracts as commanded to raise or lower the spoiler. The RVDT sends a feedback signal to the
SCM proportional to the amount of PCA movement. This signal cancels out the input signal
gnal to the EHSV when the
PCA reaches the commanded position.
Each PCA has a manual release cam which is accessible from underneath the PCA. The manual release cam opens
the PCA thermal relief valve to release trapped hydraulic fluid after the hydraulic system is depressurized.
Releasing trapped fluid is necessary to raise the spoiler panel for maintenance. The manual release cam cannot be
used to raise a spoiler panel when power is available to its PCA. Also, a spoiler panel which was raised using the
manual release cam when hydraulic power was removed will immediately retract if hydraulic power is reapplied.
Therefore, the manual release cam should be used with caution, and only when it is certain hydraulic power will
not be applied when the panel is raised.
IF 27-73
757 General Familiarization (7-2005)
ATA 27- Page 118
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Electro hydraulic Servo Valves (EHSVs)
The EHSV controls the flow of hydraulic fluid in the PCA in response to SCM commands. The command operates a
jet pipe that supplies hydraulic fluid to the EHSV control bobbin. The EHSV is spring biased in the retract position.
This causes the spoiler panel to retract if there is no command signal. Pressure is available to the EHSV whenever
the hydraulic system supplying the PCA is pressurized.
IF 27-74
757 General Familiarization (7-2005)
ATA 27 - Page 119
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Spoiler Panels
Each wing has six spoiler panels. The spoilers are numbered 1 through 12 consecutively from the left wing
outboard to the right wing outboard. Spoilers
rs 4 and 9 are ground speed brakes only. These panels are not actuated
during flight under any circumstances.
Each spoiler is attached to wing structure at four hinge points. The PCA rod ends are attached to the center of
each panel.
IF 27-75
757 General Familiarization (7-2005)
ATA 27- Page 120
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Speed brake Lever and LVDTs
The speed brake lever is located on the control stand between the stabilizer trim levers and the engine thrust
levers. An adjustable connecting rod links the lever to the speed brake mechanism.
The six speed brake LVDTs are arranged in pairs. Each pair of LVDTs is attached to the speed brake mechanism at
a common point. Speed brake lever movement will cause the LVDTs to send a signal to the SCMs corresponding to
lever position.
The speed brake lever has three positions. In the DOWN position, the spoiler panels are retracted. Pulling up on
the lever and moving it aft places the lever in the ARMED position. This prepares the speed brake system for
automatic operation. Pulling the lever all the way back to the UP position will raise the speed brakes to full
extension.
IF 27-76
757 General Familiarization (7-2005)
ATA 27 - Page 121
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flight Deck Indications
Flight deck indicators display failures in the spoiler/speed brake system. Thee amber SPOILERS light on the pilot's
overhead panel, P5, indicates a second failure in the spoiler system. An amber SPOILERS message will also
appear on the upper EICAS screen. First spoiler system failures are stored in EICAS memory and can be recalled
on the ground.
The warning system provides an amber light and level B EICAS caution message should the spoiler/speed brake
surfaces be extended at altitudes less than 800 feet. The amber SPEED BRAKES light is located on the pilot's main
panel P1. The amber EICAS caution message reads SPEED BRAKES EXT.
IF 27-77
757 General Familiarization (7-2005)
ATA 27- Page 122
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AUTO-SPEEDBRAKE CONTROL SYSTEM
Overview
The auto-speed-brake system automatically deploys the spoilers at touchdown and after a refused takeoff. The
system also automatically retracts the spoilers when a go-around is initiated after touchdown.
Auto-speed-brake system failures cause the amber AUTO SPDBRK light on the pilot's overhead panel P5 to come
on. An amber AUTO SPEEDBRAKE message will also appear on the upper EICAS display unit.
The warning system provides an amber light and level B EICAS caution message should the spoiler panels be
extended at altitudes less than 800 feet. The amber SPEED BRAKES light is located on the pilot's main panel P1.
The amber EICAS caution message reads SPEED BRAKES EXT.
IF 27-78
757 General Familiarization (7-2005)
ATA 27 - Page 123
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AUTO-SPEED-BRAKE SYSTEM COMPONENTS
Auto-Speed-brake Switches
The auto-speed-brake actuator arming switch is a single roller switch on the speed brake lever mechanism.
Placing the speed brake lever in the ARMED position activates the arming switch. The activated switch provides
power to the auto-speed-brake actuator operating relays.
The speed brake lever position switch is a double roller switch located below the arming switch. The left switch
disarms the auto-brake system when the speed brake lever is moved from the up position towards the down and
locked detent. The right switch signal is sent to the takeoff warning system which then compares speed brake
lever position to throttle position during takeoff. The auto-speed-brake switches are accessible through the control
stand side panels.
IF 27-79
757 General Familiarization (7-2005)
ATA 27- Page 124
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Auto-Speed-brake Actuator and No-Back Clutch
The auto-speed-brake actuator drives the speed-brake lever to the UP position to operate the LVDTs which input to
the spoiler control modules (SCMs). The actuator is attached to the no-back clutch at one end and structure at the
other. The actuator is accessible through control stand side panel.
The no-back clutch has two functions. The clutch enables the flight crew to manually operate the speed-brake lever
without having to back-drive the electric actuator as well as override an unscheduled auto-speed-brake deploy
signal. The clutch also enables the actuator to drive the speed-brake lever to the UP position which in turn will
operate the LVDTs which then provide input to the SCMs to raise or lower the panels.
IF 27-80
757 General Familiarization (7-2005)
ATA 27 - Page 125
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Reverse Thrust Mechanism
The reverse thrust mechanism mechanically moves the speed brake lever out of the down and locked detent
position when one or both thrust levers are moved into reverse idle range. The reverse thrust mechanism includes
a cam, shafts, linkages, and the engine reverse thrust switch. The mechanism is accessible through control stand
panels.
IF 27-81
757 General Familiarization (7-2005)
ATA 27- Page 126
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Forward Thrust Lever Position Switches
The forward thrust lever position switches are part of thee auto throttle switch pack. The switch pack is located
under the flight deck at the base of the auto throttle quadrant. Access to the switch pack is through the panels
forward of the nose gear doors.
The position switches form part of the auto-speed-brake control circuit. When the forward thrust levers are in less
than 50 percent total lever travel position, the auto-speed-brake system is operable. And when the levers are more
than 50 percent total lever travel position, the switches inhibit auto-speed-brake function.
IF 27-82
757 General Familiarization (7-2005)
ATA 27 - Page 127
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LEADING EDGE SLAT SYSTEM
Overview
The slats and the flaps increase airplane lift by increasing the wing surface. Increased wing surface lowers
airplane stall speeds. Lower airplane stall speeds allow lower takeoff and landing speeds. This reduces ground
roll, increases wheel and brake life, improves climb-out performance, and creates safer operating conditions.
There are ten slats, five in each wing. The slats are numbered 1 through 10, from left to right. Slats No. 1-4 and
No. 7-10 are the left and right sets of outboard slats. Slats No. 5 and No. 6 are the left and right inboard slats.
Automatic slat extension (autoslats), when fully extended, is used to enhance airplane stall characteristics when
flaps are in takeoff position (slats in intermediate position) and a stall warning signal is present.
The flap lever on the control stand operates the trailing edge flaps by cables between the flap lever and the flap
power drive unit (PDU). The slats are driven on primary or alternate power by one slat power drive unit (PDU) for
both inboard and outboard slats. The slat PDU is controlled on primary power by cables from the flap PDU. The
flap PDU follow-up cam positions the cable system which is connected to the slat PDU input cam. The slat PDU is
controlled on alternate power by a position selector switch. The selector switch also controls flaps. The selector
switch is operational after arming has been completed.
The slat PDU rotates torque tubes. The torque tubes operate rotary actuators geared to main tracks on each slat.
Movement of the main tracks extends or retracts the slats.
Primary power to rotate the torque tubes is hydraulic. Alternate power is electrical.
Full extension of slats from intermediate position is automatic (autoslats) to aid stall recovery.
Loss of slats No. 2-5 and No. 6-9 is detected by switches in the fixed wing leading edge.
Failure protection and annunciation for asymmetry, UCM (motion without command and motion opposite to
command), and command/position disagreement (no motion when commanded) is provided.
The inputs to and outputs from the flap/slat electronic units (FSEU-1, FSEU-2, and FSEU-3) assist in the control
of the flap/slat systems.
757 General Familiarization (7-2005)
ATA 27- Page 128
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
L.E. Slat Primary Control
Cables connected to the flap lever on the control stand drive the flap PDU input cam. Another set of cables is
connected from the flap PDU follow-up cam shaft to the input cam shaft on the slat PDU.
Operation of the flap lever will provide a controlled input to the hydraulic control valve on the flap PDU and to the
hydraulic control valve on the slat PDU.
L.E. Slat Alternate Control
Back-up operation of the slats is done with the flap/slat electronic unit (FSEU-2 and FSEU-3), an electric circuit,
the hydraulic bypass valve, and the electric motor.
There is one slat arming switch and one flap arming switch. One position selector switch controls positioning of
slats and flaps. The slat arming signal powers the slat hydraulic bypass valve to bypass (closed). The control
signal from the position selector switch is monitored by FSEU-2. FSEU-3 generates extend or retract signals to
activate power to drive the slat PDU electric motor and to cause the slat PDU electric motor clutch to engage.
The slat arming signal also powers the flap hydraulic bypass valve and the flap/slat hydraulic shutoff valve in the
depressurization module closed.
The slat hydraulic bypass valve, the flap hydraulic bypass valve, and the flap/slat shutoff valve in the
depressurization module are closed when the flap alternate drive arming switch is activated.
Primary Power
The slat drive primary power is from the left hydraulic system which is pressurized by an engine-driven pump on
the left engine. Hydraulic pressure drives the hydraulic motor on the slat PDU. In the event of a left engine failure,
the flaps and slats, as well as the landing gear, are powered by the power transfer unit.
A hydraulic control valve module is mounted on the slat power drive unit. The hydraulic motor is controlled by the
hydraulic control valve module, containing a hydraulic control valve and a dual-solenoid autoslat valve. The slat
PDU input cam is rotated by cable input initiated by the flap PDU follow-up cam. The input cam regulates the
hydraulic control valve by PDU linkages. The follow-up cam returns the hydraulic control valve to null.
Alternate Power
The slat drive alternate power is supplied by an electric motor on the slat PDU. The motor has a torque limiter and
a solenoid actuated clutch. The torque limiter protects against system overload. The motor clutch is de-energized
any time slat motion on alternate control is not commanded. The clutch engages the motor to the output shaft
when the solenoid is energized by arming the alternate drive.
757 General Familiarization (7-2005)
ATA 27 - Page 129
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Drive System
The slats are powered by two rotary actuators per slat. All rotary actuators are driven by a common torque tube
drive in the left and right wing. The left and right torque tube drives are connected to the output shaft on the slat
PDU.
The two main tracks on each slat have a gear sector that is driven by a pinion gear on each rotary actuator. The
forward end of each main track is connected to the nose of the slat. Slat auxiliary tracks are used to control the
slat angle.
Autoslats
Autoslat operation is initiated by powering of the dual-wound autoslat solenoid valve. The valve is a component of
the hydraulic control valve module on the slat PDU.
Presence of an autoslat extend signal energizes the autoslat solenoid valve and the valve directs hydraulic fluid to
move a hydraulic control valve sleeve in the extend direction. This causes the hydraulic motor to extend the slats
from intermediate to fully extended position.
Correction of airplane condition interrupts the autoslat extend signal. This results in retraction of the slats from
fully extended to intermediate position.
Slat Loss Detection Switches
Switches in the fixed wing leading edge control annunciation for loss of slats No. 2-5 and No. 6-9.
FSEU-1, FSEU-2, and FSEU-3
FSEU-1, FSEU-2, and FSEU-3 are identical and are installed in the main equipment center. The FSEUs are
physically isolated units. Each unit performs a separate function within the flap/slat systems. Each unit is capable
of performing any of the three functions.
The FSEUs receive signals for flap lever, flap, and slat position. The flap and slat alternate drive arming switches,
the alternate flap/slat position selector switch, and flap and slat PDU positions are also monitored by the FSEUs.
Parameters such as air speed and stall warning are sent to the FSEUs.
757 General Familiarization (7-2005)
ATA 27- Page 130
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-83
757 General Familiarization (7-2005)
ATA 27 - Page 131
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LEADING EDGE SLAT SYSTEM COMPONENTS
Slat PDU
The slat PDU drives all the leading edge slats. The PDU is located aft of the left wing strakelet and is mounted on
the left wing front spar. Access to the PDU is through access panels.
The slat PDU consists of two rotating control cams (an input cam and a follow-up cam), with followers and links.
Further components are a control valve input rod, a hydraulic control valve module (with a control valve and a
dual-solenoid autoslat valve), a hydraulic motor, an electric motor, a position transmitter, and a shock absorber.
Hydraulic Control Valve Module
A hydraulic control valve module is mounted on the slat PDU. The control valve module performs the PDU
hydraulic control functions. The module consists of a control valve and an autoslat valve. The left hydraulic system
is the source of hydraulic power.
The control valve controls the direction of rotation of the hydraulic motor according to the input by the cables at
the input cam shaft. The control valve will shut off hydraulic flow when the slats arrive at their selected position
as determined by the follow-up cam.
The autoslat valve is operated by a dual-solenoid. The dual-solenoid regulates the control valve for extension of
slats from intermediate to fully extended position, when an autoslat extension has been commanded. An autoslat
command will occur when slats are in intermediate position on hydraulic power and a stall condition happens. The
stall warning signal is sent from the left and/or right stall warning system to the FSEU units. Separate logic
circuits in FSEU-1 and FSEU-2 control the dual-solenoid autoslat valve.
Hydraulic Motor
A hydraulic motor is installed on the gearbox of the PDU for primary power to drive the slats.
Operation of the hydraulic motor will turn the gears in the gearbox and rotate the torque tubes to the slats.
Direction of rotation of the hydraulic motor is controlled by the control valve in the hydraulic control valve module.
Electric Motor
An electric motor is installed on the gearbox of the PDU. The motor uses alternate electric power to drive the slats.
(b) Operation of the motor turns the gears in the gearbox. The gearbox rotates the torque tubes to the slats.
Direction of rotation of the electric motor is controlled by the alternate flap/slat position selector switch. The
switch is on P3.
The motor has a torque limiter for protection against system overload and a solenoid actuated clutch which is deenergized any time slat motion on alternate control is not commanded. The clutch engages the motor to the output
shaft when the solenoid is energized.
757 General Familiarization (7-2005)
ATA 27- Page 132
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-84
757 General Familiarization (7-2005)
ATA 27 - Page 133
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Gearbox
Hydraulic or electric power is converted to mechanical power by the PDU gearbox. The gearbox rotates the torque
tubes to the slats.
The gearbox drives the follow-up cam and the position transmitter through gearing.
Position Transmitter
The input shaft on the transmitter is driven by follow-up cam rotation. Gears in the transmitter drive a brushless
phase shift resolver. The position signal is sent to FSEU-3. A logic circuit in FSEU-3 uses the position signal for
control of the PDU electric motor.
Shock Absorber
The shock absorber is installed on the PDU and transmits torque from the PDU to the slat drive torque tubes.
Functioning as a torque limiter/shock absorber, the unit limits torque to the drive system and absorbs rotational
kinetic energy from the PDU in an over-torque condition.
The shock absorber consists of an input gear, two output shafts, two torque limiters, and a shock absorber. The
input gear meshes with gearing in the PDU gearbox. The output shafts mate with couplings connected to the slat
drive torque tubes. The torque limiters are located at both ends of the unit. Each torque limiter has an overload
indicator. A ground lock pin feature is provided for manual insertion of a lock pin to prevent rotation of the slat
drive.
757 General Familiarization (7-2005)
ATA 27- Page 134
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Slat Bypass Valve
The bypass valve is installed in the left wing-body fairing forward of the wing. The valve is electric-motor operated
and serves to bypass hydraulic power when the slat alternate drive is armed and when slat asymmetry or slat uncommanded motion occurs. The valve can also be operated manually by a lever. The bypass mode of the valve is
indicated by Position 1. The normal mode is indicated by Position 2.
The valve consists of a 3-way, 2-position, rotary selector valve operated by an electric actuator. The valve has
three ports. In the event of alternate operation or protection shutdown, the valve is positioned so that the extend
port and retract port of the hydraulic motor are connected.
IF 27-85
757 General Familiarization (7-2005)
ATA 27 - Page 135
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap/Slat Depressurization Module
The flap/slat depressurization module is installed in the left main gear wheel well directly forward of the flap
PDU. The hydraulic flow to the flap and slat hydraulic motors passes through the depressurization module.
The primary function of the depressurization module is to remove hydraulic power from the flap and slat drive
systems during cruise or in the event of drive system failures. Depressurization at cruise reduces valve erosion
and minimizes the possibility of un-commanded flap/slat motion. In the event of a flap asymmetry condition or
un-commanded motion problem, the flap/slat systems are shut down hydraulically.
This module also gives flap and slat operation priority over landing gear operation. If pressure in the left hydraulic
system drops below 1800 PSI the priority valve closes, leaving a 3-gallon per minute orifice for the landing gear.
Within the module the flap/slat sequencing valve serves to reduce total flow to the flaps and slats when both
systems are operating at the same time. The valve normally serves as a flow limiter to the flap system. When slat
motion is commanded, flow to the flap system is reduced to allow adequate flow to the slat system. Total flow to
the flaps and slats is limited to assure adequate flow to other hydraulic systems.
IF 27-86
757 General Familiarization (7-2005)
ATA 27- Page 136
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Slat Drive Torque Tubes
Input between the slat PDU and the slat rotary actuators is by means of torque tubes. The torque tubes extend into
the left and right wing from the slat PDU along the forward side of the wing front spars. Within the fuselage the
torque tubes extend across the aft end of the forward cargo compartment. The torque tubes are hollow aluminum
shafts connected to each other and to the PDU, the angle gearboxes, and the rotary actuators by means of splined
couplings. The torque tubes are supported additionally by bearing supports.
IF 27-87
757 General Familiarization (7-2005)
ATA 27 - Page 137
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Angle Gearboxes
Angle gearboxes are used in the slat drive where the torque
rque tubes penetrate the fuselage left and right. The
gearboxes change torque tube alignment. Alignment is changed through bevel gears.
Rotary Actuators and Pinion Gears
Rotary actuators are installed in the leading edge of each wing. Each slat is operated by two rotary actuators. All
the rotary actuators are installed in the torque tube drive train extending left and right from the slat PDU. The
rotary actuators transfer torque to the geared slat main tracks by means of pinion gears. The pinion gears are
rotated by the rotary actuator output shafts.
The two rotary actuators at the left and right most outboard locations (slats No. 1 and 10 outboard rotary
actuators) provide a mounting pad for a slat position transmitter.
IF 27-88
757 General Familiarization (7-2005)
ATA 27- Page 138
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Slats
The geared slat main tracks associated with each rotary actuator are extended and retracted by the rotating pinion
gears driven off the rotary actuators. The forward end of each main track is connected to the slat nose. Extension
and retraction of the main tracks extend and retract the slats. The angle of each slat is controlled by auxiliary
tracks in the fixed leading edge.
Each main track at the rotary actuator locations incorporates a down stop and an up stop. The fixed leading edge
also incorporates up stops and down stops. The stops are used during rigging and testing of thee slats.
IF 27-89
757 General Familiarization (7-2005)
ATA 27 - Page 139
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Slat Loss Detection
The loss of slats No. 2-5 and No. 6-9 is detected by switches in the fixed wing leading edge. Loss of any or all of
the affected slats will ground an electrical circuit to FSEU-1 and FSEU-2 and cause failure annunciation (LE SLAT
ASYM message on EICAS display and LEADING EDGE light on P3).
IF 27-90
757 General Familiarization (7-2005)
ATA 27- Page 140
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27 - Page 141
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FSEU-1, FSEU-2, and FSEU-3
Monitoring and control of the flap and slat drive systems is done electrically by the flap/slat electronics unit.
The FSEU consists of 3 interchangeable modules; FSEU-1, FSEU-2 and FSEU-3. The 3 modules are physically and
functionally identical but they perform different functions based upon where they are installed in the equipment
rack. Each FSEU module has a microprocessor programmed to perform all FSEU functions. When the module is
installed in one of three locations in the equipment rack the wiring from the electrical connectors enables only a
specific channel, defining the module's function.
FSEU-1 performs all primary flap and slat drive system protection, indication and automatic operation. Specifically,
FSEU-1 provides:
1.
2.
3.
4.
5.
6.
7.
Un-commanded motion detection and protection of the flaps during hydraulic operation
Asymmetry detection and protection during hydraulic operation
Flap load relief functions
Autoslat controls
Flaps up depressurization
Slat position indication signals
Signals to other aircraft systems
FSEU-2 performs alternate flap and slat drive protection and annunciation.
Specific functions are:
1.
Un-commanded motion detection under alternate control
2.
Asymmetry detection under alternate control
3.
Autoslat control
4.
Position indication
5.
Signals to other aircraft systems
FSEU-3 performs alternate flap and slat drive control.
Specific functions are:
1.
2.
Alternate electric drive motor control
Signals to other aircraft systems
Built in Test Equipment (BITE) electrically tests the FSEU and interfacing electronic components. BITE software is
designed to isolate faults in the electronic system down to the line replaceable unit level.
The BITE test is activated by pressing a switch on the front of each FSEU box. The BITE test is self terminating
and the results of the test are shown on the FSEU front panel.
757 General Familiarization (7-2005)
ATA 27- Page 142
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-91
757 General Familiarization (7-2005)
ATA 27 - Page 143
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 27- Page 144
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LEADING EDGE SLAT POSITION INDICATING SYSTEM
Overview
Indication of slat position is by the flap/slat position indicator on the first officer's main instrument panel P3.
Absence of fault annunication indicates correct position of slats as related to command by the flap lever on
primary power or the alternate flap/slats position selector switch on alternate power.
Flap/Slat Position Indication on Primary Power
The logic in FSEU-1 controls activation and deactivation of a relay in the flap/slat position indicator circuit. The
relay controls transmission of a fixed 1/2-unit signal to the indicator.
When the flaps and the slats are fully retracted, the indicator needles are in UP, because the logic in FSEU-1
deactivates the relay to allow transmission of the fully-retracted signal from the flap position transmitters to the
indicator.
When either the flaps or the slats extend from the fully retracted position, the indicator needles jump from UP to
1/2, because the logic in FSEU-1 activates the relay to allow transmission of the fixed 1/2-unit signal to the
indicator.
When the flaps are in the 1-degree position and the slats are in the intermediate position, the indicator needles
jump from 1/2 to 1, because the logic in FSEU-1 deactivates the relay to allow transmission of the position signals
from the flap position transmitters to the indicator.
The needles indicate flap position only for flap lever between greater than 1-unit detent and 30-unit detent.
When the flaps retract from the 1-unit position, the indicator needles jump from 1 to 1/2, because the logic in
FSEU-1 activates the relay to allow transmission of the fixed 1/2-unit signal to the indicator.
The indicator needles jump from 1/2 to UP when the flaps and the slats are fully retracted because the logic in
FSEU-1 deactivates the relay in order to allow the transmission of the fully-retracted signal from the flap position
transmitters to the indicator.
757 General Familiarization (7-2005)
ATA 27 - Page 145
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap/Slat Position Indication on Alternate Power
The logic in FSEU-2 controls activation and deactivation of a relay in the indicator circuit. The relay controls
transmission of a fixed 1/2-unit signal to the indicator.
If the alternate flap and alternate slat drive systems are armed at the same time, the flap/slat position indication
is the same as on primary power, except that the indicator circuit relay is controlled by logic in FSEU-2.
If the alternate slat drive system is armed by itself, the indicator reflects slat position as follows:
(a) When the slats are extended to intermediate or fully extended positions, the indicator needles jump from
UP to 1/2, because the logic in FSEU-2 activates the relay to allow transmission of the fixed 1/2-unit
signal to the indicator.
(b) When the slats are retracted, the indicator needles jump from 1/2 to UP, because the logic in FSEU-2
deactivates the relay to allow transmission of the fully-retracted signal from the flap position transmitters
to the indicator.
Takeoff Warning
If the slats are not in takeoff position (intermediate position) as commanded by the flap lever and the thrust levers
are advanced into the takeoff range with the engines running, takeoff warning is announced as follows:
(a)
(b)
(c)
(d)
A red (level A) FLAPS message will appear on the EICAS display.
The master warning lights on P2 will come on.
The aural warning will sound.
The red CONFIGURATION light on P2 will come on.
757 General Familiarization (7-2005)
ATA 27- Page 146
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flap/Slat Position Indication on Alternate Power
The logic in FSEU-2 controls activation and deactivation of a relay in the indicator circuit. The relay controls
transmission of a fixed 1/2-unit signal to the indicator.
If the alternate flap and alternate slat drive systems are armed at the same time, the flap/slat position indication
is the same as on primary power, except that the indicator circuit relay is controlled by logic in FSEU-2.
If the alternate slat drive system is armed by itself, the indicator reflects slat position as follows:
1.
When the slats are extended to intermediate or fully extended positions, the indicator needles jump
from UP to 1/2, because the logic in FSEU-2 activates the relay to allow transmission of the fixed
1/2-unit signal to the indicator.
2.
When the slats are retracted, the indicator needles jump from 1/2 to UP, because the logic in FSEU-2
deactivates the relay to allow transmission of the fully-retracted signal from the flap position
transmitters to the indicator.
Takeoff Warning
If the slats are not in takeoff position (intermediate position) as commanded by the flap lever and the thrust levers
are advanced into the takeoff range with the engines running, takeoff warning is announced as follows:
IF 27-92
757 General Familiarization (7-2005)
ATA 27 - Page 147
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Slat Position Transmitters
One slat position transmitter is in the left wing, installed on the outboard rotary actuator of slat No. 1. Another slat
position transmitter is in the right wing installed on the outboard rotary actuator of slat No. 10.
IF 27-93
757 General Familiarization (7-2005)
ATA 27- Page 148
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
B757 GENERAL FAMILIARIZATION
ATA 28
FUEL SYSTEM.......................................................................................................................................... 3
Overview............................................................................................................................................. 3
Fuel Tanks........................................................................................................................................... 3
Venting................................................................................................................................................ 3
Fueling ................................................................................................................................................ 3
Engine and APU Fuel-Feed ................................................................................................................ 4
De-fueling ........................................................................................................................................... 4
Fuel Quantity Indicating and Measurement........................................................................................ 4
System Protection and Fault Isolation ................................................................................................ 4
Fuel Temperature Indication............................................................................................................... 4
Fuel Pressure Indicating...................................................................................................................... 4
FUEL TANKS ............................................................................................................................................ 5
Overview............................................................................................................................................. 5
Fuel Tank Construction ...................................................................................................................... 5
Main Fuel Tanks ................................................................................................................................. 5
Center Fuel Tank ................................................................................................................................ 6
Surge Tank .......................................................................................................................................... 6
FUEL TANK COMPONENTS .................................................................................................................. 8
Access Doors ...................................................................................................................................... 8
Over-wing Fill Ports ......................................................................................................................... 11
Sump Drain Valves ........................................................................................................................... 12
Baffle Rib Check Valve .................................................................................................................... 13
PRESSURE FUELING SYSTEM ............................................................................................................ 14
Overview........................................................................................................................................... 14
PRESSURE FUELING SYSTEM COMPONENTS ................................................................................ 16
Fuel Level Sensor ............................................................................................................................. 16
Control Card for the Fuel Level Sensor ............................................................................................ 16
Fueling Shutoff Valve ....................................................................................................................... 18
Normal Operation ............................................................................................................................. 18
Manual Operation ............................................................................................................................. 18
Fueling Adapters ............................................................................................................................... 20
Fueling Control Panel ....................................................................................................................... 21
Drain Check Valves for the Fueling Manifold DEFUEL ................................................................. 23
ENGINE FUEL FEED SYSTEM ............................................................................................................. 24
Overview........................................................................................................................................... 24
ENGINE FUEL FEED COMPONENTS ................................................................................................. 26
Engine Fuel Shutoff Valve ............................................................................................................... 26
Engine Fuel Cross feed Valve .......................................................................................................... 27
Main Tank Fuel Boost Pump ............................................................................................................ 29
Center Tank Override Pump ............................................................................................................. 30
Fuel Boost Pump Discharge Check Valve ........................................................................................ 30
Fuel Boost Pump Removal Check Valve ......................................................................................... 30
757 General Familiarization (7-2005)
ATA 28 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Tank Fuel Boost Pump Bypass Valve ..................................................................................... 32
Automatic Sumping Jet Pump .......................................................................................................... 33
Fuel Scavenge System ...................................................................................................................... 33
AUXILIARY POWER UNIT (APU) FUEL FEED SYSTEM ................................................................ 34
Overview........................................................................................................................................... 34
APU Fuel Feed System ..................................................................................................................... 34
APU FUEL FEED SYSTEM COMPONENTS ....................................................................................... 36
DC Fuel Pump .................................................................................................................................. 36
APU Shutoff Valve ........................................................................................................................... 37
APU Fuel Line and Shroud............................................................................................................... 39
APU Flame Arrestor ......................................................................................................................... 40
APU Sump Drain Valve ................................................................................................................... 40
APU Fuel Check Valve..................................................................................................................... 40
Pressure Switches ............................................................................................................................. 40
DEFUELING SYSTEM ........................................................................................................................... 42
Overview........................................................................................................................................... 42
De-fueling ......................................................................................................................................... 42
Tank-To-Tank Fuel Transfer ............................................................................................................ 42
DEFUELING SYSTEM COMPONENTS ............................................................................................... 45
De-fuel Valve.................................................................................................................................... 45
FUEL QUANTITY INDICATING SYSTEM (FQIS) ............................................................................. 46
Overview........................................................................................................................................... 46
FUEL QUANTITY INDICATING SYSTEM (FQIS) COMPONENTS ................................................. 48
Tank Units......................................................................................................................................... 48
Compensator ..................................................................................................................................... 49
Densitometer ..................................................................................................................................... 50
Center Tank Densitometer Fuel Retention System........................................................................... 51
Fuel Quantity Indicator ..................................................................................................................... 52
FUEL QTY Test Switch ................................................................................................................... 52
Load Select Indicator ........................................................................................................................ 54
Load Select Control .......................................................................................................................... 55
Fuel Quantity Processor Unit (FQPU) .............................................................................................. 56
FUEL MEASURING STICKS ................................................................................................................. 57
Overview........................................................................................................................................... 57
FUEL MEASURING STICKS COMPONENTS ..................................................................................... 57
Fuel Measuring Sticks ...................................................................................................................... 57
757 General Familiarization (7-2005)
ATA 28 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL SYSTEM
Overview
The airplane fuel system keeps fuel and distributes it to the engines and auxiliary power unit (APU). The fuel
system also has equipment for pressure fueling and de-fueling and for fuel quantity measurement.
Fuel Tanks
Two main fuel tanks and one center fuel tank keep fuel for the engines and APU. A surge tank near each wing tip
supplies main tank overflow containment and venting. A drain check valve installed inboard of each surge tank lets
fuel collected in the surge tanks return to the main fuel tanks but prevents fuel from the main fuel tank from going
into the surge tanks. Over-wing fill ports for each main fuel tank permit manual overwing fueling. Access doors on
the wing lower surface permit access into the fuel tanks for inspection or component repair. Sump drain valves at
the low point of each fuel tank permit the drainage of fuel tank contaminants, condensation, and fuel.
Venting
The fuel vent system supplies venting from each fuel tank to the surge tanks to prevent pressure/vacuum buildup. The surge tanks vent out of the fuel tanks through a vent scoop. A pressure relief valve, in each surge tank,
supplies backup venting if flame arrester blockage occurs. Float valves prevent fuel from going into the vent
channels. A flame arrestor also prevents external flame from going into the surge tank through the vent scoop.
Fueling
The fueling system supplies automatic and manual fueling capability. The fueling control panel, P28, on the right
wing leading edge supplies control for the fueling process. Four fueling shutoff valves, one for each main and two
for the center fuel tank, control the supply of fuel from the fueling adapters to the three fuel tanks. A load select
control, controlled by the Fuel Quantity Indication System (FQIS) processor, supplies automatic control of fueling.
The load select control automatically closes the applicable fueling shutoff valves when the selected fuel load is
reached.
The volumetric top-off function and the overfill control function supply backup for the load select system. The
volumetric top-off, controlled by the FQIS processor, closes the fueling shutoff valves when the fuel tanks are full.
The overfill control has an fuel level sensor in each surge tank and a control card for the fuel level sensor. The
overfill system closes the fueling shutoff valves when the fuel level sensor detects fuel entering the surge tanks.
Over-wing fill ports supply access for manual fueling of each main fuel tank.
757 General Familiarization (7-2005)
ATA 28 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine and APU Fuel-Feed
The fuel-feed systems supply fuel to the main engines and the APU. Four fuel boost pumps (AC), two in each main
fuel tank, and two fuel override pumps (AC), in the center fuel tank, usually supply fuel to the engines. The left
forward fuel boost pump in the left main fuel tank usually supplies fuel to the APU. Switch lights on the fuel
management control panel on the P5 panel control the fuel boost pumps (AC).
An APU DC fuel pump supplies backup fuel-feed to the APU if the left forward fuel boost pump (AC) fails or AC
power is not available. A pressure switch for APU fuel shutoff, found in the engine fuel-feed manifold in the left
main fuel tank, supplies pressure information for the AC/DC pump control circuit during APU operation. A cross
feed system supplies fuel through a engine fuel cross feed valve from any fuel tank to either engine. A switch light
on the P5 panel controls the engine fuel cross feed valve.
De-fueling
Two de-fueling valves let fuel go into the fueling manifold from the engine fuel-feed manifold for de-fueling. The
main fuel tanks can be suction de-fueled or pressure de-fueled. The center fuel tank must be pressure de-fueled.
A switch for each de-fueling valve on the fueling control panel, P28, controls the de-fueling valves.
Fuel Quantity Indicating and Measurement
The fuel quantity indicating system has tank units, compensators, and densitometers that supply fuel volume and
density measurements to the FQIS processor. The FQIS processor uses these measurements to compute the weight
of the fuel in each fuel tank. The fuel quantity indicator on the P5 panel shows the fuel weight in each fuel tank as
well as the total fuel quantity. Fuel measuring sticks attached to the inside of some fuel tank access doors supply
a means to measure fuel quantity manually.
System Protection and Fault Isolation
Fuel system protection and fault isolation consists of the Engine Indication and Crew Alerting System (EICAS) and
the FQIS processor functions. EICAS supplies status, caution, advisory faults or maintenance messages during
airplane operation and for later recall by ground maintenance crews. The FQIS processor supplies protection and
fault isolation only for the FQIS portion of the fuel system.
Fuel Temperature Indication
A fuel temperature sensor in the right main fuel tank supplies temperature input to the fuel temperature indicator
on the P5 panel.
Fuel Pressure Indicating
A pressure switch connected to each fuel boost pump and fuel override pump supplies low pressure indication to
each EICAS computer and to a low pressure light on the P5 panel.
757 General Familiarization (7-2005)
ATA 28 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL TANKS
Overview
Three fuel tanks, left main, right main, and center, keep the fuel necessary to run the engines and auxiliary power
unit (APU). A surge tank, outboard of each main fuel tank, collects fuel overflow from the fuel tanks and includes
equipment for venting the fuel tanks overboard.
The fuel tanks hold the necessary equipment for fueling, de-fueling, and engine fuel-feed. Equipment necessary for
APU fuel-feed and fuel quantity indicating is also contained in the fuel tank structure.
Fuel Tank Construction
The 757 uses primary wing structure for the airplane fuel tanks. The fuel tanks are installed between the front and
rear wing spars and between the upper and lower wing skin. Solid "tank end" ribs close the ends of each fuel tank,
while all other wing ribs act as fuel baffles to minimize fuel slosh.
The wing structure permits inboard fuel flow. Fuel sump areas are found at the lowest point of each fuel tank.
Sump drain valves at these sump areas permit unwanted materials to be drained from the fuel tank. The sump
drain valves also permit removal of all remaining fuel after a fuel tank is de-fueled.
All fuel tanks are fluid tight. Close metal to metal fit of all parts forms the basic seal. BMS 5-26 sealing compound
and sealed fasteners are used on all joints to complete the fluid tight seal.
Forty-five oval shaped access doors on the wing lower surface supply access to the fuel tanks.
Main Fuel Tanks
Each main fuel tank extends from the rib just inboard of the bypass valve of the fuel boost pump outboard to the
rib just inboard of the pressure relief valve for the surge tank. A baffle rib just outboard of each bypass valve
contains 11 baffle rib check valves. The check valves prevent outboard fuel flow while they permit inboard fuel
flow. The check valves make sure that the boost pump inlets remain covered with fuel when the airplane is in a
roll attitude.
Each main fuel tank holds 2176 U.S. gallons of fuel of which 4 gallons are unusable.
Sixteen access doors on the wing lower surface supply access to each main fuel tank.
A dry bay directly above the engine protects the fuel tank from hot engine fragments in the case of an engine
burst. The dry bay contains one or four, drain holes each with a flame arrestor to drain all moisture. Access panels
exist for internal inspections.
An over-wing fill port, found on the upper wing surface permits gravity feed fueling if pressure fueling equipment
is not available.
757 General Familiarization (7-2005)
ATA 28 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Center Fuel Tank
The center fuel tank is found between the inboard tank end ribs of the left and right main fuel tanks.
Seven access doors on the lower tank surface supply access to the center fuel tank.
There are 6,901 U.S. gallons of usable fuel and 35 U.S. gallons of unusable fuel in the center fuel tank.
Surge Tank
A surge tank, found outboard of each main fuel tank, contains all fuel over-flow from the main fuel tank and
supplies fuel tank venting. A drain check valve in the surge tank permits the fuel overflow to drain back into the
main fuel tank but prevents fuel flow from the main fuel tank to the surge tank.
757 General Familiarization (7-2005)
ATA 28 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-1
757 General Familiarization (7-2005)
ATA 28 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL TANK COMPONENTS
Access Doors
Forty-five access doors permit entry into the airplane fuel tanks for inspection or component repair. The oval
shaped access doors fit in each fuel tank over cutouts in the lower wing surface. For all except one access door, a
clamp ring with a knitted aluminum gasket bolted to the access door from outside the wing holds the access door
in its position. The gasket forms an electrical bond between the access door and the wing skin. A molded rubber
seal ring prevents fluid leakage around the access door. The remaining access door (vent scoop door) bolts to a
nut plate attached to the inside of the lower wing surface. A rubber seal and an O-ring prevent fuel leakage.
Removal of an access door requires fuel tank de-fueling.
Each main fuel tank has 16 access doors. Each access door measures 10 x 18 inches except for three farthest
outboard access doors. These access doors measure 8 x 18 inches. Five main tank access doors have fuel
measuring sticks attached to their inside surface.
The center fuel tank has 7 access doors. Each access door measures 10 x 18 inches. Two access doors have fuel
measuring sticks attached to their inside surface. The center fuel tank has 3 baffle doors in the center section of
the fuel tank.
Each surge tank has three access doors. Two access doors measure 8 x 18 inches. A housing of the vent flame
arrestor is attached to the inside of the center access door (vent scoop door) which is 10 x 18 inches. A pressure
relief valve is attached to the inside of the most inboard access door.
IF 28-2
757 General Familiarization (7-2005)
ATA 28 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-2A
757 General Familiarization (7-2005)
ATA 28 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 28 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Over-wing Fill Ports
An over-wing fill port is installed for each main fuel tank. Each over-wing fill port is found on the upper wing
surface about two thirds of the way out on the wing. The over-wing fill ports permit an alternate way of fueling the
main fuel tanks if pressure fueling equipment is not available.
Each over-wing fill port has an adapter, seal ring, two O-rings, a retaining nut, and a quick release filler cap. Lift
and turn the handle on the filler cap counterclockwise to unlock the filler cap. It is then lifted from the adapter.
Installation reverses the procedure and the downward turn of the handle locks and seals the filler cap against the
adapter.
IF 28-3
757 General Familiarization (7-2005)
ATA 28 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Sump Drain Valves
Six sump drain valves permit manual drainage of accumulated water and sediment from the fuel tanks. One sump
drain valve is installed flush to the lower wing surface at the low point of each main fuel and surge tank. Two
sump drain valves are installed flush to the lower surface of the center fuel tank. Drain lines attached to the sump
drain valves in the center fuel tank extend to drain line fittings on the wing-to-body fairings.
A separate flapper closes over the sump drain valve hole in the wing skin when the sump drain valve is removed
from the airplane. This permits removal of the sump drain valves without fuel tank de-fueling.
IF 28-4
757 General Familiarization (7-2005)
ATA 28 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Baffle Rib Check Valve
Eleven baffle rib check valves are installed on the baffle rib just inboard of each fueling shutoff valve in the main
fuel tank. The check valves prevent fuel flow in the outboard direction but permit fuel flow in the inboard direction.
This prevents a low fuel level between the baffle rib and the inboard tank end rib in the main fuel tank. Thus the
check valves prevent the boost pump inlets from being uncovered.
Each check valve is a free swinging flow actuated valve. Inboard fuel flow opens the check valves and outboard
flow forces the check valves closed.
IF 28-5
757 General Familiarization (7-2005)
ATA 28 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PRESSURE FUELING SYSTEM
Overview
The pressure fueling system supplies a fast procedure to fill the fuel tanks in the airplane. The FQIS processor
controls the fueling operation in the automatic mode. The fueling crew can operate the system manually if it is
necessary.
The fueling station, found on the leading edge of the right wing, has a fueling control panel and two fueling
adapters. The fueling adapters attach on the front spar and connect to the fueling manifold. The fueling adapters
permit fueling and de-fueling of the fuel tanks.
A 3 inch diameter fueling manifold extends from the fueling adapters toward the rear spar and across the airplane
to the left wing. The fueling manifold moves fuel from the fueling adapters through fueling shutoff valves into each
fuel tank. Two drain check valves are installed in the fueling manifold in the center fuel tank. The drain check
valves permit drainage of trapped fuel in the fueling manifold into the fuel tank and prevent fuel from the fuel tank
from entering the fueling manifold.
Four (pressure) fueling shutoff valves are installed between the fueling manifold and its branch lines. The branch
lines direct fuel into the individual fuel tanks. The fueling shutoff valves control the flow of fuel from the main
manifold into its branch lines. De-fueling valves connect the engine fuel-feed manifold with the (pressure) fueling
manifold.
The fueling control panel, P28, controls the fueling system. The fueling control panel has all the switches and
indicators necessary to do the fueling operation.
NOTE:
If you can not read the nomenclature on the fueling control panel, you can install
placards.
An overfill protection system has a fuel level sensor, on the rear spar of each surge tank, and a control card for
the fuel level sensor, in the electrical systems card file, P50, protects the fueling operation from overflow. The
system will shut down the fueling operation if the fuel level sensors sense fuel in the surge tank.
757 General Familiarization (7-2005)
ATA 28 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-6
757 General Familiarization (7-2005)
ATA 28 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PRESSURE FUELING SYSTEM COMPONENTS
Fuel Level Sensor
A fuel level sensor attaches to the rear spar of each surge tank. The fuel level sensor has a probe with a small
piezo-electric element sealed in the probe and an electrical connector. The probe, which extends into the fuel tank,
and the element form an electromechanical circuit. This circuit sets up acoustic vibrations in the probe when 28volts DC from the control card for the fuel level sensor energizes the element. With fuel around the probe the
circuit frequency decreases. The control card monitors changes in frequency and controls the fueling shutoff valves
accordingly.
Control Card for the Fuel Level Sensor
A control card for the fuel level sensor is installed in the electrical systems card file, P50, in the
electrical/electronics bay. The control card supplies control logic to protect the fuel system from overfilling during
fueling. The control card operates on 28-volts DC power and supplies drive signals to each fuel level sensor. The
fuel level sensors supply frequency outputs back to the control card which controls the refueling valve relays. The
frequency level that the control card senses determines whether the control card logic interrupts power to the
refueling valve relays.
757 General Familiarization (7-2005)
ATA 28 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-7
757 General Familiarization (7-2005)
ATA 28 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fueling Shutoff Valve
The fueling shutoff valve has two components; the valve, and the actuator. The two components connect through
an opening in the rear spar.
The actuator for the fueling shutoff valve is installed out of the fuel tank on the rear spar. The actuator engages
with the two valve control ports which protrude through the rear spar.
Actuator replacement does not require de-fueling.
Normal Operation
When fuel under pressure enters the valve inlet, fuel pressure and spring tension keep the fueling shutoff valve in
its closed position.
When the solenoid on the Actuator of the Fueling Shutoff Valve is energized fuel pressure forces the valve open.
To close the fueling shutoff the solenoid is de-energized and the fuel pressure inside the valve equalizes allowing
the spring loaded poppet to close the valve. Spring pressure holds the poppet closed with pressure removed from
the fueling manifold.
The valve position switch supplies the fueling station with an indication of the fueling shutoff valve position.
Manual Operation
CAUTION:
MAKE SURE THERE IS NO PRESSURE IN THE FUELING MAINFOLD BEFORE YOU TURN
THE MANUAL OVERRIDE KNOB. DAMAGE TO THE VALVE OR THE MANIFOLD CAN
OCCUR.
The fueling shutoff valve opens or closes manually if you turn the manual override knob 10 to 13 complete
revolutions. Clockwise rotation closes the fueling shutoff valve. Counterclockwise movement opens it. A lock wire
holds the manual override knob in the closed position.
NOTE:
757 General Familiarization (7-2005)
The attach screws that attach the cover plate over the manual override knob should not
be tampered with. Tampering with the attach screws on the cover plate could cause fuel
leaks.
ATA 28 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-8
757 General Familiarization (7-2005)
ATA 28 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fueling Adapters
Two fueling adapters attach to the front spar of the right wing at the fueling station.
n. The fueling adapters permit
pressure fueling and de-fueling of the airplane fuel tanks. The fueling adapters have a housing which contains a
poppet, piston and a FUEL/DEFUEL cam. The adapters mate with a standard type fueling nozzle.
With the FUEL/DEFUEL cam in the FUEL position, fuel pressure forces the poppet to open. The open poppet
prevents reverse fuel flow when the fueling operation is complete.
With the FUEL/DEFUEL cam in the DEFUEL position, the poppet and piston interlock to permit reverse fuel flow
when the fueling nozzle is attached.
IF 28-9
757 General Familiarization (7-2005)
ATA 28 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fueling Control Panel
The fueling control panel, P28, is installed forward of the front spar on the right wing between the two fueling
adapters. The fueling control panel has switches and indicators for fueling and de-fueling operations.
IF 28-10
757 General Familiarization (7-2005)
ATA 28 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 28 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Drain Check Valves for the Fueling Manifold DEFUEL
Two drain check valves, found in the center fuel tank, attach to the lower side of the fueling manifold. The drain
check valves permit drainage of the fueling manifold into the center fuel tank. The drain check valve has an inlet
port, outlet port, a hinged flapper, and piston, in its cylindrical housing. The flapper prevents reverse flow of fuel
from the fuel tank to the fueling manifold when fuel tank pressure exceeds the manifold pressure. The sliding
piston closes the drain check valve when manifold pressure exceeds fuel tank pressure by 8.5 PSIG or more.
Some aircraft may be equipped with a fueling manifold vent valve which attaches to the right fueling manifold. The
vent valve allows air to be vented from the fuel tank into the fueling manifold to drain the trapped fuel into the
fuel tank.
IF 28-11
757 General Familiarization (7-2005)
ATA 28 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FUEL FEED SYSTEM
Overview
Three fuel tanks store fuel for operating the engine or APU.
Each engine receives fuel from the left or right main fuel tank and the center fuel tank. Opening the cross feed
valve(s) permits either engine to be supplied by any tank. An engine fuel shutoff valve controls fuel flow from each
fuel tank to each engine.
Four electric motor driven fuel boost pumps supply fuel at pressure to the engines. Each main boost pump mounts
to the forward or rear spar of the center wing fuel tank. These four boost pumps draw fuel from the main fuel
tanks. In addition, each main engine fuel feed subsystem has a boost pump bypass valve to allow suction feed to
the engine should a boost pump fail.
Two override boost pumps allow fuel feed from the center fuel tank. The pumps provide fuel to the respective left
or right halves of the engine fuel feed manifold at a pressure higher than the main boost pumps. This allows the
center tank fuel to be used before fuel is used from the main fuel tanks.
Automatic sumping jet pumps scavenge water/contaminants from the fuel tank sump areas. The jet pumps,
powered by motive flow from the main boost pumps, direct the scavenged discharge to the inlet of the boost pump.
The scavenge is mixed with fuel and sent to the engine.
On some aircraft, a fuel scavenge ejector pump, powered by motive flow from the left forward boost pump,
transfers fuel from the center tank sump area through a float valve and discharges the fuel into the left main tank.
The float valve allows fuel to be discharged into the left main tank when the left main tank is less than half full.
The fuel management module, on the pilot's overhead panel P5, contains the control and indication for the engine
fuel feed system. The Engine Indication and Crew Alerting System (EICAS) also provides indication of engine fuel
feed system malfunctions. The fuel control panel on the quadrant stand P10 also contains indication of engine fuel
shutoff valve malfunction.
Three phase 115-volts AC, 400 Hertz electrical current powers the boost pumps. Twenty-eight DVC powers the
valves and control circuits.
757 General Familiarization (7-2005)
ATA 28 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-12
IF 28-12A
757 General Familiarization (7-2005)
ATA 28 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FUEL FEED COMPONENTS
Engine Fuel Shutoff Valve
The engine fuel shutoff valve is a motor actuated butterfly type valve. An engine fuel shutoff valve actuator mounts
to the rear spar of each main fuel tank. The valve body (within the fuel tank) connects to the actuator (outside the
fuel tank) by an adapter shaft. Access to the valve body requires de-fueling.
The square (preferred) actuator, has a 28-volts DC permanent magnet motor, an electrical connector, "OPENCLOSED" limit switches, and a manual override handle. The manual override handle also shows the position of the
engine fuel shutoff valve.
The round (alternate) actuator consists of a 28-volts DC motor, an electrical connector, open-closed limit switches,
and a manual override handle which also serves as a valve position indicator.
The adapter shaft connects the actuator and the valve body. The adapter shaft consists of an adapter plate and a
shaft with a universal joint. The adapter shaft has indexed splines at both ends. This ensures correct alignment
between the actuator and the valve. Access to the adapter shaft requires defueling.
IF 28-13
757 General Familiarization (7-2005)
ATA 28 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Fuel Cross feed Valve
The engine fuel cross feed valve(s) is (are) a motor actuated butterfly valve(s). The engine fuel cross feed valve(s)
mounts to the rear spar in the center wing fuel tank. The valve(s) body (within the fuel tank) attaches to the
actuator (outside the fuel tank) by an adapter shaft. Access to the valve(s) body requires de-fueling.
The square (preferred) actuator has a 28-volts DC permanent magnet motor, an electrical connector, "OPENCLOSED" limit switches, and a manual override handle. The manual override handle also shows the position of the
engine fuel cross feed valve.
The round (alternate) actuator has a 28-volts DC motor, an electrical connector, open-closed limit switches, and a
manual override handle which serves as a position indicator.
The adapter shaft connects the actuator and the valve body. The adapter shaft consists of an adapter plate and a
shaft with a universal joint. The adapter shaft has indexed splines in both ends. This ensures correct alignment of
the actuator to the valve. Access to the adapter shaft requires de-fueling.
The valve isolates or connects the left and right fuel feed manifold sections allowing any fuel tank boost pump to
pressure feed either engine.
IF 28-14
757 General Familiarization (7-2005)
ATA 28 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 28 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Tank Fuel Boost Pump
Two main tank fuel boost pumps mount to the front and rear spars of the center wing fuel tank. Each pump
consists of a motor/impeller unit installed within a housing. An inlet port, a discharge port, and a primer
discharge port and check valve form part of the housing. A fuel boost pump discharge check valve mounts to the
housing discharge port. A fuel boost pump removal check valve mounts to the housing inlet port. The removal
check valve permits removal of the motor/impeller unit without de-fueling. The motor/impeller unit contains the
motor, all electrical fittings, and the inlet and primer impellers mounted on a motor shaft. The pump provides
about 20,000 pounds of fuel per hour at 10 PSI minimum.
IF 28-15
757 General Familiarization (7-2005)
ATA 28 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Center Tank Override Pump
Two center tank override pumps mount to the rear spar of the center wing fuel tank. Each pump consists of a
motor/impeller unit installed within a housing unit. An inlet port, a discharge port, and a primer discharge port
and check valve form part of the housing. A fuel boost pump discharge check valve mounts to the housing
discharge port. A fuel boost pump removal check valve mounts to the housing inlet port. With the fuel boost pump
removal check valve in place, removal of the motor/impeller unit does not require de-fueling. The motor/impeller
unit contains the motor, all electrical fittings, and the inducer and inlet impeller mounted on the motor shaft. The
pump provides about 28,000 pounds of fuel per hour at 27 PSI minimum.
Fuel Boost Pump Discharge Check Valve
A fuel boost pump discharge check valve mounts to the discharge port of each main tank fuel boost pump and
each center tank override pump. The check valve, a spring loaded closed flapper valve, allows normal fuel flow
from the boost pump discharge to the engine and prevents reverse fuel flow through the pump.
Fuel Boost Pump Removal Check Valve
A fuel boost pump removal check valve mounts to the inlet port of each main tank fuel boost pump and each
center tank override pump. The check valve, a spring loaded closed poppet valve, prevents fuel drainage through
the boost pump housing when the motor/impeller unit is removed. Removal of the motor/impeller unit allows the
spring to seal the poppet against the seat to close the fuel path. Reinstallation of the fuel boost pump forces the
poppet open against the spring to open the fuel flow path.
757 General Familiarization (7-2005)
ATA 28 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-16
757 General Familiarization (7-2005)
ATA 28 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Tank Fuel Boost Pump Bypass Valve
If the fuel boost pumps fail, the two bypass valves of the fuel boost pumps let suction fuel feed to the engines.
The bypass valves contain a filter screen
een and a check valve body with a hinged flapper. One bypass valve connects
to the engine fuel feed manifold for each engine.
IF 27-17
757 General Familiarization (7-2005)
ATA 28 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Automatic Sumping Jet Pump
There is an automatic sumping jet pump at each forward fuel boost pump inlet (if installed), in the main fuel
tanks. There is an automatic sumping jet pump at each aft tank boost pump inlet, in the main fuel tanks. There is
an automatic sumping jet pump at each override pump inlet, in the center fuel tank. There is a fuel transfer jet
pump at each override pump inlet (if installed),
alled), in the center fuel tank. Each jet pump will scavenge water when a
fuel pump operates.
The operating boost pump pressurizes the motive port of the jet pump. Flow through the jet nozzle draws fuel and
scavenged water through the inlet screen to the induced port and out through discharge port.
Fuel Scavenge System
On some aircraft, a fuel scavenge ejector pump in the center tank provides scavenging whenever the left main tank
is below half full. The fuel scavenge system is controlled by a float operated shutoff valve in the left main tank
that opens when the tank drops below half full.
The ejector pump consists of a housing with three ports: a motive port, an induced port, and a discharge port. The
motive port connects to the left forward main boost pump primer discharge port. The induced port connects to an
inlet screen sealed to the lower surface of the center fuel tank. The discharge port is connected by tubing to the
float operated shutoff valve in the left main tank.
The float operated shutoff valve consists of a housing with an inlet port, discharge port, and a float. When the tank
is more than half full, the float is raised to a position that closes the valve and prevents the scavenge system from
operating. When the tank drops below half full, the float drops and lets the valve swing open allowing the scavenge
system to discharge fuel into the left main tank.
IF 28-18
757 General Familiarization (7-2005)
ATA 28 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AUXILIARY POWER UNIT (APU) FUEL FEED SYSTEM
Overview
The center fuel tank contains all APU fuel feed system components except the APU fuel supply line and shroud.
The APU fuel supply line and shroud lie between the center fuel tank and the APU. The DC Pump supplies the APU
with fuel upon loss of AC power or boost pump failure.
APU Fuel Feed System
The engine fuel feed system normally supplies fuel to operate the APU whenever AC power is available on the
airplane. The DC fuel pump takes over upon loss of AC power on the airplane or due to boost pump failure. The
APU fuel feed system supplies the APU with fuel.
The APU fuel feed system consists of a DC fuel pump, check valve, fuel shutoff valve, DC fuel pump pressure
switch, APU fuel shutoff pressure switch, two APU flame arrestors, an APU sump drain valve, and a fuel supply
line surrounded by a shroud. The DC fuel pump draws fuel from the left main fuel tank and only operates on AC
power loss or boost pump failure. The check valve allows the engine fuel feed system to supply the APU with fuel
and prevents a back flow of fuel to the engine feed manifold when the DC pump operates. The shutoff valve
controls fuel flow to the APU. The DC Pump pressure switch senses pump output pressure. The shroud prevents
any fuel vapor from the APU supply line from entering the airplane cabin.
The APU fuel shutoff pressure switch senses fuel pressure in the fuel feed manifold. The switch interrupts power
to the DC fuel pump if the pressure in the manifold decreases below 4 PSI. The Engine Indicating and Crew
Alerting System (EICAS) provides caution and warning messages for selected system failures.
757 General Familiarization (7-2005)
ATA 28 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 27-19
757 General Familiarization (7-2005)
ATA 28 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU FUEL FEED SYSTEM COMPONENTS
DC Fuel Pump
The DC fuel pump, a positive displacement type, consists of a pump and driving motor within a housing. The
pump housing, within the fuel tank, contains a check valve and a plunger valve. The housing outlet contains the
check valve. The housing inlet contains the plunger valve. The driving motor, 28-volts DC powered, includes a
thermal fuse, a check valve, a discharge valve, and a pressure switch. The pump and drive motor mount on the
outside of the fuel tank. The motor turns at 6600 RPM and the pump provides a continuous flow of 3.1 GPM at 18
PSI. Fuel provides cooling and lubrication, through the check valve in the drive motor.
IF 28-20
757 General Familiarization (7-2005)
ATA 28 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Shutoff Valve
The APU fuel shutoff valve is a motor actuated ball valve. The shutoff valve mounts to the rear spar of the center
wing fuel tank. The valve body (inside the fuel tank) attaches to the actuator (outside the fuel tank) so that defueling is not required for actuator removal. A thermal relief, ported to the tank side
de of the valve, cracks and
reseats between 75 and 45 PSIG. A slotted ball within the valve body controls fuel flow through the valve.
The square (preferred) actuator of the APU fuel shutoff valve has a 28-volts DC permanent magnet motor, an
electrical connector, OPEN - CLOSED limit switches, and a manual override handle. The manual override handle
also shows the position of the APU fuel shutoff valve.
The round (alternate) actuator consists of a 28-volts DC drive motor, an electrical connector, open-closed limit
switches, and a manual override handle which also functions as a position indicator.
IF 28-21
757 General Familiarization (7-2005)
ATA 28 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 28 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Fuel Line and Shroud
The APU fuel line connects from the DC fuel pump to the APU shutoff valve, and aft under the passenger
compartment floor to the APU. The APU fuel line carries fuel from the left main fuel tank to the APU.
A shroud surrounds the APU fuel line, between the dry side of the fuel tanks and the APU firewall. The shroud is
to prevent leakage of fuel vapor.
The APU fuel line drain mast is on the right lower side of the tail cone. The drain line to the drain mast connects
to the APU fuel line just forward of the APU firewall.
IF 28-22
757 General Familiarization (7-2005)
ATA 28 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Flame Arrestor
AIRPLANES WITH AN APU SUMP DRAIN VALVE AFT OF THE LEFT MAIN GEARWHEEL WELL; three APU
flame arrestors mount in the APU fuel line shroud drain tubing. Two flame arrestors are installed in the drain
tubing just above the APU fuel sump drain valve aft of the left main gear wheel well. The other flame arrestor
mounts in the drain tubing just above the center fuel tank.
AIRPLANES WITH AN APU SUMP DRAIN VALVE BETWEEN THE MAIN GEAR WHEEL WELLS; two APU flame
arrestors mount in the APU fuel line shroud drain tubing. One flame arrestor is installed in the drain tubing just
above the APU fuel sump drain valve between the main gear wheel wells. The other flame arrestor mounts in the
drain tubing just above the center fuel tank.
The flame arrestors prevent external flame from entering the airplane through the APU fuel sump drain valve. Each
flame arrestor acts as a heat sink to keep fuel vapor below the ignition point.
APU Sump Drain Valve
Either the APU sump drain valve is located just aft of the left main gear wheel well and is installed on the inside
of the wing-to-body fairing or the APU sump drain valve is located between the main wheel well doors and is
installed on the inside of the keel beam.
The sump drain valve is a manually operated drain valve. It consists of two spring loaded poppets, a primary
poppet and a secondary poppet. Pushing upward on the drain valve primary poppet opens the drain valve and
allows fuel to drain from the drain lines. The secondary poppet closes to seal the valve during replacement of the
primary poppet.
APU Fuel Check Valve
The APU fuel check valve, a one way flapper valve, allows the engine fuel feed system to supply the APU with fuel.
The valve prevents fuel flow to the engine feed manifold when the DC fuel pump operates.
Pressure Switches
Two identical pressure switches, a DC fuel pump pressure switch and an APU fuel pump shutoff pressure switch,
provide APU fuel feed system indication and shutoff control, respectively. Both pressure switches are cylindrical in
shape with a pressure port on one end and an electrical connector on the other end. The switches open at 4 PSIG.
757 General Familiarization (7-2005)
ATA 28 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-23
757 General Familiarization (7-2005)
ATA 28 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DEFUELING SYSTEM
Overview
The de-fueling system consists mainly of de-fueling valves. Other necessary components include the boost pumps,
the boost pump bypass valves, the fueling valves, and the cross feed valve. Three methods for de-fueling a fuel
tank include pressure de-fueling, suction de-fueling, and tank-to-tank transfer.
De-fueling
The de-fueling system consists of a de-fueling valve at the left and right ends of the center fuel tank connecting
the fuel feed manifold to the main fueling manifold. Fuel extraction takes place at the fueling station through
fueling adapters.
Pressure de-fueling and suction de-fueling provide the means for de-fueling the fuel tanks. A boost pump bypass
suction line connects each main fuel tank to the fuel feed manifold. Suction de-fueling draws fuel into the boost
pump bypass through the fuel feed manifold and main fueling manifold and out of the fueling adapters. Only the
main fuel tanks may be suction de-fueled; pressure de-fueling uses the boost pumps to pump the fuel out of the
tanks through the same fuel path.
Only on-ground de-fueling is possible. Airborne fuel jettisoning is not possible. Fueling station controls regulate
de-fueling.
Tank-To-Tank Fuel Transfer
De-fueling of a fuel tank by transferring the fuel to the other fuel tank requires use of de-fueling valves, fueling
valves, the cross feed valve, and the boost pumps.
Only on-ground tank-to-tank fuel transfer is possible. Fueling station controls and the cross feed valve switch light
on the pilot's overhead panel P5 control tank-to-tank fuel transfer.
757 General Familiarization (7-2005)
ATA 28 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-24
757 General Familiarization (7-2005)
ATA 28 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 28 - Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DEFUELING SYSTEM COMPONENTS
De-fuel Valve
The de-fueling valve is a motor actuated butterfly valve. The de-fueling valve mounts to the rear spar of the center
wing fuel tank. The valve body (within the fuel tank) connects to the actuator (outside the fuel tank) by an
adapter/shaft. Access to the valve body requires de-fueling.
The de-fueling valve square (preferred) actuator has a 28-volts DC permanent magnet motor, an electrical
connector, OPEN – CLOSED limit switches, and a manual override handle. The manual override handle also shows
the position of the de-fueling valve.
The de-fueling valve round (alternate) actuator has a 28-volts DC drive motor, an electrical connector, OPEN CLOSED limit switches, and a manual override handle. The manual override handle
le also shows the position of the
de-fueling valve.
Applying 28-volts DC power opens or closes the valve, depending upon the command
The de-fueling valve position determines whether the main fueling manifold is connected to the fuel feed manifold.
IF 28-25
757 General Familiarization (7-2005)
ATA 28 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL QUANTITY INDICATING SYSTEM (FQIS)
Overview
The fuel quantity indicating system (FQIS) provides fuel quantity measurement, calculation, and indication. The
system contains tank units, compensators, densitometers, indicators, and a fuel quantity processor unit (FQPU).
The FQPU does fuel quantity computation and fault isolation for the FQIS system. The FQPU also controls pressure
fueling for the airplane.
The fuel quantity indicating system is a microcomputer controlled system using the variable capacitance principle
for fuel quantity measurement. The capacitance measured by the tank units is roughly proportional to the amount
of fuel in the tank.
The FQIS contains the FQPU, a fuel quantity indicator, 3 load select indicators, a load select control, 3
compensators, 3 densitometers, 4 tank wiring harnesses, and 33 tank units. The FQPU is the principle component
of the FQIS. It is a digital, dual-channel, microprocessor base computer, located in the main equipment center on
rack E3-4.
The FQPU transmits excitation signals to and reads signals from the tank units and compensators; it also receives
densitometer data and computes fuel volume and weight. Each channel of the FQPU includes its own failure
monitoring, and built-in test equipment (BITE). Fault isolation is provided for all sensors and the fuel quantity
indicator. Faults both transient and permanent are stored in a non-volatile memory, for recall on demand by the
maintenance crew.
The FQPU unit provides the following outputs:
1. Signals to EICAS and the FUEL CONFIG light to warn of low fuel, fuel imbalance or fuel in center tank with
override pumps off, or system faults.
2. Digital fuel quantity (primary and total) displayed on the fuel quantity indicator on the pilots overhead
panel, P5.
3. Digital fuel load selected, displayed on the load select indicators on the fueling control panel, P28.
4. Volumetric and fuel quantity shutoff signals sent to the fueling valves for automatic shutoff.
The tank units and compensators are similar in construction and mounting provisions. Both units are capacitance
type sensors but the tank units sense the height of fuel in the tank and the compensators measure the dielectric
constant of the fuel.
The densitometer senses fuel density. Knowing the density and the volume of fuel enables the FQPU to calculate
the mass or weight of the fuel.
A FUEL QTY test switch checks that the flight compartment displays including EICAS messages are operable. This
switch only checks the displays and does not indicate the condition of the FQIS system.
757 General Familiarization (7-2005)
ATA 28 - Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-26
757 General Familiarization (7-2005)
ATA 28 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL QUANTITY INDICATING SYSTEM (FQIS) COMPONENTS
Tank Units
There are 12 tank units in each main tank, and 9 units in the center tank. All tank units are identical, except for
their length. This depends on the depth of the tank at that particular location.
The tank unit senses the fuel level in the fuel tank. It is a linear, capacitance type sensor. Tank characteristics are
accounted for by computer software. The concentric aluminum tubes of the tank unit act as plates of a capacitor,
and the dielectric is air and/or fuel depending on the amount of fuel in the tank. This air to fuel ratio produces a
capacitance value which is used to compute the volume of fuel.
IF 28-27
757 General Familiarization (7-2005)
ATA 28 - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Compensator
One compensator mounts in each fuel tank. All compensators are identical. The compensators are similar in
construction and mounting provisions to the tank units.
The compensator senses the dielectric constant of the fuel. It is a capacitance sensing device similar to the tank
unit. The compensator mounts at the bottom of the tank, completely covered by fuel until the tank is almost empty.
As a result, the compensator capacitance is affected only by changes in the dielectric constant of the fuel.
IF 28-28
757 General Familiarization (7-2005)
ATA 28 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Densitometer
One densitometer mounts in each fuel tank. The densitometer measures the fuel density.
The densitometer consists of two parts, the emitter and the electronics unit. All densitometers mount on the rear
spar with the emitter inside the fuel tank, and the electronics unit outside of the tank. Removal of the electronics
unit does not require de-fueling.
The densitometer electronics unit mounts on the studs of the emitter mounting plate, on the aft surface of the rear
spar. Removal takes place without de-fueling.
The two detector tubes are filled with xenon gas. When a high voltage of 1400-volts from the power supply board is
applied to the anode of the detector tubes, the interaction of the xenon gas, and the gama radiation from the
emitter, produces a variation in voltage at the anodes. The variation in voltage, or pulse, passes to the signal
processing electronics. The two pulse train output signals of the electronics unit go to the FQPU for calculation of
fuel density. Fuel density is proportional to the natural logarithmic value of the ratio of the two pulse output
signals.
IF 28-29
757 General Familiarization (7-2005)
ATA 28 - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Center Tank Densitometer Fuel Retention System
A fuel retention system is installed on the center tank densitometer
nsitometer and consists of a retention canister and a float
valve. The system allows the center tank densitometer to be covered with fuel before the center tank fuel level is
high enough to cover the densitometer. This allows the center tank densitometer to determine fuel density, while
refueling, at an earlier time than would be possible without the system.
IF 28-30
757 General Familiarization (7-2005)
ATA 28 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fuel Quantity Indicator
A fuel quantity indicator mounts on the pilot's overhead panel, P5. It provides digital liquid crystal displays of
individual tank, and total fuel quantities; and fuel temperature.
The fuel quantity indicator can be tested by placing the FUEL QTY test switch on P61 panel to FUEL QTY. The total
fuel readout should display 188.8. The individual tank readouts should display 88.8. The fuel temperature indicator
should display -185 ± 2. Returning the test switch to the neutral position causes the readouts to display the
amount of usable fuel.
FUEL QTY Test Switch
The FUEL QTY test switch is located on the right side panel, P61 in the flight compartment.
Placing the FUEL QTY test switch in the FUEL QTY position will cause the following:
1.
2.
3.
4.
5.
6.
7.
8.
The left, center, and right fuel quantity indicators on pilot’s overhead panel P5 will display 88.8.
The total fuel quantity indicator on P5 panel will display 188.8.
The fuel temperature indicator on P5 panel will display -185 ± 2.
The FUEL CONFIG light on P5 panel will illuminate.
The EICAS caution message LOW FUEL will appear.
The EICAS status message FUEL QTY IND will appear.
The EICAS status message FUEL QTY CHANNEL will appear.
The EICAS maintenance message FUEL QTY BITE will appear.
NOTE:
In order to view the EICAS maintenance message FUEL QTY BITE, 45 seconds must
elapse before again placing the FUEL QTY test switch in the FUEL QTY position. Then
the ECS/MSG switch on the right side panel P61 must be immediately pushed and
released.
9. The FQPU will perform a VERIFY test on the fuel quantity indicating system.
NOTE:
The FQPU will not perform a VERIFY test if the fueling station door is open or if the
VERIFY switch on the FQPU has just been pressed. The VERIFY test will be completed in
approximately 45 seconds.
The FUEL QTY test switch has a time out of approximately 8 seconds. After approximately 8 seconds, the fuel
quantity indicator and all other displays will revert back to their previous condition. This 8 second time out is to
prevent a broken FUEL QTY test switch from disabling the fuel quantity indicator and the other displays.
757 General Familiarization (7-2005)
ATA 28 - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
If prior to using the FUEL QTY test switch, a fault had caused the fuel quantity indicator to blank, then the fuel
quantity indicator may give a false reading until the VERIFY test is completed. If the fault still exists after the
VERIFY test, then the fuel quantity indicator will blank again.
Once the FUEL QTY test switch is used, 45 seconds must elapse before using the FUEL QTY test switch again.
Using the FUEL QTY test switch more than once in a 45 second time frame will introduce nuisance fault codes and
nuisance status codes into the FQPU.
IF 28-31
757 General Familiarization (7-2005)
ATA 28 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Load Select Indicator
One load select indicator located on the fuel station, P28 panel, monitors each tank.
There are two black-on-white LCDs for each indicator. The upper display
splay reads fuel quantity in each tank in LBS x
1000. The lower display shows the fuel quantity selected to be loaded into the tank in LBS x 1000. A 7-1/2 watt,
automatically controlled heater assures proper LCD operation in cold weather. Four permanently installed
incandescent lamps provide indicator back lighting. Power is applied to the indicators whenever the fueling station
door is opened, unless the power is supplied by the battery.
To test the load select indicators press the TEST IND switch on the fueling control panel, P28. The upper and
lower displays should both read 88.8.
IF 28-32
757 General Familiarization (7-2005)
ATA 28 - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Load Select Control
The load select control is located on the fueling control panel, P28.
The load select control consists of an explosion-proof case with a flame arrestor type breathing plug. The case
houses three thumbwheel type switches, three resistor boards, and an electrical connector.
The load select control permits the fueling crew to select the amount of fuel to be loaded into each individual tank.
This is accomplished by setting the thumbwheel switches then pressing the SET switch for the required tank. The
load select control then sends an analog signal to the FQPU representative of the fuel to be loaded.
IF 28-33
757 General Familiarization (7-2005)
ATA 28 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fuel Quantity Processor Unit (FQPU)
The fuel quantity processor unit (FQPU, also referred to as the "processor") is a digital, dual channel computer.
Memory provisions include; read only memory (ROM) for storing the programs, random access memory (RAM) for
scratchpad functions, and electrically alterable read only memory (EAROM) for storing faults. All input/output
ports on the FQPU are memory mapped.
The FQPU mounts on the E3-4 shelf in the main equipment center. Two hold-down hooks at the front of the FQPU
secure the unit on the rack. An ARINC 600-1 style 2 connector, located at the back of the FQPU, connects it with
the other components of the FQIS. On the front panel of the FQPU are controls for the built-in test, and a red dot
(light emitting diode) matrix, fault and status code display.
IF 28-34
757 General Familiarization (7-2005)
ATA 28 - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL MEASURING STICKS
Overview
The fuel measuring sticks allow manual measurement of fuel quantity in the fuel tanks. The fuel measuring sticks
should be used when the fuel quantity indicator in the flight compartment is not operating correctly.
Upon delivery of any new airplane, a fuel measuring stick document is provided along with the airplane. The fuel
measuring stick document provides instructions for the determination of fuel quantity by using the fuel measuring
sticks. The main body of this document is a set of conversion tables relating fuel measuring stick readings and
airplane attitude readings to obtain the fuel quantity.
NOTE:
Distribution and storage of the fuel measuring stick documents are determined by each
individual airline.
FUEL MEASURING STICKS COMPONENTS
Fuel Measuring Sticks
Each main tank has five fuel measuring sticks on the lower wing surface. All fuel measuring sticks located in the
main tank are mounted to a main tank access door. The center tank contains four fuel measuring sticks. Two
center tank fuel measuring sticks mount to center tank access doors. The remaining two fuel measuring sticks
mount to the structure of the center tank lower surface.
The fuel measuring stick assembly contains a housing, magnet and float, and a stick sub assembly. The housing,
with magnet and float, is sealed and mounted to the inside of the lower wing surface. A float stop prevents the
float from coming off the housing. The stick sub assembly fits into the housing from outside the airplane. Rotating
the latch in a counterclockwise direction releases the stick sub assembly from the housing, allowing it to fall.
Depending on the fuel level, the magnet in the float attracts the steel armature at the top of the stick sub
assembly.
The calibrated marks on the stick sub assembly shows fuel quantity in pounds.
757 General Familiarization (7-2005)
ATA 28 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-35
757 General Familiarization (7-2005)
ATA 28 - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 28-36
757 General Familiarization (7-2005)
ATA 28 - Page 59
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 29
HYDRAULIC POWER SYSTEM ............................................................................................................. 4
Overview............................................................................................................................................. 4
Main Hydraulic Systems ..................................................................................................................... 7
Auxiliary Hydraulic Systems .............................................................................................................. 7
Ground Servicing System ................................................................................................................... 8
Indicating Systems .............................................................................................................................. 8
HYDRAULIC POWER SYSTEM COMPONENTS ............................................................................... 11
Engine Driven Pump (EDP) ............................................................................................................. 11
Alternating Current Motor Pump (ACMP) ....................................................................................... 12
Hydraulic Reservoir .......................................................................................................................... 14
Reservoir Pressurization Module ...................................................................................................... 15
Return Filter Modules ....................................................................................................................... 16
EDP Pressure/Case Drain Filter Module .......................................................................................... 17
ACMP Pressure/Case Drain Filter Module ...................................................................................... 18
Left and Right System Modules ....................................................................................................... 18
Center System Module...................................................................................................................... 18
EDP Supply Shutoff Valve ............................................................................................................... 19
Isolated ACMP Shutoff Valves ........................................................................................................ 20
Heat Exchanger ................................................................................................................................. 21
Ground Power Connections .............................................................................................................. 22
Hydraulic Control Panel ................................................................................................................... 23
Ground Servicing Station ................................................................................................................. 24
HYDRAULIC SERVICING..................................................................................................................... 25
Fill the Hydraulic Reservoir - Manual Procedure............................................................................. 25
Fill the Hydraulic Reservoir - Pressure Procedure ........................................................................... 29
RAM AIR TURBINE (RAT) SYSTEM................................................................................................... 32
Overview........................................................................................................................................... 32
RAM AIR TURBINE (RAT) SYSTEM COMPONENTS....................................................................... 33
RAT Strut.......................................................................................................................................... 33
Hydraulic Pump ................................................................................................................................ 34
Turbine .............................................................................................................................................. 35
Deployment Actuator........................................................................................................................ 36
Actuator Control Valve..................................................................................................................... 37
Safety Valve...................................................................................................................................... 38
Checkout Module.............................................................................................................................. 39
Tachometer ....................................................................................................................................... 40
RAT Compartment Door .................................................................................................................. 41
Door Actuation Link ......................................................................................................................... 42
Extend the RAT ................................................................................................................................ 43
Retract the RAT ................................................................................................................................ 44
757 General Familiarization (7-2005)
ATA 29 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC POWER TRANSFER UNIT (PTU) SYSTEM ................................................................ 47
Overview........................................................................................................................................... 47
Hydraulic Power Transfer Unit (PTU) System Components ................................................................... 48
Power Transfer Unit (PTU) .............................................................................................................. 48
PTU Pressure Filter Module ............................................................................................................. 49
PTU Case Drain Filter Module ......................................................................................................... 50
PTU Control Valve ........................................................................................................................... 51
PTU Flow Limiter Valve (Not on all aircraft) .................................................................................. 52
PTU Control Pressure Switch (Not on all aircraft) ........................................................................... 53
Left EDP Pressure Switch (Not on all aircraft) ................................................................................ 54
HYDRAULIC PRESSURE INDICATING SYSTEM ............................................................................. 55
Overview........................................................................................................................................... 55
HYDRAULIC PRESSURE INDICATING SYSTEM COMPONENTS................................................. 56
EICAS Display ................................................................................................................................. 56
Low Pressure Lights ......................................................................................................................... 57
RAT Pressure Light .......................................................................................................................... 58
Pressure Transmitters........................................................................................................................ 59
HYDRAULIC FLUID TEMPERATURE INDICATING SYSTEM ....................................................... 60
Overview........................................................................................................................................... 60
HYDRAULIC FLUID TEMPERATURE INDICATING SYSTEM COMPONENTS........................... 61
EICAS display .................................................................................................................................. 61
Overheat Lights................................................................................................................................. 62
Overheat Light Temperature Switches ............................................................................................. 63
Temperature Transmitters ................................................................................................................. 64
HYDRAULIC FLUID QUANTITY INDICATING SYSTEM ............................................................... 65
Overview........................................................................................................................................... 65
HYDRAULIC FLUID QUANTITY INDICATING SYSTEM COMPONENTS ................................... 66
EICAS Display ................................................................................................................................. 66
Quantity Warning Lights .................................................................................................................. 67
Remote Quantity Indicator................................................................................................................ 68
Sight Glasses ..................................................................................................................................... 69
Quantity Transmitter ......................................................................................................................... 69
Quantity Monitor Unit ...................................................................................................................... 70
757 General Familiarization (7-2005)
ATA 29 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 29 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC POWER SYSTEM
Overview
Three separate hydraulic systems provide fluid at 3000 PSI to operate the airplane systems. The hydraulic systems
are identified as left, right, and center.
High pressure lines and critical return lines are made from titanium. Aluminum is used for non-critical return lines
and steel tubing is used in fire zones. Hydraulic tubing is color coded by system. The left system is coded red,
right is green, and center is blue.
Two auxiliary hydraulic systems provide reserve power. These are the ram air turbine (RAT) and the power
transfer unit (PTU).
A ground servicing system fills all hydraulic reservoirs from one location.
The indicating systems inform the pilots of the operating conditions of each hydraulic system.
757 General Familiarization (7-2005)
ATA 29 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 29-1
757 General Familiarization (7-2005)
ATA 29 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 29 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Hydraulic Systems
The center system components are located in the aft left wing/body fairing. The pressure sources for the center
system are two alternating current motor pumps (ACMP). The ACMPs run constantly when the ELEC pump select
switches are in the ON position.
The left and right system components are located on each engine, engine strut, and main wheel well forward
fairing. The pressure sources for each of the left and right systems consist of one engine driven pump (EDP) and
one alternating current motor pump (ACMP). The EDP runs all the time that the engine is running. When the ELEC
pump switch is ON, the ACMP runs all the time. These systems provide fluid under pressure of 3000 PSI to
portions of the flight control, auto flight, landing gear, and thrust reverser systems.
Each hydraulic system has a fluid reservoir which is pressurized by air from the pneumatic system. Filter modules
clean the fluid after being pressurized or after passing through the pump case drain and after returning from user
systems. Heat exchangers in the fuel tanks cool the pump case drain fluid before it returns to the reservoir.
Ground power connections are provided in each engine strut and in the aft left wing/body fairing for attachment of
an external hydraulic pressure source.
Auxiliary Hydraulic Systems
The ram air turbine (RAT) is stowed in the aft right wing/body fairing. The RAT consists of a hinged strut with a
hydraulic pump powered by a ram air driven turbine. The RAT deploys automatically if power is lost in both
engines. The RAT can be deployed manually with the RAT deployment switch in the flight compartment. The RAT
supplies pressure to the center system for operation of the flight controls.
The power transfer unit (PTU) in the left wheel well operates if power is lost in the left engine or if left EDP
output pressure is low. The PTU can also be switched on manually. The PTU consists of a hydraulic motor
connected to a hydraulic pump. The motor is driven by right hydraulic system pressure and the pump supplies
pressure to the left system. There is no fluid interchange between the right and left hydraulic systems. Left system
pressure supplied by the PTU operates the following systems:
1.
2.
3.
4.
5.
Nose gear steering
Flaps and slats
Landing gear
ON AIRPLANES WITH HYDRAULIC MOTOR-DRIVEN GENERATOR; hydraulic motor-driven generator
ON FREIGHTER AIRPLANES; main deck cargo door
757 General Familiarization (7-2005)
ATA 29 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Ground Servicing System
A central ground servicing station is located in the aft left wing/body fairing. Hydraulic fluid is added to the
reservoirs of all three systems from this station. A fill valve selects which reservoir is to receive fluid. A remote
quantity indicator shows fluid level in the reservoir selected by the fill valve. Fluid can be added under pressure
from a ground service cart or with the manual fill pump installed at the servicing station.
Indicating Systems
The indicating system consists of warning lights on the hydraulic control panel and messages on the engine
indicating and crew alerting system (EICAS). The indicating systems monitor fluid pressure, temperature, and
quantity and reservoir air pressure.
757 General Familiarization (7-2005)
ATA 29 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 29-2
757 General Familiarization (7-2005)
ATA 29 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 29 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC POWER SYSTEM COMPONENTS
Engine Driven Pump (EDP)
The engine driven pumps (EDP) are one of the pressure sources for the left and right systems. The EDP is driven
by the engine gearbox and is located on the right side of the engine. Each EDP supplies 37 GPM
M at 3700 RPM and
2850 PSI. A valve in the pump varies the amount of fluid delivered to hold system pressure constant until system
demand exceeds rated flow of pump. When system pressure reaches 3100 PSI the EDP delivery flow is zero for all
pump speeds.
The EDP moves fluid through the pump case to lubricate and cool the pump. This fluid flows out the case drain
port to the system heat exchanger and returns to the reservoir.
The output of the EDP can be shut off by a depressurization valve on the pump. This valve closes the pump outlet
port when the engine pump switch on the hydraulic control panel is moved to DEPRESS. The valve also closes
when the engine fire switch on the aft electronic control panel P8 is pulled out to FIRE position.
IF 29-3
757 General Familiarization (7-2005)
ATA 29 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Alternating Current Motor Pump (ACMP)
The left and right systems each have one ACMP and the center system has two ACMPs. The left and right system
ACMPs are in the left and right main wheel well forward fairings, respectively. The two center system ACMPs are
in the aft left wing/body fairing. Each pump is attached to vibration absorbing mounts.
The ACMP supplies 13.5 GPM at 1500 PSI. At pressure above 1500 PSI the pump output decreases to 7.0 GPM at
2850 PSI. When pressure reaches 3025 PSI the pump delivery flow is zero.
The pump is driven by a 115/200-volts 3-phase 400 Hz motor. The motor is controlled by the ELEC pump switches
on the hydraulic control panel. The right system ACMP will also operate when the RESERVE BRAKE switch on main
instrument panel P1 is on. The ACMP moves fluid through the pump case to lubricate and cool the pump. This
fluid flows out the case drain port to the system heat exchanger and returns to the reservoir.
The 28-volt dc control circuit for each ACMP contains a time delay. The duration of time delay for each pump is
different and ranges from 2 to 5 seconds for the various pumps. The ACMP time delays prevent a large surge on
the electrical power system as the ACMPs restart if a temporary interruption
ption of dc power occurs.
IF 29-4
757 General Familiarization (7-2005)
ATA 29 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 29-5
757 General Familiarization (7-2005)
ATA 29 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Hydraulic Reservoir
The hydraulic reservoir for each system supplies hydraulic fluid for the flow demands of the user systems. The
reservoirs for the left and right systems are in the left and right main wheel wells, respectively. The center system
reservoir is in the aft left wing/body fairing.
The three reservoirs are round tanks of the same design but of different size. Each reservoir holds 6.6 gallons (25
liters, 5.5 imperial gallons) for left and right systems and 3.5 gallons (13 liters, 2.9 imperial gallons) for the center
system. The reservoirs are pressurized to 45-50 PSI by the pneumatic system to assure a supply of fluid to the
pumps.
Each reservoir has a quantity transmitter and a sight glass for checking fluid level. Pressure relief,
depressurization, drain, and sampling valves and a fluid trap are on each reservoir.
IF 29-6
757 General Familiarization (7-2005)
ATA 29 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Reservoir Pressurization Module
This module filters and distributes air from the pneumatic system to the three reservoirs. The module is on the
keel beam forward of the main wheel wells. The module includes a filter, two check valves and a manual bleed
valve.
IF 29-7
757 General Familiarization (7-2005)
ATA 29 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Return Filter Modules
The return module in each system filters all fluid returning to the reservoir. The left and right system return
modules are in the left and right wheel well forward fairings, respectively. The center system return module is in
the aft left wing/body fairing.
Each module includes a throw-away filter. A differential pressure indicator shows when the filter is blocked. A bypass valve routes fluid around a blocked filter. A spring-loaded valve in the module prevents fluid draining when
the filter element is removed. The module has two check valves to prevent backflow through the filter.
The center system module also includes a ground power return
turn connector and the system pressure relief valve.
IF 29-8
757 General Familiarization (7-2005)
ATA 29 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EDP Pressure/Case Drain Filter Module
The EDP modules filter both pressure and case drain fluid from the EDPs. The EDP modules are in each engine
strut.
Each module includes two throw-away filters. Differential pressure indicators show when each filter is blocked. A
spring-loaded valve at each filter prevents fluid draining when the filter element is removed.
Each module also contains a pump pressure switch, a system pressure switch, a pump temperature
mperature switch, and
two check valves. On some airplanes the module also contains a system pressure relief valve. On others, the
pressure relief valve is located in the left and right wheel well forward fairing.
IF 29-9
757 General Familiarization (7-2005)
ATA 29 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ACMP Pressure/Case Drain Filter Module
The ACMP modules filter both pressure and case drain fluid from the ACMPs. The left and right system return
modules are in the left and right wheel well forward fairings, respectively. The center system module is in the aft
left wing/body fairing.
Left and Right System Modules
Each module includes two throw-away filters. Differential pressure indicators show when each filter is blocked. A
spring-loaded valve at each filter prevents fluid draining when the filter element is removed. Each module also has
a pump pressure
re switch and two check valves.
Center System Module
The center system filter module is a dual unit which filters fluid from two ACMPs. This module is the same as two
left or right system modules installed in one housing. The center system module also has a system pressure
switch.
IF 29-10
757 General Familiarization (7-2005)
ATA 29 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EDP Supply Shutoff Valve
The supply shutoff valve controls fluid supply to the EDP and is normally in the open position. The valve is
operated by a 28-volts DC motor which is controlled by the engine fire switch. A position indicator on the valve
provides a valve CLOSED or OPEN indication. The valve is installed in the EDP supply line in the engine strut.
IF 29-11
757 General Familiarization (7-2005)
ATA 29 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Isolated ACMP Shutoff Valves
Two valves provide a reserve hydraulic system to power normal brakes in the event of loss of the right and left
hydraulic systems. These valves consist of a supply and a pressure shutoff valve. Each valve is operated by a 28volts DC motor.
The right system ACMP normally draws fluid from the reservoir through a standpipe. When the shutoff valves are
activated, the supply shutoff valve provides fluid to the right
ght ACMP from the bottom of the reservoir. The pressure
shutoff valve isolates the output of the right ACMP to power only the normal brakes.
The RESERVE BRAKE switch on the captain's main instrument panel P1 operates both valves and turns on the right
ACMP.
IF 29-12
757 General Familiarization (7-2005)
ATA 29 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Heat Exchanger
A heat exchanger in each system cools pump case drain fluid as the fluid returns to the reservoir. The three
identical heat exchangers are in the wing fuel tanks and use fuel as a coolant. The left and right system heat
exchangers are at the inboard end of the left and right main fuel tanks, respectively. The center system heat
exchanger is adjacent to the left system exchanger. A minimum of 600 gallons (4020 pounds/1827 kilograms) of
fuel is required in each main tank to provide hydraulic fluid cooling.
IF 29-13
757 General Familiarization (7-2005)
ATA 29 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Ground Power Connections
Each hydraulic system can be pressurized from a hydraulic ground cart connected to the ground power
connections. The left and right system connections are in the left and right engine struts, respectively. The center
system connections are in the aft left wing/body fairing. At each location, the ground power connections consist of
self-sealing quick-disconnects for the pressure and return lines.
IF 29-14
757 General Familiarization (7-2005)
ATA 29 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Hydraulic Control Panel
The control panel is on the pilot's overhead panel, P5 and contains the pump select switches and system warning
lights.
The pump select switch for each EDP and ACMP is a push-to-activate type and contains an ON light and pump low
pressure light. The switch has a mechanical ON/DEPRESS indicator for the EDPs and ON/OFF for the ACMPs.
The panel has warning lights for low pressure and low fluid quantity in each system. Warning lights are provided
for low pressure and overheat in each pump.
IF 29-15
757 General Familiarization (7-2005)
ATA 29 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Ground Servicing Station
The ground servicing system fills all hydraulic reservoirs from one location. The system components are in the aft
left wing/body fairing. The system has a hand pump with suction line and a connection for pressure fill from a
ground cart. There is also a remote quantity indicator, a fill selector valve, and a fill filter module.
IF 29-16
757 General Familiarization (7-2005)
ATA 29 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC SERVICING
NOTE:
To get the correct fluid level in the reservoirs, you must fill the reservoirs in these
conditions:
1. All landing gear in the down position
2. All landing gear doors closed
3. The steering and flight controls in the neutral position
When you fill the reservoir in the right system, the pressure gage for the brake accumulator must show at least
2500 PSI.
Fill the Hydraulic Reservoir - Manual Procedure
Access Wing to Body - Aft Lower Half (Left) Access Panel 197KL Central Hydraulic Service Center
Supply electrical power.
Open the access door, 197KL, for the central hydraulic service center.
Do these steps before you fill the reservoir in the right hydraulic system:
1. Make sure the pressure gage on the brake accumulator, in the aft left wing-to-body fairing, shows a
minimum of 2500 PSI.
CAUTION:
YOU MUST HAVE FLUID IN THE RIGHT SYSTEM RESERVOIR, BEFORE YOU OPERATE THE
HYDRAULIC PUMPS OR YOU MUST PRESSURIZE THE RIGHT SYSTEM WITH A
HYDRAULIC SERVICE CART. IF YOU OPERATE THE HYDRAULIC PUMPS WITHOUT
SUFFICIENT FLUID IN THE RESERVOIR, YOU CAN CAUSE DAMAGE TO THE PUMPS.
2. If the pressure gage does not show a minimum of 2500 PSI, pressurize the right hydraulic system.
3. Remove hydraulic power, after the right hydraulic system pressurizes the brake accumulator to a minimum
of 2500 PSI.
757 General Familiarization (7-2005)
ATA 29 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill The Hydraulic Reservoir (Continued):
WARNING:
USE THE PROCEDURE IN AMM 32-00-15 TO REMOVE THE DOOR LOCKS. THE DOORS
OPEN AND CLOSE QUICKLY AND CAN CAUSE INJURY TO PERSONS OR DAMAGE TO
EQUIPMENT.
4. For the left or right system, remove the door locks from the landing gear doors and close the doors.
CAUTION:
USE CLEAN HYDRAULIC FLUID, BMS 3-11, AND CLEAN EQUIPMENT TO FILL THE
HYDRAULIC SYSTEM RESERVOIRS. IF YOU DO NOT USE CLEAN HYDRAULIC FLUID AND
EQUIPMENT, CONTAMINATION OF THE HYDRAULIC SYSTEM CAN OCCUR.
NOTE:
All currently qualified BMS 3-11, Type IV hydraulic fluids are interchangeable and
intermixable in any proportion.
Remove the suction hose from the pump handle.
Clean the suction hose with a rag.
Put the end of the hose in a container of hydraulic fluid.
Remove the pump handle from the brackets.
Put the pump handle in the socket of the manual fill pump.
WARNING:
757 General Familiarization (7-2005)
BE CAREFUL TO NOT FILL THE RESERVOIR TO MORE THAN THE NECESSARY LEVEL. IF
YOU PUT TOO MUCH FLUID IN THE RESERVOIRS, THE FLUID CAN GO INTO THE DUCTS
OF THE PNEUMATIC SYSTEM AND THE AIR CONDITIONING PACKS. THIS CAN CAUSE
SMOKE AND DANGEROUS FUMES TO GO INTO THE FLIGHT COMPARTMENT AND THE
PASSENGER CABIN. IF CONTAMINATION OF THE PNEUMATIC SYSTEM OCCURS AGAIN
AND AGAIN, IT CAN CAUSE DAMAGE TO THE TITANIUM DUCTS.
ATA 29 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill The Hydraulic Reservoir (Continued):
Do these steps for each hydraulic reservoir you will fill:
1. Put the reservoir fill valve to the position (L, R, or C) for the hydraulic system in which you will fill the
reservoir.
2. Operate the manual fill pump to fill the reservoir.
3. Monitor the reservoir fill indicator while you fill the reservoir.
4. Stop when the reservoir fill indicator shows at the top of the area that has a color, as follows:
a) Red for the left system (6.6 gal./25 liters)
b) Blue for the center system (3.5 gal./13 liters)
c) Green for the right system (6.6 gal. /25 liters)
5. Make sure the six EICAS circuit breakers on the overhead panel, P11, are closed.
6. Push the ELEC/HYD switch on the EICAS MAINT panel, on the right side panel, P61.
757 General Familiarization (7-2005)
ATA 29 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill The Hydraulic Reservoir (Continued):
7. Make sure the (L, R, C) HYD QTY indication on the EICAS display shows as follows:
a) 1.00 (± 0.08 for the left or right system)
b) 1.00 (± 0.14 for the center system)
Put the reservoir fill valve in the CLOSED position.
Remove the pump handle from the socket of the manual fill pump.
Put the pump handle in the brackets.
Remove the suction hose from the container of hydraulic fluid.
Drain the hydraulic fluid from the suction hose.
Remove the hydraulic fluid from the suction hose with a rag.
Put the suction hose into the pump handle.
Remove electrical power if it is not necessary.
(19) Close the access panel, 197KL, for the central hydraulic service center.
757 General Familiarization (7-2005)
ATA 29 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill the Hydraulic Reservoir - Pressure Procedure
Access Wing to Body - Aft Lower Half (Left) Access Panel 197KL Central Hydraulic Service Center.
Supply electrical power.
Open the access door, 197KL, for the central hydraulic service center.
Do these steps before you fill the reservoir in the right hydraulic system:
1. Make sure the pressure gage on the brake accumulator, in the aft left wing-to-body fairing, shows a
minimum of 2500 PSI.
CAUTION:
YOU MUST HAVE FLUID IN THE RIGHT SYSTEM RESERVOIR, BEFORE YOU OPERATE THE
HYDRAULIC PUMPS OR YOU MUST PRESSURIZE THE RIGHT SYSTEM WITH A
HYDRAULIC SERVICE CART. IF YOU OPERATE THE HYDRAULIC PUMPS WITHOUT
SUFFICIENT FLUID IN THE RESERVOIR, YOU CAN CAUSE DAMAGE TO THE PUMPS.
2. If the pressure gage does not show a minimum of 2500 PSI, pressurize the right hydraulic system.
3. Remove hydraulic power, after the right hydraulic system pressurizes the brake accumulator to a minimum
of 2500 PSI.
WARNING:
USE THE PROCEDURE IN AMM 32-00-15 TO REMOVE THE DOOR LOCKS. THE DOORS
OPEN AND CLOSE QUICKLY AND CAN CAUSE INJURY TO PERSONS OR DAMAGE TO
EQUIPMENT.
4. For the left or right system, remove the door locks from the landing gear doors and close the doors.
CAUTION:
USE CLEAN HYDRAULIC FLUID, BMS 3-11, AND CLEAN EQUIPMENT TO FILL THE
HYDRAULIC SYSTEM RESERVOIRS. IF YOU DO NOT USE CLEAN HYDRAULIC FLUID AND
EQUIPMENT, CONTAMINATION OF THE HYDRAULIC SYSTEM CAN OCCUR.
NOTE:
All currently qualified BMS 3-11, Type IV hydraulic fluids are interchangeable and
intermixable in any proportion.
757 General Familiarization (7-2005)
ATA 29 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill The Hydraulic Reservoir (Continued):
Connect the ground service cart to the pressure fill connection.
WARNING:
BE CAREFUL TO NOT FILL THE RESERVOIR TO MORE THAN THE NECESSARY LEVEL. IF
YOU PUT TOO MUCH FLUID IN THE RESERVOIRS, THE FLUID CAN GO INTO THE DUCTS
OF THE PNEUMATIC SYSTEM AND THE AIR CONDITIONING PACKS. THIS CAN CAUSE
SMOKE AND DANGEROUS FUMES TO GO INTO THE FLIGHT COMPARTMENT AND THE
PASSENGER CABIN. IF CONTAMINATION OF THE PNEUMATIC SYSTEM OCCURS AGAIN
AND AGAIN, IT CAN CAUSE DAMAGE TO THE TITANIUM DUCTS.
Do these steps for each hydraulic reservoir you will fill:
1. Put the reservoir fill valve to the position (L, R, or C) for the hydraulic system in which you will fill the
reservoir.
CAUTION:
USE A MAXIMUM OF 150 PSI WHEN YOU FILL THE HYDRAULIC RESERVOIR. TOO MUCH
PRESSURE CAN CAUSE DAMAGE TO THE HYDRAULIC RESERVOIR.
2. Operate the hydraulic service cart.
3. Monitor the reservoir fill indicator while you fill the reservoir.
4. Stop when the reservoir fill indicator shows at the top of the area that has a color, as follows:
a) Red for the left system (6.6 gal./25 liters)
b) Blue for the center system (3.5 gal./13 liters)
c) Green for the right system (6.6 gal./25 liters).
5. Make sure the six EICAS circuit breakers on the overhead panel, P11, are closed.
6. Push the ELEC/HYD switch on the EICAS MAINT panel, on the right side panel, P61.
7. Make sure the (L, R, C) HYD QTY indication on the EICAS display shows as follows:
a) 1.00 (±| 0.08 for the left or right system)
b) 1.00 (± 0.14 for the center system)
757 General Familiarization (7-2005)
ATA 29 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill The Hydraulic Reservoir (Continued):
Put the reservoir fill valve in the CLOSED position.
Stop the hydraulic ground service cart.
Disconnect the hydraulic ground service cart from the pressure fill connection.
Install a cap on the pressure fill connection.
Remove electrical power if it is not necessary.
Close the access door, 197KL, for the central hydraulic service center.
757 General Familiarization (7-2005)
ATA 29 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RAM AIR TURBINE (RAT) SYSTEM
Overview
The ram air turbine (RAT) provides reserve hydraulic power to the center system for operation of the flight
controls. The RAT is stowed in the aft right wing/body fairing. When deployed, the RAT swings out into the air
stream and remains extended until it is retracted on the ground. The RAT deploys if power is lost in both engines.
The RAT can be deployed manually with the RAT deployment switch in the flight compartment.
When the airplane is on the ground, the RAT can be deployed and retracted with the RAT ground manual switch in
the right wheel well. Prior to retracting the RAT, the propeller blades must be moved to align index marks on the
hub and strut.
The RAT consists of a turbine and a hydraulic pump mounted on a hinged strut housing. The turbine consists of a
hub with two propeller blades which are driven by the air stream. The RAT is deployed and retracted by the
deployment actuator. The actuator retracts with hydraulic pressure and deploys with spring pressure.
757 General Familiarization (7-2005)
ATA 29 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RAM AIR TURBINE (RAT) SYSTEM COMPONENTS
RAT Strut
The RAT strut supports the hydraulic pump and turbine. The strut is pivoted at the upper end where it mounts to
the airplane. Swivel valves at the top of the strut transfer fluid from the strut to the center hydraulic system.
Internal passages in the strut carry fluid between the swivel valves and hydraulic pump.
A spring-loaded blade-lock plunger is mounted at the lower end of the strut. The plunger locks the turbine blades
in the correct position for stowing the RAT. The plunger also prevents rotation of thee blades until clear of the
airplane during deployment. The plunger is actuated by a cable attached to a bracket on the outboard swivel valve.
A blade index switch actuated by the plunger stops retraction of the RAT if the plunger does not engage the turbine
hub.
IF 29-17
757 General Familiarization (7-2005)
ATA 29 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Hydraulic Pump
The hydraulic pump supplies 11.3 GPM at 2140 PSI. The pump flow is controlled by the output pressure. At 3025
PSI the pump delivery flow is zero.
The pump can be used as a hydraulic motor to drive the turbine for a test of the governor. To back-drive the
turbine, a valve in the checkout module reverses the fluid flow from the center hydraulic system. The system
pressure is then used as a power source to drive the pump as a motor. In this mode, the pump drives the turbine
in the normal direction. Prior to back-driving the turbine, a safety screen is installed which completely encloses the
turbine blades for personnel protection. The safety screen must be removed before the RAT is retracted.
IF 29-18
757 General Familiarization (7-2005)
ATA 29 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Turbine
The turbine consists of a hub with two propeller blades.. The hub is connected to the hydraulic pump by a
driveshaft. The blades are attached to a mechanical governor in the hub. The governor varies the blade pitch angle
to control the turbine speed.
IF 29-19
757 General Familiarization (7-2005)
ATA 29 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Deployment Actuator
The deployment actuator is a hydraulic cylinder attached to airplane structure and to the RAT strut. The actuator
moves the strut between the deployed and retracted positions. Right hydraulic system pressure drives the actuator
to the retracted position where it is held by an internal mechanical lock. The actuator is driven to the deployed
position by springs which compress during retraction. The actuator is unlocked by either of two solenoids, one for
manual deployment and one for automatic deployment. Hydraulic fluid from the actuator flows into the right
system return line as the RAT deploys.
IF 29-20
757 General Familiarization (7-2005)
ATA 29 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Actuator Control Valve
The control valve connects the deployment actuator to either the right system return line or to the RAT safety
valve. The valve is on the fuselage skin inside the aft right wing/body fairing. The two-position valve is operated
by a 28-volts DC motor which is controlled by the RAT ground manual switch. The actuator pressure port is
normally connected to the right system return line. For deployment or retraction with the RAT ground manual
switch, the actuator pressure port is connected to the RAT safety valve.
IF 29-21
757 General Familiarization (7-2005)
ATA 29 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Safety Valve
The safety valve is on the aft wall of the right main wheel well. The valve is a three-position valve operated by two
28-volts DC solenoids. The valve solenoids are controlled by the RAT ground manual switch. In the retraction
position, the valve connects right system pressure to the deployment actuator pressure port. In the deployment
position, the valve connects the actuator pressure port to the right system return line. In the off position, the valve
blocks flow from the actuator pressure port to stop actuator movement.
IF 29-22
757 General Familiarization (7-2005)
ATA 29 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Checkout Module
The checkout module is on the aft wall of the right main wheel. The module contains a back-drive valve, filters,
pressure switch, check valves, and a volumetric fuse cartridge valve. The module is used to check the operation of
the RAT with the airplane on the ground.
The back-drive valve is a manual valve which reverses the RAT pump connections to the center hydraulic system.
In the back-drive position, center system pressure is routed to the supply port of the RAT pump. This causes the
pump to act as a motor driving the turbine in the normal direction.
The module filters both pressure and case drain fluid from the RAT pump. The filters are a non-cleanable type.
A pressure switch senses pressure output of the RAT pump and turns on the RAT pressure light in the flight
compartment.
Check valves prevent the RAT system from being pressurized by the center system while the RAT is not in use. An
orifice in the back-drive valve permits a small flow of fluid to warm the strut and pump. The fluid returns to the
center system through the pump case drain line.
The volumetric fuse cartridge valve reduces the pressure load on the RAT pump until the turbine reaches governed
speed.
IF 29-23
757 General Familiarization (7-2005)
ATA 29 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Tachometer
The tachometer provides a visual indication of the operating status of the RAT during ground test. The tachometer
consists of an indicator and a speed sensor. The tachometer indicator is inside the fairing panel at the lower edge
of the right wheel well aft bulkhead.
The tachometer indicator has two indicator lights and a lamp test switch. A green light indicates turbine speed is
at the normal governed speed. A red light indicates turbine over-speed. If no light is on, the turbine speed is below
the normal governed speed. The lamp test switch is used to check that the indicator lights are not burned out.
When the switch is pressed, with the RAT operating, both indicator lights should illuminate.
Power for the tachometer indicator and the source of turbine speed sensing is provided by a speed sensor located
in the lower end of the RAT strut between the hub and pump. The speed sensor is a permanent magnet generator
that produces an electrical signal which is used to power the indicator lights.
IF 29-24
757 General Familiarization (7-2005)
ATA 29 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RAT Compartment Door
The compartment door is hinged on the outboard side and opens as the RAT is deployed. The door and opening in
the fairing have seals to isolate the compartment when the RAT is stowed.
IF 29-25
757 General Familiarization (7-2005)
ATA 29 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Actuation Link
The actuation link joins the RAT strut to the compartment door. The link has a bearing at each end. The link forces
the door open or closed as the RAT deploys or retracts.
IF 29-26
757 General Familiarization (7-2005)
ATA 29 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Extend the RAT
Access MLG Wheel Well Wing to Body - Aft Lower Half (Right)
WARNING:
USE THE PROCEDURE IN AMM 32-00-15 TO INSTALL THE DOOR LOCKS. THE DOORS
OPEN AND CLOSE QUICKLY AND CAN CAUSE INJURY TO PERSONS OR DAMAGE TO
EQUIPMENT.
Open the doors for the landing gear and install the door locks.
Supply electrical power.
Make sure these circuit breakers on the main power distribution panel, P6, are closed:
•
6F1, RAT MAN or RAT MAN PWR
Make sure these circuit breakers on the overhead panel, P11, are closed:
1. 11D26, HYDRAULIC RAT CONT or HYDRAULIC RAT AUTO CONT
2. 11D27, HYDRAULIC RAT AUTO or HYDRAULIC RAT AUTO PWR
Remove the access panel for the ground manual switch from the fairing on the aft bulkhead of the right wheel well.
WARNING:
KEEP PERSONS AND EQUIPMENT AWAY FROM THE PATH OF THE RAT AND THE RAT
COMPARTMENT DOOR. THE RAT AND THE RAT COMPARTMENT DOOR MOVE QUICKLY
AND CAN CAUSE INJURY TO PERSONS OR DAMAGE TO EQUIPMENT.
Hold the ground manual switch in the down (deploy) position to extend the RAT.
757 General Familiarization (7-2005)
ATA 29 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Extend the RAT (Continued):
Open these circuit breakers on the P11 panel, and install the RAT circuit breaker lock set:
1. 11D26, HYDRAULIC RAT CONT or HYDRAULIC RAT AUTO CONT
2. 11D27, HYDRAULIC RAT AUTO or HYDRAULIC RAT AUTO PWR
Open these circuit breakers on the P6 panel, and install the RAT circuit breaker lockset:
1. 6F1, RAT MAN or RAT MAN PWR
2. 6F2, RAT MAN CONT
Remove the pressure from the right hydraulic system.
Install the protective covers on the RAT turbine blades and the turbine hub.
Retract the RAT
NOTE:
Use two persons to do the RAT retraction procedure. One person operates the ground
manual switch to retract the RAT. The other person monitors the RAT movement to make
sure the RAT blades do not ouch the airplane structure.
Access MLG Wheel Well Wing to Body - Aft Lower Half (Right)
CAUTION:
REMOVE THE RAT SAFETY SCREEN BEFORE YOU RETRACT THE RAT. IF THE RAT IS
RETRACTED WITH THE SAFETY SCREEN ON, DAMAGE CAN OCCUR TO THE AIRPLANE
EQUIPMENT.
If it is installed, remove the RAT safety screen.
Supply electrical power.
757 General Familiarization (7-2005)
ATA 29 - Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Retract the RAT (Continued):
Make sure the battery switch, on the P5 panel, is turned ON.
Remove the RAT circuit breaker lockset and close these circuit breakers on the overhead panel, P11:
1. 11D26, HYDRAULIC RAT CONT or HYDRAULIC RAT AUTO CONT
2. 11D27, HYDRAULIC RAT AUTO or HYDRAULIC RAT AUTO PWR
Remove the RAT circuit breaker lockset and close these circuit breakers on the main power distribution panel, P6:
1. 6F1, RAT MAN or RAT MAN PWR
2. 6F2, RAT MAN CONT
Pressurize the right hydraulic system.
Remove the protective covers from the RAT turbine blades and the turbine hub.
Move the turbine hub to align the index marks on the turbine hub and the RAT strut.
WARNING:
KEEP PERSONS AND EQUIPMENT AWAY FROM THE PATH OF THE RAT AND THE RAT
COMPARTMENT DOOR. THE RAT AND THE RAT COMPARTMENT DOOR MOVE QUICKLY
AND CAN CAUSE INJURY TO PERSONS OR DAMAGE TO EQUIPMENT.
Hold the ground manual switch in the up (stow) position until the RAT is fully retracted.
Make sure the RAT UNLOCKED light on the overhead panel, P5, is off.
757 General Familiarization (7-2005)
ATA 29 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Retract the RAT (Continued):
Close the six EICAS circuit breakers on the P11 panel.
Push the ELEC/HYD switch on the EICAS MAINT panel, on the right side panel, P61.
Make sure the RAT UNLOCKED and RAT messages do not show on the EICAS display.
Install the access panel for the ground manual switch in the fairing on the aft bulkhead of the right wheel well.
WARNING:
USE THE PROCEDURE IN AMM 32-00-15/201 TO REMOVE THE DOOR LOCKS. THE
DOORS OPEN AND CLOSE QUICKLY AND CAN CAUSE INJURY TO PERSONS OR DAMAGE
TO EQUIPMENT.
Remove the door locks from the landing gear doors and close the doors.
Remove the power from the right hydraulic system if it is not necessary.
757 General Familiarization (7-2005)
ATA 29 - Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC POWER TRANSFER UNIT (PTU) SYSTEM
Overview
The power transfer unit (PTU) system provides an alternate power source for the left hydraulic system, if left
engine power or left EDP pressure is lost. This system powers only the flaps, slats, landing gear, and nose gear
steering. The PTU uses right system pressure to provide left system pressure without transferring fluid between
the two systems. The PTU contains in one housing a hydraulic motor and a hydraulic pump joined by a driveshaft.
The left engine-out relay No. 2 turns the PTU on if power is lost in the left engine.
On some aircraft the PTU is also turned on by the PTU control pressure switch if left EDP pressure falls below
approximately 1275 PSI.
The PTU can be turned on with the PTU manual control switch on right side panel P61.
757 General Familiarization (7-2005)
ATA 29 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Hydraulic Power Transfer Unit (PTU) System Components
Power Transfer Unit (PTU)
The power transfer unit consists of a hydraulic motor and a hydraulic pump in one housing. The motor and pump
are joined by a driveshaft inside the housing. The hydraulic motor uses 28 GPM at 3000 PSI to provide a pump
output of 22.5 GPM at 2175 PSI. As pump output pressure increases, pump delivery flow decreases until flow is
near zero at 2600 PSI. The PTU is on the inboard wall of the left main wheel well.
IF 29-27
757 General Familiarization (7-2005)
ATA 29 - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PTU Pressure Filter Module
The PTU module filters the pump pressure flow before the fluid enters the left system. The filter module is next to
the PTU, in the left wheel well. The module has a disposable filter element. A differential pressure indicator shows
when the filter is blocked. A spring-loaded valve in the module prevents fluid draining when the filter element is
removed. The module has a check valve, a control pressure switch, and an indication pressure switch.
IF 29-28
757 General Familiarization (7-2005)
ATA 29 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PTU Case Drain Filter Module
The PTU case drain filter module consists of a disposable filter element, a differential pressure indicator, a check
valve, a relief valve, and a spring-loaded shutoff valve. The element filters case drain fluid from the PTU pump.
The differential pressure indicator shows when the filter is blocked. The checkk valve prevents reverse flow when
the left system is pressurized. The relief valve opens to allow fluid to bypass the filter, if the filter becomes
plugged. When the filter case is removed, the shutoff valve and check valve prevent loss of fluid. The filter module
is in the left main wheel well, next to the PTU.
IF 29-29
757 General Familiarization (7-2005)
ATA 29 - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PTU Control Valve
The control valve is a shutoff valve in the PTU motor supply line. The valve is on the inboard wall of the right
wheel well. The two-position valve is operated by a 28-volts DC motor. In the ON position, the valve supplies right
system pressure to operate the hydraulic motor. In the OFF position, the valve shuts off pressure to the motor.
IF 29-30
757 General Familiarization (7-2005)
ATA 29 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PTU Flow Limiter Valve (Not on all aircraft)
The flow limiter valve controls fluid flow rate to the PTU as the right hydraulic system pressure varies. The flow
rate varies from above 4.5 GPM at 1800 PSI supply to about 40 GPM at a supply pressure of 2400 PSI. The flow
limiter valve is on the inboard wall of the right wheel well.
IF 29-31
757 General Familiarization (7-2005)
ATA 29 - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PTU Control Pressure Switch (Not on all aircraft)
The control pressure switch monitors left EDP pressure and turns on the PTU if the left EDP output pressure drops
below approximately 1275 PSI. The control pressure switch is in the left EDP pressure line and is adjacent to the
EDP pressure/case drain filter module in the left engine strut.
IF 29-32
757 General Familiarization (7-2005)
ATA 29 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Left EDP Pressure Switch (Not on all aircraft)
The left EDP pressure switch, which controls left EDP low pressure indication, also turns on the PTU if left EDP
pressure drops below approximately 1275 PSI (Ref 29-11-00 for EDP pressure switch location and more
information).
IF 29-33
757 General Familiarization (7-2005)
ATA 29 - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC PRESSURE INDICATING SYSTEM
Overview
A hydraulic pressure indicating system is provided for each main hydraulic system. This system provides
indication of fluid pressure and low pressure warning for each hydraulic system. Low pressure warning lights are
on the hydraulic control panel in the flight compartment. A pressure indicating light for the RAT is on the engine
start/RAT control panel in the flight compartment. The EICAS display contains a digital readout of hydraulic
system pressure. This display also provides low pressure warning messages. The pressure indicating system
inputs to the EICAS computer are powered by 28-volts DC from circuit breakers on overhead panel P11. The low
pressure warning lights are powered by 28-volts DC from master dim and test circuit breakers on panel P11.
757 General Familiarization (7-2005)
ATA 29 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC PRESSURE INDICATING SYSTEM COMPONENTS
EICAS Display
The EICAS display shows a digital readout of pressure in each system when the status mode or the ELEC/HYD
maintenance mode is selected. Messages which describe problems relating to low pressure warning lights
automatically appear on the EICAS display. The messages for the pressure indicating system consist of the
following:
L HYD SYS PRESS
R HYD SYS PRESS
C HYD SYS PRESS
L HYD ENG PUMP
R HYD ENG PUMP
C HYD ELEC 1
C HYD ELEC 2
L HYD ELEC PUMP
R HYD ELEC PUMP
L HYD SYS MAINT
R HYD SYS MAINT
C HYD SYS MAINT
IF 29-34
757 General Familiarization (7-2005)
ATA 29 - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Low Pressure Lights
Three system low pressure lights are on the hydraulic control panel. The system light turns on when activated by
the low pressure switch. In the left and right systems, the switch is on the EDP pressure/case
re/case drain filter module.
In the center system, the switch is on the ACMP pressure/case drain filter module.
A low pressure light is provided on the hydraulic control panel for each pump. The lights are in the pump select
switches. A pump low pressure light illuminates if the output of a pump in operation becomes low. The low
pressure switches are on the pump pressure/case drain filter modules.
IF 29-35
757 General Familiarization (7-2005)
ATA 29 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
RAT Pressure Light
A pressure indicating light for the RAT hydraulic pump is in the manual deployment switch on the engine
start/RAT control panel. This light turns on when the RAT pump is supplying pressure. The pump pressure switch
is on the RAT checkout module.
IF 29-36
757 General Familiarization (7-2005)
ATA 29 - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pressure Transmitters
A pressure line mounted transmitter in each system converts
ts pressure to an electrical signal. The transmitters
provide a voltage proportional to system pressure. The EICAS computer changes voltage signals from the
transmitters into pressure readings on the EICAS display. The transmitters for the left and center systems are in
the left wheel well. The right system transmitter is in the right wheel well.
IF 29-37
757 General Familiarization (7-2005)
ATA 29 - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC FLUID TEMPERATURE INDICATING SYSTEM
Overview
A hydraulic fluid temperature indicating system is provided for each main hydraulic system. This system indicates
the fluid temperature in each reservoir and the overheat in each pump. The pump overheat warning lights are on
the hydraulic control panel in the flight compartment. The fluid temperature in each reservoir is shown as a digital
readout on the EICAS display. The EICAS display also provides overheat warning messages. The overheat warning
lights are powered by 28-volts DC from master dim and test circuit breakers on overhead panel P11.
757 General Familiarization (7-2005)
ATA 29 - Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC FLUID TEMPERATURE INDICATING SYSTEM COMPONENTS
EICAS display
The EICAS display shows a digital readout of fluid temperature in each reservoir when the ELEC/HYD maintenance
mode is selected. Messages which describe problems relating to pump overheat warning lights automatically
appear on the display. The messages consist of the following:
L ELEC HYD OVHT
R ELEC HYD OVHT
C HYD 1 OVHT
C HYD 2 OVHT
L ENG HYD OVHT
R ENG HYD OVHT
IF 29-38
757 General Familiarization (7-2005)
ATA 29 - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overheat Lights
An overheat warning light for each pump is on the hydraulic controll panel. Each light is below the related pump
select switch.
IF 29-39
757 General Familiarization (7-2005)
ATA 29 - Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overheat Light Temperature Switches
The overheat temperature switch for each ACMP is inside the pump. The switch for each EDP is on the EDP
pressure/case drain filter module.
IF 29-40
757 General Familiarization (7-2005)
ATA 29 - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Temperature Transmitters
Each reservoir temperature transmitter is installed on the system reservoir.
IF 29-41
757 General Familiarization (7-2005)
ATA 29 - Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC FLUID QUANTITY INDICATING SYSTEM
Overview
A hydraulic fluid quantity indicating system is provided for each main hydraulic system. This system indicates
fluid quantity in each hydraulic reservoir. Low quantity warning lights are on the hydraulic control panel in the
flight compartment. The EICAS display contains a digital readout of fluid quantity in each reservoir. The display
also provides low quantity messages. A remote quantity indicator is at the central ground servicing station. The
quantity indicating system is powered by 115-volts AC from overhead circuit breaker panel P11. The low quantity
warning lights are powered by 28-volts DC from master dim and test circuit breakers on panel P11.
757 General Familiarization (7-2005)
ATA 29 - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC FLUID QUANTITY INDICATING SYSTEM
EM COMPONENTS
EICAS Display
The EICAS display shows a digital readout of fluid quantity in each reservoir in both the status and maintenance
modes. The quantity readout is a decimal number indicating a fraction of full.
An OF (overfull) indication is displayed next to the digital readout of fluid quantity in the EICAS maintenance mode
for readings above 1.22. The OF indication is generated in the EICAS computer. In addition, a (L, R, C) HYD QTY
O/FULL message is displayed on the EICAS maintenance page.
A RF (refill) indication is displayed next to the digital readout of fluid quantity in the EICAS status and
maintenance modes for readings below 0.75. The RF indication is generated in the EICAS computer.
A (L, R, C) HYD QTY level C message is displayed on EICAS to indicate a low fluid level in the system reservoir.
IF 29-42
757 General Familiarization (7-2005)
ATA 29 - Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Quantity Warning Lights
A low quantity warning light for each hydraulic system is on the hydraulic control panel. The warning light
illuminates to indicate low fluid level in the system reservoir. This light will also illuminate if reservoir air
pressure is low.
IF 29-43
757 General Familiarization (7-2005)
ATA 29 - Page 67
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Remote Quantity Indicator
A quantity gage is at the central ground servicing station. This gage shows the fluid level in the reservoir being
serviced.
IF 29-44
757 General Familiarization (7-2005)
ATA 29 - Page 68
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Sight Glasses
One sight glass is on each reservoir to visually check the full fluid level.
Quantity Transmitter
The quantity transmitter is a variable capacitor inside each reservoir. The capacitance of the transmitter varies
with the change of depth of fluid within thee reservoir.
IF 29-45
757 General Familiarization (7-2005)
ATA 29 - Page 69
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Quantity Monitor Unit
The monitor unit changes the signal from the quantity transmitters into voltage signals to the EICAS computer, the
low quantity lights, and the remote quantity gage.
IF 29-46
757 General Familiarization (7-2005)
ATA 29 - Page 70
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOTES:
757 General Familiarization (7-2005)
ATA 29 - Page 71
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 30
ICE AND RAIN PROTECTION - DESCRIPTION AND OPERATION ................................................................................ 3
Wing Thermal Anti-Icing .................................................................................................................................... 3
Engine Inlet Thermal Anti-Icing.......................................................................................................................... 3
Pitot Probe Anti-Icing ......................................................................................................................................... 3
Angle of Attack Probe Heat ................................................................................................................................ 3
Total Air Temperature Probe Heat ..................................................................................................................... 3
Engine Probe Heat .............................................................................................................................................. 3
Flight Compartment Window Anti-Icing .............................................................................................................. 3
Water and Drain Line Heaters ............................................................................................................................ 3
Ice and Rain Protection (Continued): .................................................................................................................. 4
Windshield Wiper System ................................................................................................................................... 4
Rain Repellent System ........................................................................................................................................ 4
WING THERMAL ANTI-ICING SYSTEM........................................................................................................................ 5
Overview............................................................................................................................................................. 5
WING THERMAL ANTI-ICING COMPONENTS .......................................................................................................... 6
Wing Thermal Anti-Ice (TAI) Valve .................................................................................................................... 6
Telescoping Duct ................................................................................................................................................ 7
Flexible Coupling ................................................................................................................................................ 8
Spray Tubes ....................................................................................................................................................... 9
Wing and Engine Anti-Ice Control Panel ........................................................................................................... 10
Test Panel ........................................................................................................................................................ 10
ENGINE INLET THERMAL ANTI-ICING SYSTEM ....................................................................................................... 11
Overview........................................................................................................................................................... 11
ENGINE INLET THERMAL ANTI-ICING SYSTEM COMPONENTS ............................................................................... 12
Engine Inlet Thermal Anti-Ice (TAI) Valves ...................................................................................................... 12
Engine Inlet Thermal Anti-Ice (TAI) Pressure Switches ................................................................................... 13
RB211-535E4 SERIES ENGINES WITH THE TAI VALVE POSITION SWITCH ..................................................... 13
Engine Inlet Thermal Anti-Ice (TAI) Ducts ....................................................................................................... 14
ENGINES WITH RR SB71-9634 ......................................................................................................................... 14
Wing and Engine Anti-Ice Control Panel ........................................................................................................... 15
PITOT PROBE ANTI-ICING SYSTEM ......................................................................................................................... 16
Overview........................................................................................................................................................... 16
PITOT PROBE ANTI-ICING SYSTEM COMPONENTS................................................................................................. 17
Pitot Probes ...................................................................................................................................................... 17
Auxiliary Annunciator Panel ............................................................................................................................. 18
Miscellaneous Test Panel ................................................................................................................................. 18
ANGLE OF ATTACK PROBE HEAT SYSTEM.............................................................................................................. 19
Overview........................................................................................................................................................... 19
ANGLE OF ATTACK PROBE HEAT SYSTEM COMPONENTS ................................................................................. 20
757 General Familiarization (7-2005)
ATA 30 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AOA Probes ...................................................................................................................................................... 20
Auxiliary Annunciator Panel ............................................................................................................................. 21
Miscellaneous Test Panel ................................................................................................................................. 21
TOTAL AIR TEMPERATURE PROBE HEAT SYSTEM ................................................................................................. 22
Overview........................................................................................................................................................... 22
TOTAL AIR TEMPERATURE PROBE HEAT SYSTEM COMPONENTS .................................................................... 23
TAT Probe ........................................................................................................................................................ 23
Auxiliary Annunciator Panel ............................................................................................................................. 24
Miscellaneous Test Panel ................................................................................................................................. 24
ENGINE PROBE HEAT SYSTEM ................................................................................................................................ 25
Overview........................................................................................................................................................... 25
ENGINE PROBE HEAT SYSTEM COMPONENTS.................................................................................................... 26
Engine Probes .................................................................................................................................................. 26
FLIGHT COMPARTMENT WINDOW ANTI-ICING SYSTEM ......................................................................................... 27
Overview........................................................................................................................................................... 27
FLIGHT COMPARTMENT WINDOW ANTI-ICING SYSTEM COMPONENTS ............................................................ 28
Windows ........................................................................................................................................................... 28
Window Heat Control Unit (WHCU) ................................................................................................................... 28
Window Heat Control Panel .............................................................................................................................. 29
Miscellaneous Test Panel ................................................................................................................................. 29
WINDSHIELD WIPER SYSTEM .................................................................................................................................. 30
Overview........................................................................................................................................................... 30
WINDSHIELD WIPER SYSTEM COMPONENTS...................................................................................................... 31
Windshield Wiper/Rain Repellent Control Panel .............................................................................................. 31
Windshield Wiper Motor/Converter .................................................................................................................. 32
Windshield Wiper Arm ...................................................................................................................................... 33
Windshield Wiper Blade ................................................................................................................................... 33
WINDSHIELD RAIN REPELLENT SYSTEM ................................................................................................................. 34
Overview........................................................................................................................................................... 34
Airplanes with Hydrophobic Coating on the Windshield; .................................................................................. 34
WINDSHIELD RAIN REPELLENT SYSTEM COMPONENTS ........................................................................................ 35
Windshield Wiper/Rain Repellent Control Panel .............................................................................................. 35
Solenoid Valves ................................................................................................................................................ 36
Spray Nozzles ................................................................................................................................................... 37
Rain Repellent Container .................................................................................................................................. 38
Pressure Gage .................................................................................................................................................. 38
Accumulator ..................................................................................................................................................... 39
WATER AND DRAIN LINE HEATERS SYSTEM........................................................................................................... 41
Overview........................................................................................................................................................... 41
WATER AND DRAIN LINE HEATERS SYSTEM COMPONENTS .............................................................................. 42
Heater Tapes, Blankets .................................................................................................................................... 42
Heaters Thermostats ........................................................................................................................................ 42
Drain Pipe Heating Gaskets .............................................................................................................................. 42
Drain Mast Heaters .......................................................................................................................................... 42
757 General Familiarization (7-2005)
ATA 30 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ICE AND RAIN PROTECTION - DESCRIPTION AND OPERATION
Overview
The anti-icing system ensures operation and control of airplane systems under ice conditions. The anti-icing system is made up of five
subsystems: wing thermal anti-icing, engine inlet thermal anti-icing, probe anti-icing, flight compartment window anti-icing, and water
and drain line heaters.
Wing Thermal Anti-Icing
The wing thermal anti-icing system uses engine bleed air to prevent ice formation on the leading edge slats. System components include
thermal anti-ice (TAI) valves, telescoping ducts, flexible couplings, spray tubes, wing and engine anti-ice control panel, and test panel.
Engine Inlet Thermal Anti-Icing
The engine inlet thermal anti-icing system uses engine bleed air to prevent ice formation on the engine inlet cowls. Engine P1 probes are
electrically heated. System components include engine inlet thermal anti-ice (TAI) valves, engine inlet TAI pressure switches, engine inlet
cowl ducts, engine P1 probes, and wing and engine anti-ice control panel.
Pitot Probe Anti-Icing
Pitot probes are electrically heated to prevent erroneous readings due to ice formation. System components are pitot probes, the test
panel, and the annunciator panel.
Angle of Attack Probe Heat
Angle of attack probes are electrically heated to prevent erroneous readings due to ice formation. System components are angle of attack
probes, the test panel, and the annunciator panel.
Total Air Temperature Probe Heat
The total air temperature probe is electrically heated to prevent erroneous readings due to ice formation. System components are the total
air temperature probe, test panel, and annunciator panel.
Engine Probe Heat
Engine P1 probes are electrically heated to prevent erroneous readings due to ice formation.
Flight Compartment Window Anti-Icing
Flight compartment windows are electrically heated by controllers with built-in-test (BIT) capability. System components include window
heat control units, test panel, and window heat control panel.
Water and Drain Line Heaters
Water and drain lines are heated to prevent ice formation. System components include heater tapes, heater tape thermostats, heating
gaskets, and drain mast heaters.
757 General Familiarization (7-2005)
ATA 30 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Ice and Rain Protection (Continued):
Windshield Wiper System
Electrically operated windshield wipers increase visibility during rain and snow. Components of the windshield wiper system include:
windshield wiper motors, windshield wiper arms, windshield wiper blades, and windshield wiper/rain repellent control panel.
Rain Repellent System
Rain repellent works with the wipers to increase visibility in heavy rain or snow. The rain repellent system consists of a rain repellent
bottle, rain repellent valves, spray nozzles, accumulator, and windshield wiper/rain repellent control panel.
757 General Familiarization (7-2005)
ATA 30 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WING THERMAL ANTI-ICING SYSTEM
Overview
Engine bleed air provides in-flight thermal anti-icing to three outboard slats on each wing. System components
include: wing thermal anti-ice (TAI) valves, telescoping ducts, spray tubes, wing and engine anti-ice control panel,
and a test panel.
System electrical power comes from a DC bus, through a circuit breaker on overhead circuit breaker panel P11.
757 General Familiarization (7-2005)
ATA 30 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WING THERMAL ANTI-ICING COMPONENTS
Wing Thermal Anti-Ice (TAI) Valve
The wing TAI valve controls the flow of engine bleed air into the leading edge anti-ice
ti-ice ducts. One valve is located
above the strut position in each wing. The valve is set to keep duct air pressure between 20-28 PSI.
An indicator on the valve shows when it is CLOSED or NOT CLOSED. The valve can be manually set to the full
open or full closed position by rotating a hex head bolt on the valve housing. The valve can only be locked in the
closed position.
IF 30-1
757 General Familiarization (7-2005)
ATA 30 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Telescoping Duct
Each wing has a telescoping duct connecting the wing supply manifold
ld to the leading edge distribution duct. Bleed
air can be fed to the leading edge slats regardless of slat position.
IF 30-2
757 General Familiarization (7-2005)
ATA 30 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flexible Coupling
Flexible couplings join the leading edge ducts where the slats are separated.
parated. The couplings allow for alignment
errors during slat movements. There are 2 couplings per wing, between the leading edge slats.
IF 30-3
757 General Familiarization (7-2005)
ATA 30 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Spray Tubes
Ducting on the three outboard slats on each wing have spray holes to direct bleed air to the leading edge surfaces.
IF 30-4
757 General Familiarization (7-2005)
ATA 30 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Wing and Engine Anti-Ice Control Panel
The wing and engine anti-ice control panel M10397 controls wing thermal anti-icing with the WING ANTI-ICE
switch. The M10397 panel also controls engine inlet thermal anti-icing. The M10397 panel is on pilots' overhead
panel P5.
Test Panel
The WING ANTI-ICE WINDOW/PROBE HEAT switch on the M10398 test panel allows a ground test of the wing TAI
system. The M10398 panel is on right side panel P61.
IF 30-5
757 General Familiarization (7-2005)
ATA 30 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE INLET THERMAL ANTI-ICING SYSTEM
Overview
The engine inlet cowls are anti-iced by engine bleed air. System components include engine inlet thermal anti-ice
(TAI) valves, engine inlet TAI low and high pressure switches, engine inlet TAI ducts, and a wing and engine antiice control panel. System electrical power comes from the 28-volts DC battery bus, through circuit breakers on the
overhead circuit breaker panel, P11.
Engine anti-ice maintenance operation should be limited to bleed duct temperature of 450 degrees F maximum for
less than 5 minutes duration.
757 General Familiarization (7-2005)
ATA 30 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE INLET THERMAL ANTI-ICING SYSTEM COMPONENTS
Engine Inlet Thermal Anti-Ice (TAI) Valves
The engine inlet TAI valve controls the flow of engine bleed air to the engine inlet TAI duct. One valve is located in
the left side of each engine, aft of the pressure blow-out door. The valve is set to keep duct pressure between 2428 PSI.
The TAI valve can be manually set and locked in one of two positions (VALVE LOCKED PART OPEN or VALVE
LOCKED CLOSED) by a manual override assembly and an engraved locking plate. These are installed at the top of
the valve and do not move or operate when the valve is in the normal position.
IF 30-6
757 General Familiarization (7-2005)
ATA 30 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Inlet Thermal Anti-Ice (TAI) Pressure Switches
Two pressure switches (high and low) are located on the left side of each engine. Thee high pressure switch is
above the TAI valve and the low pressure switch is below the TAI valve.
The low pressure switch is used to indicate valve position. The high pressure switch senses a failure in the
pressure regulating function of the TAI valve.
RB211-535E4 SERIES ENGINES WITH THE TAI VALVE POSITION SWITCH
A valve position switch is installed at the top of the engine TAI valve and below the manual override assembly (the
jackscrew and tongue). During the normal valve operation, the valve position switch is electrically in series with
the low pressure switch.
The movement of the engine TAI valve mechanically operates the valve position switch, which opens or closes
when the valve does. The valve position switch prevents an unwanted display of the VALVE disagree light on the
M10397 (P5) panel.
IF 30-7
757 General Familiarization (7-2005)
ATA 30 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Inlet Thermal Anti-Ice (TAI) Ducts
The engine TAI ducts carry engine bleed air along the left side of each engine to the inlet cowls.
ENGINES WITH RR SB71-9634
The ducting includes a thermal anti-icing pressure indicator. This gives a visual indication if the TAI supply duct is
defective. An operated pressure indicator will be seen during a pre-flight check.
IF 30-8
757 General Familiarization (7-2005)
ATA 30 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Wing and Engine Anti-Ice Control Panel
The wing and engine anti-ice control panel, M10397, controls the engine inlet thermal anti-icing. Control is
provided with two push-on/push-off L and R ENGINE ANTI-ICE switches. The M10397 panel also
so controls the wing
thermal anti-icing. The panel is located on the pilots' overhead panel, P5.
IF 30-9
757 General Familiarization (7-2005)
ATA 30 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PITOT PROBE ANTI-ICING SYSTEM
Overview
Four pitot probes (two on each side) are electrically heated to prevent erroneous readings due to ice conditions.
Pitot probe anti-icing system components include: pitot probes, the auxiliary annunciator panel, and the
miscellaneous test panel.
System electrical power comes from the 28-volts DC buses, and the 115-volts AC buses. Circuit breakers are on
overhead circuit breaker panel P11, and main power distribution panel P6.
757 General Familiarization (7-2005)
ATA 30 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PITOT PROBE ANTI-ICING SYSTEM COMPONENTS
Pitot Probes
Two pitot probes are located on each side, halfway down the fuselage, under the left and right No. 2 flight deck
windows. Two levels of pitot probe heating are used, depending on whether the plane is on the ground or in flight.
IF 30-10
757 General Familiarization (7-2005)
ATA 30 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Auxiliary Annunciator Panel
Auxiliary annunciator panel M10394 is located on pilots' overhead panel P5. The annunciator panel contains four
amber lights, one for each pitot probe. The lights are identified as: CAPT PITOT, F/O PITOT, L AUX PITOT, and R
AUX PITOT. These lights come on to identify a probe heater failure.
Miscellaneous Test Panel
Miscellaneous test panel M10398 provides for ground testing of the probe heaters. The test panel is on right side
panel P61.
IF 30-11
757 General Familiarization (7-2005)
ATA 30 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ANGLE OF ATTACK PROBE HEAT SYSTEM
Overview
Two angle of attack (AOA) probes (one on each side) are electrically heated to prevent erroneous readings due to
ice conditions. AOA probe heat system components include: AOA probes, the auxiliary annunciator panel, and the
miscellaneous test panel.
System electrical power comes from the 28-volts DC buses, and the 115-volts AC buses. Circuit breakers are on
overhead circuit breaker panel P11, and main power distribution panel P6.
757 General Familiarization (7-2005)
ATA 30 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ANGLE OF ATTACK PROBE HEAT SYSTEM COMPONENTS
AOA Probes
One AOA probe is located on each side, halfway down the fuselage under the left and right No. 2 flight deck
windows. Each AOA probe has a vane heater and a case heater. The vane is heated to maintain an efficient vane
airfoil. The case is heated to ensure free rotation of the vane. Only one level of AOA probe heating is used both on
the ground and in flight.
The vane and case heaters are independently wired to enable positive fault detection in the vane heater circuit.
IF 30-12
757 General Familiarization (7-2005)
ATA 30 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Auxiliary Annunciator Panel
Auxiliary annunciator panel M10394 is located on pilots' overhead panel P5. The annunciator panel contains two
amber lights, one for each AOA probe. The lights are identified as: L AOA, R AOA. These lights come on to identify
a probe heater or system failure.
Miscellaneous Test Panel
Miscellaneous test panel M10398 provides for ground testing of the probe heaters. The test panel is on right side
panel P61.
IF 30-13
757 General Familiarization (7-2005)
ATA 30 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TOTAL AIR TEMPERATURE PROBE HEAT SYSTEM
Overview
The total air temperature (TAT) probe is electrically heated to prevent erroneous readings due to ice conditions.
TAT system components include: the TAT probe, the auxiliary annunciator panel, and the miscellaneous test panel.
System electrical power comes from the right main 28-volts DC bus and the right main 115-volts AC bus. Circuit
breakers are on overhead circuit breaker panel P11, and main power distribution panel P6.
757 General Familiarization (7-2005)
ATA 30 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TOTAL AIR TEMPERATURE PROBE HEAT SYSTEM COMPONENTS
TAT Probe
The TAT probe is located on the right side, halfway down the fuselage, aft of the nose gear. Only one level of TAT
probe heating is used.
IF 30-14
757 General Familiarization (7-2005)
ATA 30 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Auxiliary Annunciator Panel
Auxiliary annunciator panel M10394 is located on pilots' overhead panel P5. The annunciator panel contains an
amber TAT probe light. The light comes on to identify a TAT probe heater failure.
Miscellaneous Test Panel
Miscellaneous test panel M10398 provides for ground testing of the probe heater. The test panel is on right side
panel P61.
IF 30-15
757 General Familiarization (7-2005)
ATA 30 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE PROBE HEAT SYSTEM
Overview
Two engine probes (one for each engine) are electrically heated to prevent erroneous readings due to ice
conditions. Engine probe heat system components consist of the engine probes. System electrical power comes
from the main left and right 28-volts DC buses, and the main left and right 115-volts AC buses. Circuit breakers
are on overhead circuit breaker panel P11.
The engine probes are identified as P1 probes and they will be referred to as the engine probes in this procedure.
757 General Familiarization (7-2005)
ATA 30 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE PROBE HEAT SYSTEM COMPONENTS
Engine Probes
One engine probe is located on the upper inside surface of the inlet cowl. The probe measures total pressure in
the inlet air stream entering the engine LP compressor.
The probe is heated by an electrical heating element which provides one level of heating. The heater uses 115volts AC power.
IF 30-16
757 General Familiarization (7-2005)
ATA 30 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FLIGHT COMPARTMENT WINDOW ANTI-ICING SYSTEM
Overview
The captain's and first officer's No. 1, No. 2 and No. 3 windows are heated. The main components are the
windows, window heat control units, the window heat control panel, and the miscellaneous test panel. System
power comes from the main left 28-volts DC bus, and the main left and right 115-volts AC buses.
757 General Familiarization (7-2005)
ATA 30 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FLIGHT COMPARTMENT WINDOW ANTI-ICING SYSTEM COMPONENTS
Windows
A conductive film or wire net between the outer pane and the core provides electrical heating of the window.
Window Heat Control Unit (WHCU)
Two window heat control units (WHCUs), each with built-in-test equipment (BITE), regulate heating and isolate
system failures. The left WHCU is on shelf 2 of left forward equipment center rack E1. The right WHCU is on shelf
1 of E1. BITE pushbuttons, fault indicating lights, and BITE instructions are on the unit face.
The WHCU uses 200 volts (No. 1 window) and 115 volts (No. 2 and 3 windows) ac power for heating. Twenty-eight
volt dc power is used by the WHCUs for system testing.
IF 30-17
757 General Familiarization (7-2005)
ATA 30 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Window Heat Control Panel
Window heat control panel M10395 is located on pilots' overhead panel P5. The panel contains four WINDOW HEAT
switches to control window heating.
Miscellaneous Test Panel
Miscellaneous test panel M10398 provides for testing of the window heat system. The test panel is on right side
panel P61.
IF 30-18
757 General Familiarization (7-2005)
ATA 30 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WINDSHIELD WIPER SYSTEM
Overview
Two windshield wipers clear the captain's and first officer's No. 1 windows during takeoff, approach, and landing
in rain or snow. Each wiper is driven by a separate motor to ensure clear vision through one window if one motor
fails. System components include: the windshield wiper/rain repellent control panel, windshield wiper
motor/converter, windshield wiper arm, and windshield wiper blades. The system uses main left and right 28-volts
DC power, through circuit breakers on overhead circuit breaker panel P11.
757 General Familiarization (7-2005)
ATA 30 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WINDSHIELD WIPER SYSTEM COMPONENTS
Windshield Wiper/Rain Repellent Control Panel
Windshield wiper/rain repellent control panel M10023 is on pilots' overhead panel P5. The control panel contains a
switch to control the windshield wiper system. Switches controlling the rain repellent system
em are also on the panel.
IF 30-19
757 General Familiarization (7-2005)
ATA 30 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Windshield Wiper Motor/Converter
The two windshield wipers are driven by separate motor/converters.
converters. The motor/converter is a single unit. Each
unit is mounted to an access panel below its respective window.
The rotary motion of the motor is changed by the converter to produce the sweeping motion of the wiper arm. The
motor runs at low or high speed, depending on the wiper speed chosen by a switch setting in the flight deck. A
parking switch in the motor/converter sets the wiper blade to the park position when the wiper is turned off. The
windshield motors use 28-volts DC power. A connector on each unit supplies power and control.
IF 30-20
757 General Familiarization (7-2005)
ATA 30 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Windshield Wiper Arm
The wiper arm connects the output shaft of the motor/converter to the wiper blade.
Windshield Wiper Blade
The wiper blade consists of a wiper edge connected to a metal channel. This channel attaches to the wiper arm.
IF 30-21
757 General Familiarization (7-2005)
ATA 30 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WINDSHIELD RAIN REPELLENT SYSTEM
Overview
Rain repellent is sprayed on the captain's and/or first officer's No. 1 window to provide a water repellent coating.
The system uses main left and right 28-volts DC power, through circuit breakers on overhead circuit breaker panel
P11.
System components include:
1. The windshield wiper/rain repellent control panel
2. Solenoid valves
3. Spray nozzles
4. A rain repellent container
5. A pressure gage
6. An accumulator
Airplanes with Hydrophobic Coating on the Windshield;
The rain repellent system consists of a hydrophobic coating applied to the number 1 window. No pilot actions are
necessary to operate the system.
Periodic maintenance is necessary to test and restore the coating on the windshield surface.
757 General Familiarization (7-2005)
ATA 30 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WINDSHIELD RAIN REPELLENT SYSTEM COMPONENTS
Windshield Wiper/Rain Repellent Control Panel
Windshield wiper/rain repellent control panel M10023 is on pilots' overhead panel P5. The control panel contains
two switches to control the rain repellent system. A switch controlling the windshield wipers is also on the panel.
IF 30-22
757 General Familiarization (7-2005)
ATA 30 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Solenoid Valves
Two electrically operated solenoid valves control the flow of rain repellent to the No. 1 windows. The right solenoid
valve is behind first officer's instrument panel P3. The left solenoid valve is behind captain's instrument panel P1.
IF 30-23
757 General Familiarization (7-2005)
ATA 30 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Spray Nozzles
Two nozzles on the fuselage and forward of the windshield, direct the rain repellent
ent spray onto the windshields.
IF 30-24
757 General Familiarization (7-2005)
ATA 30 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rain Repellent Container
A pressurized rain repellent container is located forward of the left hand-side flight deck partition. The container is
replaced when empty.
Pressure Gage
A pressure gage indicates when repellent container replacement is needed. The pressure gage is located on the
flight deck partition, aft of the first observer's seat, next to the rain repellent container.
ainer.
IF 30-25
757 General Familiarization (7-2005)
ATA 30 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Accumulator
An accumulator is installed to prevent back flow of rain repellent fluid into the pneumatic system in the case of
check valve failure. The accumulator is located forward of the pilot's center instrument panel, left of the airplane
centerline.
IF 30-26
757 General Familiarization (7-2005)
ATA 30 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 30 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WATER AND DRAIN LINE HEATERS SYSTEM
Overview
Water supply and drain lines are electrically heated to prevent ice formation. System components include heater
tapes, blankets, heater tape thermostats, drain pipe heating gaskets, and drain mast heaters. System electrical
power comes from the 115-volts AC ground service bus through circuit breakers on right miscellaneous electrical
equipment panel P37.
757 General Familiarization (7-2005)
ATA 30 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WATER AND DRAIN LINE HEATERS SYSTEM COMPONENTS
Heater Tapes, Blankets
Heater tapes, blankets are wrapped on some water supply and drain lines to prevent freezing. Temperatures of the
heaters are controlled by thermostats. Some heaters are equipped with integral thermostats, while others use
external thermostats.
Heaters Thermostats
Thermostats are used to control the temperatures of the heaters. The thermostats will open when their
temperatures exceeds 60 ±5°F (15.5 ±3°C), and will re-close when temperatures drops below 45 ±5°F (7.2 ± 3°C).
Drain Pipe Heating Gaskets
Heating gaskets are installed on the fitting ends of the toilet drain pipes. The heating gaskets are not temperaturecontrolled by thermostats.
Drain Mast Heaters
The forward and aft drain masts are heated to allow for in-flight drainage without freezing. Drain mast heater
temperature is controlled by an air/ground relay. Low heat is provided on the ground, and high heat is provided
in-flight.
757 General Familiarization (7-2005)
ATA 30 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 30-27
757 General Familiarization (7-2005)
ATA 30 - Page 43
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 32
LANDING GEAR SYSTEM ............................................................................................................................................ 6
Overview............................................................................................................................................................. 6
Nose Landing Gear and Doors ............................................................................................................................ 6
Main Landing Gear and Doors ............................................................................................................................ 6
Extension and Retraction .................................................................................................................................... 6
Wheels and Brakes ............................................................................................................................................. 6
Steering .............................................................................................................................................................. 6
Position and Warning ......................................................................................................................................... 7
Air/Ground Relays.............................................................................................................................................. 7
NOSE LANDING GEAR AND DOORS ............................................................................................................................ 8
NOSE GEAR............................................................................................................................................................. 8
NOSE GEAR COMPONENTS ........................................................................................................................................ 9
Drag Strut........................................................................................................................................................... 9
Lock Link ............................................................................................................................................................ 9
Lock Spring ........................................................................................................................................................ 9
Shock Strut ........................................................................................................................................................... 10
Torsion Link ......................................................................................................................................................... 12
Tow Fitting............................................................................................................................................................ 13
NOSE GEAR DOORS .................................................................................................................................................. 14
NOSE GEAR DOORS COMPONENTS .......................................................................................................................... 15
Forward Doors .................................................................................................................................................. 15
Forward Door Operating Mechanism ................................................................................................................ 15
Aft Doors .......................................................................................................................................................... 16
Aft Door Operating Mechanism ........................................................................................................................ 16
MAIN LANDING GEAR & DOORS ............................................................................................................................... 17
Main Landing Gear ........................................................................................................................................... 17
Main Landing Gear Components ........................................................................................................................... 18
Trunnion Link ................................................................................................................................................... 18
Drag Strut......................................................................................................................................................... 18
Shock Strut....................................................................................................................................................... 19
Torsion Links ................................................................................................................................................... 20
Truck Assembly................................................................................................................................................ 21
Side Strut and Down-lock Assembly ................................................................................................................ 22
Reaction Link .................................................................................................................................................... 23
LANDING GEAR EXTENSION AND RETRACTION SYSTEM ........................................................................................ 24
Overview........................................................................................................................................................... 24
Extension and Retraction of the Nose Landing Gear ......................................................................................... 24
Landing Gear Alternate Extension ..................................................................................................................... 24
LANDING GEAR CONTROL SYSTEM.......................................................................................................................... 25
LANDING GEAR SYSTEM CONTROL COMPONENTS ................................................................................................. 26
757 General Familiarization (7-2005)
ATA 32 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Control Lever Module ....................................................................................................................................... 26
Selector Valve ................................................................................................................................................... 27
EXTENSION AND RETRACTION OF THE MAIN LANDING GEAR ................................................................................ 28
EXTENSION AND RETRACTION OF THE MAIN LANDING GEAR COMPONENTS ....................................................... 29
Retract Actuator................................................................................................................................................ 29
Door Actuator ................................................................................................................................................... 30
Down-lock Actuator .......................................................................................................................................... 31
Up-lock Actuator ............................................................................................................................................... 32
Door-Operated Gear Sequence ......................................................................................................................... 33
Down-lock Operated Door Sequence Valve ....................................................................................................... 34
Up-lock Operated Sequence Valve .................................................................................................................... 35
Up-lock Assembly ............................................................................................................................................. 35
Truck Positioner Actuator ................................................................................................................................. 36
Truck Positioner Shuttle Valve ......................................................................................................................... 37
NOSE GEAR EXTENSION AND RETRACTION COMPONENTS .................................................................................... 38
Retract Actuator................................................................................................................................................ 38
Lock Actuator ................................................................................................................................................... 39
Gear-Operated Sequence Valve......................................................................................................................... 40
Door-Operated Sequence Valve ........................................................................................................................ 41
Door Actuator ................................................................................................................................................... 41
Flow Control Valve ............................................................................................................................................ 41
Gear Sequence Valve Bypass Valve or Lock Unload Valve Module ................................................................... 42
LANDING GEAR ALTERNATE EXTENSION SYSTEM .................................................................................................. 43
Overview........................................................................................................................................................... 43
LANDING GEAR ALTERNATE EXTENSION SYSTEM COMPONENTS ......................................................................... 44
Alternate Extension Switch ............................................................................................................................... 44
Power Pack....................................................................................................................................................... 45
Door Lock Release Actuator.............................................................................................................................. 46
Door Safety Valve ............................................................................................................................................. 46
Door Release Interlock Actuator ....................................................................................................................... 46
Alternate Up-lock Release Actuator .................................................................................................................. 48
Alternate Extension Hydraulic Shuttle Valve ..................................................................................................... 49
Hydraulic Pressure Switch................................................................................................................................ 50
Door Closed and Locked Switches .................................................................................................................... 52
Door Ground Control Switches.......................................................................................................................... 54
WHEELS AND BRAKES.............................................................................................................................................. 56
Overview........................................................................................................................................................... 56
Tires and Wheels.............................................................................................................................................. 56
Hydraulic Brake System ................................................................................................................................... 56
Antiskid/Auto-brake System ............................................................................................................................ 56
Parking Brake System ...................................................................................................................................... 56
TIRES AND WHEELS ................................................................................................................................................. 57
Overview........................................................................................................................................................... 57
Main and Nose Gear Tires ................................................................................................................................ 57
757 General Familiarization (7-2005)
ATA 32 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TIRES AND WHEELS COMPONENTS ......................................................................................................................... 58
Main Gear Wheels............................................................................................................................................. 58
Nose Gear Wheels............................................................................................................................................. 60
Nose Wheel Spin Brake .................................................................................................................................... 61
HYDRAULIC BRAKE SYSTEM .................................................................................................................................... 63
Overview........................................................................................................................................................... 63
HYDRAULIC BRAKE SYSTEM COMPONENTS ........................................................................................................... 66
Forward Brake Pedal and Quadrant Linkage .................................................................................................... 66
Brake Metering Valve Module ........................................................................................................................... 68
Brake Accumulator ........................................................................................................................................... 70
Alternate Brake Selector Valve (ABSV) ............................................................................................................ 70
Accumulator Isolation Valve (AIV) .................................................................................................................... 70
Hydraulic Brake Assembly................................................................................................................................ 72
Indication.......................................................................................................................................................... 74
ANTISKID/AUTOBRAKE SYSTEM ............................................................................................................................ 76
Overview........................................................................................................................................................... 76
Antiskid System ............................................................................................................................................... 76
Autobrake system ............................................................................................................................................. 78
ANTISKID/AUTOBRAKE SYSTEM COMPONENTS .................................................................................................... 80
Antiskid/Autobrake Crew Control Panels and Annunciators................................................................................. 80
ANTISKID ON/OFF Switch ............................................................................................................................... 80
ANTISKID Fault Light ....................................................................................................................................... 80
Autobrake Selector Switch................................................................................................................................ 82
AUTOBRAKES Light.......................................................................................................................................... 82
Antiskid/Autobrake Control Unit ...................................................................................................................... 83
Antiskid Module (Normal) ................................................................................................................................ 84
Antiskid Module (Alternate) ............................................................................................................................. 84
Antiskid Shuttle Valves ..................................................................................................................................... 85
Antiskid Wheel Speed Transducers .................................................................................................................. 86
Autobrake Module ............................................................................................................................................ 87
Autobrake Shuttle Valve Assembly ................................................................................................................... 88
NOSE WHEEL STEERING SYSTEM ............................................................................................................................ 89
Overview........................................................................................................................................................... 89
NOSE WHEEL STEEERING SYSTEM COMPONENTS ................................................................................................. 92
Tiller, Gearbox, and Torque Limiter...................................................................................................................... 92
Rudder Pedal Steering Interconnect Mechanism................................................................................................... 94
Spring Cartridge and Piston Position Quadrant .................................................................................................... 96
Summing Mechanism and Broken Cable Compensator .................................................................................... 97
Steering Metering Valve Module and Actuators .................................................................................................... 98
Steering Collar and Torsion Links ...................................................................................................................... 100
AIR/GROUND RELAY SYSTEM ............................................................................................................................... 102
Overview......................................................................................................................................................... 102
AIR/GROUND RELAY SYSTEM COMPONENTS ....................................................................................................... 103
Air/Ground Relays.......................................................................................................................................... 103
757 General Familiarization (7-2005)
ATA 32 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Nose Gear Not Compressed Sensors .............................................................................................................. 104
Main Gear Truck Tilt Sensors ........................................................................................................................ 104
LANDING GEAR POSITION INDICATING AND WARNING SYSTEM .......................................................................... 106
Overview......................................................................................................................................................... 106
LANDING GEAR POSITION INDICATING AND WARNING SYSTEM COMPONENTS .................................................. 107
Indicator Lights .............................................................................................................................................. 107
EICAS Display................................................................................................................................................. 107
Main Landing Gear Down and Locked Sensors .............................................................................................. 110
Main Landing Gear Up and Locked Sensor..................................................................................................... 111
Main Landing Gear Door Closed Sensors ....................................................................................................... 112
Nose Gear Locked Sensors ............................................................................................................................. 113
Nose Gear Up Position Sensors ...................................................................................................................... 113
Nose Gear Down Position Sensors ................................................................................................................. 113
Nose Gear Door Closed Sensors ..................................................................................................................... 113
PROXIMITY SWITCH SYSTEM ................................................................................................................................ 114
Overview......................................................................................................................................................... 114
Proximity Switch System Components .................................................................................................................... 115
Proximity Switch Electronics Unit (PSEU) ...................................................................................................... 115
Proximity Sensors........................................................................................................................................... 116
757 General Familiarization (7-2005)
ATA 32 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR SYSTEM
Overview
The landing gear system is a combination of sub-systems which provide control and support of the airplane while
on the ground. The system includes the main and nose landing gear, an extension and retraction system, braking,
steering, and a position indicating system.
Nose Landing Gear and Doors
The nose landing gear consists of one nose gear and associated doors, which support the nose of the airplane. The
gear is located near the nose of the airplane beneath the flight deck.
Main Landing Gear and Doors
The main landing gear consists of two gears and associated doors which support approximately 85 per cent of the
airplanes' weight. The gears are located under the wing, inboard of the engine nacelles.
Extension and Retraction
The extension and retraction system consists of control cables, and hydraulic actuators and valves which control
and operate landing gear movement.
Wheels and Brakes
The airplane is supported on the ground by wheel and tire assemblies mounted on the landing gear. The main gear
wheels are equipped with brakes which aid in stopping the airplane.
An automatic brake control and antiskid protection system aids airplane braking.
A brake temperature monitor system monitors the temperatures of individual brakes and supplies this information
to the flight crew.
Steering
Directional control of the airplane on the ground is provided by a nose wheel steering system. The system consists
of control cables, and hydraulic actuators and valves which rotate the nose wheels.
757 General Familiarization (7-2005)
ATA 32 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Position and Warning
The position and warning system monitors and indicates the status of landing gear and doors position as a check
for proper operation. Sensors on the gears and doors monitor the position, and the information is displayed by
indicator lights in the flight deck.
Air/Ground Relays
Proximity sensors on the nose and main landing gear and truck positioner hydraulic inlet pressure switches
provide inputs to the proximity switch electronics unit (PSEU) to indicate whether the airplane is on the ground or
in the air. The PSEU controls air/ground relay switching which controls the various airplane systems for air or
ground mode operation.
757 General Familiarization (7-2005)
ATA 32 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE LANDING GEAR AND DOORS
NOSE GEAR
The nose gear provides directional control and supports the forward end of the airplane while on the ground. The
nose gear consists of a shock strut, drag strut, lock link, lock springs, torsion link, and tow fitting.
IF 32-1
757 General Familiarization (7-2005)
ATA 32 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE GEAR COMPONENTS
Drag Strut
The drag strut carries forward and aft loads and stabilizes the shock strut. The drag strut is hinged in the middle
and consists of an upper and a lower strut. The upper strut attaches to the wheel well side walls; the lower strut
attaches to the shock strut outer cylinder. During retraction the drag strut folds in the middle.
Lock Link
The lock link consists of a forward lock link, and an aft lock link. The forward end connects at the hinge point of
the drag strut; the aft end attaches to the wheel well aft wall. The links lock in an over center position to keep the
nose gear extended or retracted until the lock actuator (Ref 32-34-00) unlocks it.
Lock Spring
The lock springs attach to the aft lock link and to the forward lock link. The springs keep the lock link in the over
center position until the lock actuator supplies a force to stretch the springs out of the locked position.
IF 32-2
757 General Familiarization (7-2005)
ATA 32 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Shock Strut
The nose gear shock strut is the main supporting member of the nose gear. The shock strut consists of an inner
cylinder and an outer cylinder. The outer cylinder is connected to the wheel well side walls with trunnion pins; the
inner cylinder slides within the outer cylinder.
An oil-air mixture is kept within the cylinders by two active seals, one static and one dynamic. Two sets of spare
seals are stored in the shock strut to aid active seal replacement.
The shock strut is divided by the orifice plate into upper, middle, and lower chambers. The upper chamber
contains oil and nitrogen, the middle and lower chambers contain oil.
During shock strut compression the tapered metering pin adjusts the flow of oil from the middle chamber into the
upper chamber to control the absorption of impact loads. At the same time some oil will flow from the middle
chamber through holes in the upper bearing into the lower chamber.
During strut extension the rebound restrictor decreases the flow of oil from the lower chamber into the middle
chamber to control the rate of strut extension.
Upper and lower centering cams between the inner and outer cylinders assist in aligning the nose gear in the
forward position during retraction.
The shock strut includes lugs for the attachment of the retract actuator, steering actuators, torsion links, drag
strut, forward and aft tow fittings, aft door operating mechanism, and the one-piece axle.
757 General Familiarization (7-2005)
ATA 32 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-3
757 General Familiarization (7-2005)
ATA 32 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Torsion Link
The torsion link consists of an upper and lower link connected to the aft side of the shock strut. The upper link is
attached to the outer cylinder and the lower link is attached to the inner cylinder. The two links are connected
together to allow for vertical extension and compression of the shock strut while preventing rotation between the
inner and outer cylinders except while steering.
CAUTION:
DO NOT USE UPPER TORSION LINK AS A STEP WHEN DISCONNECTED FROM LOWER
LINK. TARGET FOR NOSE GEAR NOT COMPRESSED PROXIMITY SENSOR WILL BE
CONTACTED AND DEFLECTION OF TARGET MAY OCCUR.
The torsion links are disconnected from each other to allow the cylinders to rotate to angles larger than normal
steering limits for towing purposes. When the links are disconnected from each other, the upper link should never
be used as a step while performing maintenance, as any excessive force on the link could damage the air/ground
sensor targets on the sides of the link.
IF 32-4
757 General Familiarization (7-2005)
ATA 32 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Tow Fitting
A forward and an aft tow fitting are attached to the shock strut inner cylinder to allow towing
wing in the forward or aft
direction.
IF 32-5
757 General Familiarization (7-2005)
ATA 32 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE GEAR DOORS
The nose gear doors consist of forward and aft doors. Thee doors close over the wheel well to provide aerodynamic
smoothness during flight.
IF 32-6
757 General Familiarization (7-2005)
ATA 32 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE GEAR DOORS COMPONENTS
Forward Doors
The forward doors are hydraulically operated and close when the gear is retracted or extended. The doors are of
clamshell type and include a left and right door.
Forward Door Operating Mechanism
The mechanism consists of a bell crank, connected to the door actuator; and rods, connected to the bell crank and
forward doors. The mechanism is a mechanical link from the door actuator to the doors.
IF 32-7
757 General Familiarization (7-2005)
ATA 32 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aft Doors
The aft doors are mechanically connected to the gear and only close when the gear is retracted. The doors are of
clamshell type and include a left and right door.
Aft Door Operating Mechanism
The mechanism consists of two rods and a bell crank. The aft rod is connected to the shock strut and bell crank,
the forward rod is connected to the bell crank and the aft door. A left mechanism controls the left door and a right
mechanism controls the right door.
IF 32-8
757 General Familiarization (7-2005)
ATA 32 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
MAIN LANDING GEAR & DOORS
The main landing gear consists of two main gears which absorb landing, taxiing, and take-off loads and support
the airplane while on the ground. The gears are located on each side of thee airplane and are mounted between the
rear wing spar and the main landing gear beam. During flight, the gears retract inboard into the wing cavities and
fuselage wheel wells and are covered by doors. The doors fair into the airplane body contour to reduce
aerodynamic drag.
Main Landing Gear
Most of the airplane weight is supported on the ground by two main gears. Each gear is installed on the wing
between the rear wing spar and the main landing gear beam.
Each gear includes a shock strut, drag strut, side strut and down-lock assembly, trunnion link, reaction link,
torsion links, and truck assembly.
IF 32-9
757 General Familiarization (7-2005)
ATA 32 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Landing Gear Components
Trunnion Link
The trunnion link is installed between the shock strut and the rear wing spar. Loads from the drag strut are
transmitted through the trunnion link to airplane structure. The forward end of the link mounts in a spherical
bearing on the rear wing spar. Under severe impact, the pin connection at the bearing will fail, and allow the gear
to break away from airplane structure with minimal damage. Attachment for the drag strut and strut door are
located on the trunnion link.
Drag Strut
The drag strut absorbs forward and aft loads on the gear. The strut is a one-piece brace which forms a triangle
with the trunnion link and shock strut.
The upper end attaches to the trunnion link near the forward trunnion bearing. The lower end attaches to the shock
strut at the upper torsion link attachment. Fittings on the drag strut provide mounting for the main gear strut door.
IF 32-10
757 General Familiarization (7-2005)
ATA 32 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Shock Strut
The shock strut is the primary supporting member of the main gear. The strut consists of inner and outer
cylinders which are charged with an air-oil mixture to absorb loads during landing, taxiing, and take-off. The shock
strut mounts into a spherical bearing on the main landing gear beam and attaches to the trunnion link.
Upper and lower bearings inside the outer cylinder provide sliding surfaces between the two cylinders. Two active
seals, one static and one dynamic, seal the air-oil mixture between the two cylinders. Two sets of spare seals are
stored in annular grooves in the lower bearing.
Shocks are absorbed by the flow of hydraulic fluid through the annular space between a metering pin and an
orifice plate. The metering pin is tapered to progressively adjust the flow of hydraulic fluid. This provides uniform
control of the loads on airplane structure. A rebound snubber, located just below the upper bearing, acts as a oneway restrictor. During compression, the snubber allows free flow of hydraulic fluid. During extension, the snubber
restricts flow of hydraulic fluid and thus controls rebound of the shock strut.
The outer cylinder provides attachments for the drag strut, lower side strut, and upper torsion link. Also attached
to the outer cylinder is an up-lock roller which engages in the up-lock hook during gear retraction to lock the gear
up. The inner cylinder provides attachments for the truck assembly and lower torsion link.
IF 32-11
757 General Familiarization (7-2005)
ATA 32 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Torsion Links
The torsion links prevent the shock strut inner cylinder from rotating as it moves in and out of the outer cylinder
during normal operation. This maintains truck directional headings at all times.
The upper link connects to the shock strut outer cylinder, the lower link connects to the inner cylinder, and the two
links are connected together to form a hinged joint.
IF 32-12
757 General Familiarization (7-2005)
ATA 32 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Truck Assembly
The truck assembly distributes airplane loads from the shock strutt to the wheels and tires. The assembly consists
of a truck beam, axles, brake rods, and a protective shield.
The truck beam is the primary supporting member of the truck assembly. It pivots about the center where it is
attached to the lower end of the shock strut. Tow fittings are installed on each end of the beam for forward and aft
towing.
Two axles are installed in the truck beam, and are locked into place by retaining tow fittings. Four brakes are
installed on protective sleeves which surround the axles. Four wheels are mounted directly on the axles.
Brake rods link each brake to the shock strut. The rods transmit brake torque to the gear.
Wire bundles on the truck beam are protected from objects thrown from the tires by a protective shield clamped
underneath the truck beam.
IF 32-13
757 General Familiarization (7-2005)
ATA 32 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Side Strut and Down-lock Assembly
The side-strut and down-lock assembly locks the main gear into the extended position, and provides lateral
support of the gear. The assembly consists of the side strut and the down-lock, both of which are two-member
braces which fold during extension and retraction.
The side strut consists of an upper and a lower strut hinged in the center. The lower strut attaches to a rotating
swivel on the shock strut. The upper strut attaches to a universal swivel on the reaction link.
The down-lock consists of an upper and a lower link, also hinged in the center. The lower link attaches to the
hinge point of the side strut. The upper link attaches to a rotating spindle on the main gear jury support brace.
Two down-lock springs are located on the lower down-lock link to aid in locking the main gear into the extended
position. A down-lock pin is inserted through the two links to lock the gear for safety during ground operations.
IF 32-14
757 General Familiarization (7-2005)
ATA 32 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Reaction Link
The reaction link transmits lateral loads from the side strut to airplane structure. The outboard end of the link
attaches to the rear wing spar above the trunnion link. The inboard end attaches to a support link connected to a
body frame.
IF 32-15
757 General Familiarization (7-2005)
ATA 32 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR EXTENSION AND RETRACTION SYSTEM
Overview
The extension and retraction system does the control and operation for all landing gear movement. The extension
and retraction system includes the systems that follow:
1. Extension and retraction system for the main landing gear
2. Extension and retraction system for the nose landing gear
3. A control system
4. Alternate extension system
Extension and Retraction of the Nose Landing Gear
The extension and retraction system for the nose landing gear includes hydraulic actuators and valves. These
components give the sequence and operate the movement of the nose landing gear and the forward doors for the
nose landing gear.
Landing Gear Alternate Extension
The alternate extension system for the landing gear includes hydraulic actuators, valves, and a pump. These
components release the up-locks on the landing gear, open the doors, and let the landing gear extend when the
normal extension system does not operate. The alternate extension system is also used to open the landing gear
doors from the ground to get access to the wheel wells for maintenance.
757 General Familiarization (7-2005)
ATA 32 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR CONTROL SYSTEM
The control system for the landing gear controls the operation of the extension and retraction systems for the nose
and main landing gear. The control system includes a control lever, a selector valve, and a system of quadrants
and cables.
IF 32-16
757 General Familiarization (7-2005)
ATA 32 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR SYSTEM CONTROL COMPONENTS
Control Lever Module
The control lever is the mechanism which starts landing gear extension and retraction. The lever is contained in a
three-position module which is installed behind the pilot's center instrument panel. The lever moves in a guide
which has three detents (UP, OFF, and DN) and is spring-loaded to stay in the detent that is set.
A lever lock solenoid is included in the module for safety. This electrical solenoid moves to structurally stop
movement of the lever to the UP detent when a ground mode signal is received from the air/ground relay system.
This prevents retraction of the landing gear when the airplane is on the ground. A manual override rod
mechanically moves the solenoid when there is a failure in the system which causes the solenoid not to be
energized.
IF 32-17
757 General Familiarization (7-2005)
ATA 32 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Selector Valve
The selector valve sends hydraulic
draulic fluid from the left hydraulic system reservoir to the actuators and valves in the
extension and retraction systems for the main and nose landing gear. The valve has three positions which agree
with the positions of the control lever. The selector valve and the control lever are connected by quadrants and
cables. Movement of the control lever moves the selector valve into the position that agrees with the control lever.
The selector valve is installed in the wheel well for the right main landing gear. It can be found in the top, forward,
inboard corner of the wheel well.
IF 32-18
757 General Familiarization (7-2005)
ATA 32 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EXTENSION AND RETRACTION OF THE MAIN LANDING GEAR
The extension and retraction system for the main landing gear includes hydraulic actuators and valves. These
components give the sequence and operate the movement of the main landing gear and the doors for the main
landing gear.
The extension and retraction system for the main landing gear includes valves and actuators which retract and
extend the main landing gear and operate the doors for the main landing gear. The components of the system are
as follows:
1. Retract actuator
2. Door actuator
3. Down-lock actuator
4. Up-lock actuator
5. Door-operated landing gear sequence valve
6. Down-lock-operated door sequence valve
7. Up-lock-operated sequence valve
8. Truck positioner actuator
9. Truck positioner shuttle valve
10. Up-lock assembly
757 General Familiarization (7-2005)
ATA 32 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EXTENSION AND RETRACTION OF THE MAIN LANDING
ING GEAR COMPONENTS
Retract Actuator
The retract actuator is a hydraulic piston-type actuator that operates hydraulically in two directions. The actuator
snubs internally at each end of the piston travel. This actuator applies the force to retract and extend the landing
gear. The actuator is installed in the wing, outboard of the landing gear. The head end attaches to the beam that
holds the main landing gear. The rod end of the actuator attaches to the upper end of the shock strut.
IF 32-19
757 General Familiarization (7-2005)
ATA 32 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Actuator
The door actuator is also a hydraulic piston-type actuator that operates hydraulically in two directions. This
actuator snubs internally at each end of the piston travel. The door actuator applies the force to open and close the
door for the main landing gear. An internal mechanical lock in the head end of the actuator locks the door when it
is closed. Hydraulic pressure that is applied to the actuator to open the door releases the mechanical lock. The
actuator is installed near the center of the keel beam in the wheel well for the main landing gear. The head end
attaches to a fitting on the keel beam. The rod end attaches to the beam that holds the center hinge on the door
for the main landing gear.
IF 32-20
757 General Familiarization (7-2005)
ATA 32 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Down-lock Actuator
The down-lock actuator is also a hydraulic piston-type actuator that operates hydraulically in two directions. This
actuator does not snub internally. The down-lock actuator moves the down-lock links to let the landing gear retract
and also helps the down-lock springs lock the down-lock links when the landing gear extends. The actuator is
installed on the upper down-lock link. The head end attaches to the down-lock spindle, and the rod end attaches to
the upper down-lock link near the apex of the down-lock links.
IF 32-21
757 General Familiarization (7-2005)
ATA 32 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Up-lock Actuator
The up-lock actuator is a hydraulic piston-type
n-type actuator that operates hydraulically in one direction. This actuator
applies the force to move the up-lock hook to release the main landing gear and let it extend. The actuator is
installed on the up-lock assembly on the outboard fairing of the wheel well for the main landing gear.
IF 32-22
757 General Familiarization (7-2005)
ATA 32 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door-Operated Gear Sequence
The door-operated gear sequence valve is a three-position valve. The valve controls the flow of the hydraulic fluid
into the down-lock, up-lock, and retract actuators. This valve controls the sequence of landing gear movement in
relation to door position for the landing gear. The valve is connected to and operated by the door actuator. The
valve is installed above the door actuator on the keel beam in the wheel well for the main landing gear.
IF 32-23
757 General Familiarization (7-2005)
ATA 32 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Down-lock Operated Door Sequence Valve
The down-lock operated door sequence valve is a two-position valve. The valve controls the flow of hydraulic fluid
to the door actuator. This valve is hydraulically connected to the up-lock operated sequence valve. The valve's
primary function is to supply hydraulic pressure to the door actuator to close the door when the main landing gear
is down and locked. The valve is installed on the jury brace, which is attached to the rear wing spar, just outboard
of the upper down-lock link. Valve position is operated by a target on the upper down-lock link which engages a
roller on the valve when the landing gear is down and locked.
IF 32-24
757 General Familiarization (7-2005)
ATA 32 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Up-lock Operated Sequence Valve
The up-lock operated sequence valve is a two-position valve. It controls the flow of hydraulic fluid to the door
actuator. This valve is hydraulically connected to the down-lock operated door sequence valve. Its primary function
is to supply hydraulic pressure to the door actuator to open or close the door when the main landing gear is up
and locked. A restrictor check valve installed in the hydraulic fitting in the line that connects this valve and the
down-lock operated door sequence valve decreases the speed at which the door closes for safety. The up-lock
operated sequence valve is installed on the up-lock assembly on the outboard fairing of the wheel well for the
main landing gear. The valve is connected, but not directly, to the up-lock actuator and movement of the up-lock
actuator operates the valve position.
Up-lock Assembly
The up-lock assembly is a spring-loaded hook and linkage mechanism. This assembly locks the landing gear up in
the wheel well after it is retracted. The up-lock hook engages in the up-lock roller which is installed on the shock
strut. The mechanism is operated by the up-lock actuator. The up-lock assembly is installed on the outboard
fairing of the wheel well for the main landing gear.
IF 32-25
757 General Familiarization (7-2005)
ATA 32 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Truck Positioner Actuator
The truck positioner actuator is a hydraulic piston-type actuator. It operates hydraulically in one direction. This
actuator applies the force to put the truck assembly at an angle to permit clearance with the structure when the
landing gear moves to the extended or retracted position. The actuator is installed on the aft side of the shock
strut near the truck assembly. The head end attaches to the shock strut inner cylinder, and the rod end attaches to
the truck beam.
IF 32-26
757 General Familiarization (7-2005)
ATA 32 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Truck Positioner Shuttle Valve
The left truck positioner shuttle valve is a flow-limited control valve. Normally this valve permits free flow of
hydraulic fluid to the truck positioner actuator. If a low pressure input is received from the lines between the valve
and the actuator, fluid flow to the actuator is blocked. This is done with a hydraulic fuse.
On some aircraft the extension and retraction system for the right main landing gear has two hydraulic fuses
installed. There is one fuse installed in the landing gear up and one installed in the landing gear down control
lines. On these airplanes the right truck positioner shuttle valve fuse is not installed. If a low pressure input is
received the related fuse will stop flow to the full system. This will prevent the loss of the other airplane systems
supplied by the left hydraulic system because of a break or a leak.
IF 32-27
757 General Familiarization (7-2005)
ATA 32 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE GEAR EXTENSION AND RETRACTION COMPONENTS
Retract Actuator
The actuator is attached to the aft wall of the wheel well for the nose landing gear and to the shock strut trunnion.
The actuator hydraulically retracts to retract the landing gear. Hydraulic pressure is supplied to the retract actuator
only during landing gear retraction.
IF 32-28
757 General Familiarization (7-2005)
ATA 32 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Lock Actuator
The lock actuator hydraulically pushes the lock links into the over-center position. This locks the nose landing gear
in the extended or retracted position. The actuator also hydraulically pushes the lock links out of the over-center
position when the extension and retraction cycles start.
The actuator extends during landing gear extension to lock the lock link. The actuator retracts during landing gear
retraction to move the lock link from the locked position. The actuator is connected to the aft wall of the wheel
well and the aft lock link.
IF 32-29
757 General Familiarization (7-2005)
ATA 32 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Gear-Operated Sequence Valve
The gear-operated sequence valve makes sure that the nose landing gear is fully extended or retracted before the
doors close. The valve is mechanically connected to the drag strut assembly. It is installed on the aft wall of the
wheel well. The valve sends hydraulic fluid to the lock actuator and the door actuator.
IF 32-30
757 General Familiarization (7-2005)
ATA 32 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door-Operated Sequence Valve
The door-operated sequence valve is attached to the forward wall of the wheel well. The valve is connected to the
door actuator with rods and cranks. The valve makes sure that the doors are open before the landing gear extends
or retracts. Hydraulic fluid flows through the valve to the lock actuator and the gear sequence valve bypass valve.
Door Actuator
The door actuator opens and closes the forward doors for the nose landing gear during landing gear extension and
retraction. The actuator contains an internal mechanical lock to keep the doors in the closed position. The actuator
is installed on the forward wall of the wheel well. It is connected to the doors and the door-operated sequence
valve by rods and cranks.
Flow Control Valve
The flow control valve is a pressure-operated, two-position, in-line valve. It controls hydraulic fluid flow to and
from the extend pressure side of the door actuator. The valve also limits fluid flow to return from the door actuator
when the landing gear is down and locked and the doors are not fully closed. This function slows the speed at
which the door closes. This is for safety when the landing gear doors are opened or closed for ground
maintenance. At all other times, the fluid flow is not limited.
The flow control valve is installed in the hydraulic lines in the wheel well for the nose landing gear.
IF 32-31
757 General Familiarization (7-2005)
ATA 32 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Gear Sequence Valve Bypass Valve or Lock Unload Valve Module
The gear sequence valve bypass valve (lock unload valve module)
dule) receives a mechanical input from the drag strut.
This permits hydraulic fluid to flow to the retract actuator, around the door-operated sequence valve, as the nose
landing gear gets near the retracted position. The module is installed on the left wall of the wheel well.
IF 32-32
757 General Familiarization (7-2005)
ATA 32 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR ALTERNATE EXTENSION SYSTEM
Overview
Landing gear alternate extension is done by an electrical/hydraulic/mechanical system that operates
independently from the normal extension and retraction system. This system releases the locks on the landing gear
and doors when the normal hydraulic system does not work. This system is also used to open the landing gear
doors for maintenance procedures in or near the wheel well areas.
The alternate extension system has the components that follow:
1. Alternate extension switch
2. Power pack
3. Door lock release actuators
4. Alternate up-lock release actuators
5. Door safety valves
6. Hydraulic pressure switch.
The door ground operations system has these components (minus the alternate extension switch) plus those that
follow:
1. Door release interlock actuators
2. Door closed switches
3. Door locked switches
4. Door unsafe lights
5. Door ground control switches.
757 General Familiarization (7-2005)
ATA 32 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR ALTERNATE EXTENSION SYSTEM COMPONENTS
Alternate Extension Switch
The alternate extension switch is a momentary push-button switch.
itch. When it is pushed, electrical power is sent to
the power pack to start alternate extension. The switch is installed on the pilot's center instrument panel, P3-1,
immediately below the control lever for the landing gear.
An amber light that is part of the switch comes on when the alternate extension switch is pushed. The light stays
on when the alternate extend power pack operates and goes out when the alternate extend power pack shuts off.
IF 32-33
757 General Familiarization (7-2005)
ATA 32 - Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Power Pack
The power pack is an electrically-operated hydraulic pump. The power pack supplies a specified volume of
hydraulic fluid from the left system reservoir to the actuators and valves in the alternate extension system. The
power pack is installed on the keel beam in the right wheel well for the main landing gear. The location is forward
of the door actuator for the main landing gear.
IF 32-34
757 General Familiarization (7-2005)
ATA 32 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Lock Release Actuator
The door lock release actuator is a piston-type actuator that operates hydraulically in one direction. An internal
valve sends hydraulic fluid out of the actuator when the piston is fully extended. The valve (and on some airplanes,
the piston) is spring-loaded to go back to the normally retracted position when hydraulic pressure is removed. The
actuator supplies the force to mechanically release the internal lock in the door actuator.
A door lock release actuator is supplied for each landing gear. Each actuator for the main landing gear is installed
immediately aft of the door actuator, on the keel beam in the wheel well for the main landing gear. The actuator
for the nose landing gear is installed on the left wall of the wheel well for the nose landing gear.
Door Safety Valve
The door safety valve is a two-position valve which is installed in the door close line. It makes sure the landing
gear door does not accidentally close. The valve permits free flow of hydraulic fluid to the door actuator during
normal extension. During alternate extension, the valve is mechanically set in position and locked to do the
functions that follow:
1. Prevent hydraulic flow through the door close line that goes from the sequence valve to the door actuator
2. Open the line from the door actuator retract port to return
The valve is set in this position to make sure that no hydraulic fluid will be caught in the hydraulic line to cause
resistance when you try to open the door. It also prevents supply of hydraulic pressure to the door actuator to arm
the door to close without the door actually moved to the closed position.
There is one door safety valve for each landing gear. Each main landing gear valve is installed on the keel beam in
the wheel well for the main landing gear. The valve can be found immediately aft of the door actuator. The nose
landing gear valve is installed near the top of the forward wall of the wheel well for the nose landing gear.
Door Release Interlock Actuator
The door release interlock actuator has a hydraulic piston-type actuator and a solenoid valve assembly. This
assembly releases the lock from the door safety valve to permit the landing gear door to be closed. To operate, an
electrical input to the assembly and normal hydraulic pressure are necessary. The electrical signal causes the
solenoid valve to move and supply hydraulic pressure to the actuator. The actuator supplies the force to change the
position of the door safety valve to remove the lock.
There is one door release interlock actuator for each landing gear. Each main landing gear actuator is installed
immediately aft of the door actuator, on the keel beam in the wheel well for the main landing gear. The nose
landing gear actuator is installed at the forward end of the left wall of the wheel well for the nose landing gear.
757 General Familiarization (7-2005)
ATA 32 - Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-35
757 General Familiarization (7-2005)
ATA 32 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Alternate Up-lock Release Actuator
The alternate up-lock release actuator is a piston-type actuator that operates hydraulically in one direction. The
piston is spring-loaded to go back to the normally retracted position when hydraulic pressure is removed. The
main landing gear actuator supplies the force to move the up-lock hook to release the landing gear from the up
and locked position. The nose landing gear actuator supplies the force to move the lock links from the over-center
position. This releases the nose landing gear from the up and locked position.
There is one alternate up-lock release actuator for each landing gear. Each main landing gear actuator is installed
on the up-lock assembly. The up-lock assembly is found on the outboard fairing of the wheel well for the main
landing gear. The nose landing gear actuator is installed on the ceiling of the wheel well for the nose landing gear
above the landing gear trunnion.
IF 32-36
757 General Familiarization (7-2005)
ATA 32 - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Alternate Extension Hydraulic Shuttle Valve
The alternate-extension-hydraulic-shuttle valve is a two-position shuttle valve. It is installed in the landing gear
"up" line between the landing gear selector valve and the normal extension system
stem mechanisms. The valve position
during normal system operation permits the flow of fluid in the landing gear retraction "up" line. When alternate
extension is started, hydraulic pressure for alternate extension changes the position of the shuttle valve. This then
sends normal system "up" line fluid to return. This permits alternate extension to occur without the hydraulic
system pressure removed. This condition occurs when the position of the landing gear selector valve cannot be
changed from the "up to the "down" position. When the position of the valve is changed the higher (3,000) PSIG
normal system up pressure is removed from the normal system up-lock mechanisms. This permits the lower
(2250) PSIG alternate extension pressure to operate the landing gear doors and landing gear through the alternate
extension release mechanism.
IF 32-37
757 General Familiarization (7-2005)
ATA 32 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Hydraulic Pressure Switch
A pressure switch is installed in the hydraulic line for the alternate extension system. The switch is used to control
the time of operation for the power pack in the alternate extension system. After hydraulic pressure has extended
the door lock release and alternate up-lock release actuators, the landing gear and doors are released. When this
occurs the pressure in the system starts to increase. When the pressure switch receives an input of 1700-2200 PSI,
the pressure switch operates to remove electrical power from the power pack. On some airplanes, if the system
does not stop in 30 seconds, a LDG GEAR MONITOR message will show on EICAS.
The pressure switch is installed on a bracket at the top of the wheel well for the nose landing gear.
IF 32-38
757 General Familiarization (7-2005)
ATA 32 - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Closed and Locked Switches
The door closed switch is a mechanical plunger-type switch. It is used to supply an input when the landing gear
door is open. The switch operates directly by movement of the door. When the door opens, the plunger extends and
the switch turns on a red door not safe light. This light is used to tell persons that the door is not in a safe
condition. There is one switch for each landing gear door. The door switch for the main landing gear is installed
near the forward end of the outboard wall of the wheel well for the main landing gear. The door switch for the
nose landing gear is installed on the forward wall of the wheel well for the nose landing gear.
The door locked switch is a mechanical rotary-type switch. It is used to supply an input when the door safety valve
is turned and locked in the OFF (safety) position. The switch operates when its roller touches the safety valve cam.
When the safety valve is locked in the safety position, the door is hydraulically safe. The locked switch will then
make the red door not safe light go off. The door switch for the main landing gear is installed behind the door
safety valve in the wheel well for the main landing gear. The door switch for the nose landing gear is installed
below the door release interlock actuator in the wheel well for the nose landing gear.
IF 32-39
757 General Familiarization (7-2005)
ATA 32 - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-40
757 General Familiarization (7-2005)
ATA 32 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Ground Control Switches
The door ground control switches are used to open and close the landing gear doors to get access to the wheel
wells. There are two switches, the ALL DOORS OPEN ARM and the ALL DOORS OPEN switches. They are operated
at the same time to open all three landing gear doors together. These switches are momentary toggle switches
which supply electrical power to the power pack the same as the ALTN EXTENSION switch. There is a switch to
close the doors for the nose landing gear and a switch to close the doors for the two main landing gear. They
operate independently. With hydraulic power supplied and the control lever for the landing gear in DN, the MAIN or
NOSE GEAR DOOR CLOSE switch is operated to close the related doors. These switches are also momentary toggle
switches. They supply electrical power to the door-release-interlock-actuator solenoid.
The two (door open) switches and the main landing gear (door close) switch are found on the electrical service
panel for the MLG wheel wells, P72. This panel is found on the bottom of the fuselage immediately aft of the doors
for the main landing gear. The nose landing gear (door close) switch is found on the left equipment panel for the
nose landing gear, P63. This panel can be found on the shock strut for the nose landing gear.
757 General Familiarization (7-2005)
ATA 32 - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-41
757 General Familiarization (7-2005)
ATA 32 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WHEELS AND BRAKES
Overview
This section on wheels and brakes includes maintenance coverage on the following related topics:
1. Tires and wheels
2. Hydraulic brake system
3. Antiskid/auto-brake system
4. Parking brake system
5. Brake temperature monitoring system
Tires and Wheels
The tires and wheels support the airplane during ground operations. Four tires and wheels are installed on each
main gear and two on the nose gear.
Hydraulic Brake System
The hydraulic brake system is used to provide a means to slowdown and/or stop the airplane after landing
touchdown and during taxi operations. Brakes installed in each of the main landing gear wheels are actuated
hydraulically by manual brake pedal movement or automatically through the auto-brake system.
Antiskid/Auto-brake System
The antiskid system automatically releases the brakes to prevent skids or loss of control during braking.
The auto-brake system automatically applies the brakes after landing, to slow the airplane at a deceleration rate
selected by the pilots before landing.
Parking Brake System
The parking brake system enables the airplane brakes to be set for 8 hours minimum for parking.
Brake Temperature Monitoring System
The brake temperature monitoring system provides a means for the flight crew to monitor the temperature of the
brakes to help them avoid being overheated during braking operations.
757 General Familiarization (7-2005)
ATA 32 - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TIRES AND WHEELS
Overview
The airplane is supported by 10 wheel and tire assemblies; two on the nose landing gear and four on each main
landing gear truck.
Two spin brakes stop spinning of the nose wheels on retraction.
Main and Nose Gear Tires
The main gear tires are size H40 x 14.5-19 with a 22 ply rating. The tires have a speed rating of 225 mph.
The nose gear tires are size H31 x 13-12 with a 20 ply rating. The tires have a speed rating of 225 mph.
Refer to the servicing section of maintenance manual for proper main and nose gear tire inflation pressure.
IF 32-42
757 General Familiarization (7-2005)
ATA 32 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TIRES AND WHEELS COMPONENTS
Main Gear Wheels
The main landing gear wheels are made from forged aluminum and are a split wheel assembly for ease in
mounting tires. The inner surface of the rim of each wheel half is the sealing surface against which the tire bead
rests. The inner and outer wheel halves are mated and fastened together with 16 equally spaced tie bolts, secured
with washers and nuts. Leakage of air from the tire through the wheel half mating surfaces is prevented by a
packing mounted on the surface of the inner wheel half. Another packing mounted on the inner surface of the inner
wheel half seals the hub area of the wheel against dirt and moisture. Eight inserts (keys) installed over bosses in
the inner wheel half rotate the brake rotors as the wheel turns. Stainless steel heat shields mounted underneath
and between the inserts, keep excessive brake generated heat from the wheel and tire.
A tire inflation valve with an integral tire pressure gage is installed in each outer wheel segment. An over-inflation
valve is installed in each outer wheel half to relieve pressure at 375-450 PSIG to prevent excess pressure buildup.
Two tapered roller bearings are installed in the hub of each wheel. (4) Four thermal relief plugs in each inner
wheel half prevent tire blowouts due to excess brake heat. The plugs will melt to release tire pressure at
approximately 390°F (199°C).
757 General Familiarization (7-2005)
ATA 32 - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-43
757 General Familiarization (7-2005)
ATA 32 - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Nose Gear Wheels
The nose landing gear wheel assembly is made from forged aluminum and is a split wheel assembly for ease in
mounting tires. The inner surface of the rim of each wheel half is the sealing surface against which the tire bead
rests. The inner and outer wheel half assemblies are mated and fastened together with 10 equally spaced tie bolts,
secured with washers and nuts. Leakage of air from the tire through the wheel half mating surfaces is prevented
by a packing mounted on the surface of the inner wheel half. Retaining rings installed in the hub of the inner and
outer wheel halves hold the bearing grease seals and bearing cones in place. These seals retain the bearing
lubricant and keep out dirt and moisture.
A tire inflation valve with an integral tire pressure gage is installed in the outer wheel segment. An over-inflation
valve is installed in each outer wheel half to relieve at 375-450 PSIG to prevent excess pressure buildup.
Two tapered roller bearings are installed in the hub of each wheel.
IF 32-44
757 General Familiarization (7-2005)
ATA 32 - Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Nose Wheel Spin Brake
Two wheel spin brakes, attached to thee nose wheel well ceiling, stop the nose wheels from spinning after takeoff
and gear retraction. The brake pads are installed on the contact surfaces of the spin brakes to provide good
frictional contact.
Some airplanes have spin brake assemblies that have a spring arm with a brake wear pad made of a composite
material. An improved spin brake assembly that is installed on subsequent airplanes has a spring arm with
replaceable aluminum wear bars.
IF 32-45
757 General Familiarization (7-2005)
ATA 32 - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC BRAKE SYSTEM
Overview
The hydraulic brakes are controlled manually by applying foot pressure to the captain's and/or first officer's brake
pedals, or automatically by the auto-brake system. In the manual mode of operation, which is covered in this
section of the maintenance manual, the four brakes on the left gear are operated by the pilot's left pedal(s) and
the four brakes on the right gear are operated by the pilot's right pedal(s). The captain's and first officer's brake
pedals are joined together by linkage so that braking force will be the combined force applied to the pedals by both
pilots.
The hydraulic brake system includes the main landing gear brake assemblies (8 places), mechanical control
linkage from the pilot's foot pedals to the normal and alternate hydraulic metering valves (left and right sides),
hydraulic valves (accumulator isolation, alternate brake selector, normal and alternate metering), check valves,
wheel brake (de-spin) actuator, hydraulic lines and pilot's brake system indications. The brakes are multiple disk
rotor-stator type units that are activated hydraulically.
The pedal and quadrant linkage consists of the pilot's foot pedal mechanism (bell cranks, pushrods, quadrants)
which connects to the left and right metering valve modules by a series of cables running aft through the floor
beams. Two sets of cables are installed for system redundancy and they are guided by a series of pulleys mounted
on the body structure.
The brake hydraulic system is normally powered by the right hydraulic system. However, if right system pressure
is lost, the left system pressure is automatically selected by the alternate brake selector valve to operate the
brakes. A brake accumulator is provided in the event loss of both right and left hydraulic systems occurs. A limited
number of brake applications (5 or 6) can be accomplished using the accumulator source. A reserve source of
brake hydraulic power can be selected if other sources become inactive. The reserve source consists of one
electric motor-driven pump and a dedicated supply of hydraulic fluid in the right hydraulic system used for brakes.
Hydraulic system indications provided to the pilots include a brake system pressure gage, a low hydraulic brake
pressure light, and Engine Indicating and Crew Alerting System (EICAS) caution messages.
The parking brake system is included as a integral part of the hydraulic brake system and provides a means of
keeping brakes on without applying constant foot brake pedal pressure.
The antiskid system operates through the hydraulic brake system to prevent skids during braking. During normal
braking, pressure is unmodified by the antiskid valves. However, if one or more wheels enter a skid condition, the
antiskid valves will automatically reduce hydraulic pressure to the affected brakes. When the skid has been
averted, the antiskid valves will revert to normal, and allow normal braking to continue. The antiskid and autobrake systems are combined into one maintenance manual section.
757 General Familiarization (7-2005)
ATA 32 - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overview (Continued):
The brake temperature monitoring system is installed to monitor the temperature of each of the eight brakes and
aid the pilots in avoiding an overheat condition when the brakes are being operated. Brake temperature values are
displayed on the status mode of the EICAS system.
Gear retract braking, to stop the wheels from spinning prior to them entering the wheel well after take off, is
accomplished from the gear-up hydraulic pressure line. This pressure actuates the auto-brake (de-spin) actuators
on the left and right alternate brake metering valves which in turn actuates the brakes.
757 General Familiarization (7-2005)
ATA 32 - Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-46
757 General Familiarization (7-2005)
ATA 32 - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HYDRAULIC BRAKE SYSTEM COMPONENTS
Forward Brake Pedal and Quadrant Linkage
The captain's and first officer's brake pedal and quadrant linkages are similar and are integrated with the rudder
pedal assemblies. Rotation of the left and right rudder pedals about the heel pedals is transmitted through control
rods to the left and right lower bell cranks at the base of the rudder pedals. This motion is in turn transmitted
through fore-aft control rods to the forward bell cranks which are connected to cable quadrants. Maximum brake
pedal travel is approximately 16 degrees.
The left brake cables, normal and alternate, are driven by two quadrants which are connected by a short shaft to
the captain's left brake bell crank. The right brake cables, normal and alternate, are driven by two quadrants which
are connected by a short shaft to the first officer's right brake bell crank.
The captain's right brake bell crank is mounted on the shaft that is part of the captain's left brake bell crank and
quadrant assembly and is free to rotate on this shaft. This bell crank is connected by a transverse control rod to
the first officer's right brake bell crank and thus, to the right cable quadrant.
The first officer's left brake bell crank is in a similar manner, connected to the captain's left brake bell crank and
thus to the left cable system. This interconnecting mechanism allows the pilot's equal and simultaneous control.
757 General Familiarization (7-2005)
ATA 32 - Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-47
757 General Familiarization (7-2005)
ATA 32 - Page 67
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Brake Metering Valve Module
The left and right brake metering valve quadrant assemblies are installed on the forward bulkhead of the left and
right main gear wheel wells. The two sets of cables drive quadrants mounted on the normal and alternate brake
metering valve assemblies.
The normal and alternate metering valve assemblies are separated to ensure adequate hydraulic system separation
between the right (normal) and left (alternate) systems. The quadrants drive an actuating lever which in turn
depresses the valve slide plunger to meter hydraulic pressure to the brakes. A tension spring is installed on each
quadrant to provide brake pedal feel forces and return the actuating mechanism to the off position when pedals
are released. An adjustable stop bolt on the normal portion of the installation establishes the off position of both
quadrants.
Hydraulic feedback of brake pressure within the brake metering valve returns the valve slide to the off position to
release brake pressure when the actuating force on the valve slide is removed.
The quadrants are rigged with a gap between the actuating lever and the slide; the gap between the normal
actuating lever and the valve slide is less than the gap in the alternate unit to enable normal system activation
first. In the metering range, metered pressure is proportional to the force applied on the end of the slide.
A gear retraction braking actuator is installed on the alternate brake metering valves (left and right) to stop the
wheels from spinning prior to entering the wheel well after takeoff.
IF 32-48
757 General Familiarization (7-2005)
ATA 32 - Page 68
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 69
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Brake Accumulator
The brake accumulator provides a passive pressure source for approximately 5 to 6 brake applications when left
and right hydraulic systems are not available. The accumulator is also used to extend parking brake holding time
and to stabilize the brake system supply pressure during initial brake application and any subsequent antiskid
cycling.
The accumulator is automatically charged by the right hydraulic system and the charge is maintained by check
valve isolation of the supply line to prevent loss of pressure when the right system is depressurized.
The accumulator is located at approximate STA 1080 in the right wheel well on the main fore-aft bulkhead.
Alternate Brake Selector Valve (ABSV)
The ABSV is located on the forward STA 1040 bulkhead in the right wheel well.
The ABSV is used to isolate the left hydraulic system source from the alternate brake system. When the right
hydraulic (normal) system source is de-pressurized, the ABSV will cycle, in order to allow the left hydraulic
(alternate) system source to provide brake pressurization. This ABSV value cycling open takes place when control
pressure drops to approximately 48 percent or less of the normal pressurized system.
Accumulator Isolation Valve (AIV)
The AIV is located on the forward STA 1040 bulkhead in the right wheel well, just outboard of the ABSV.
The AIV is normally open. The AIV cycles closed to isolate the brake system accumulator when the left (alternate)
hydraulic system reaches 48 percent or more of accumulator pressure. The AIV cycles to the open position when
the left hydraulic (alternate) system drops to 48 percent or less of accumulator pressure to provide a backup
system (accumulator) of brake pressurization.
The AIV and ABSV utilize the same basic valve unit with various ports connected into the brake system in a
different manner to suit the different valve functions.
757 General Familiarization (7-2005)
ATA 32 - Page 70
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-49
757 General Familiarization (7-2005)
ATA 32 - Page 71
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Hydraulic Brake Assembly
The brake assembly is a conventional rotor-stator multiple disc unit. The brake housing is mounted on bushings
which ride on main landing gear axle sleeves. A brake rod (torque arm) attaches between the lower brake housing,
at a pin joint, and the yoke of the shock strut. The brake is thus allowed to rotate on the axle with corresponding
landing gear truck position changes. Lubrication fittings are provided for brake rod and axle bearing joints.
Brake pressure is applied hydraulically through six pistons which push against a pressure plate. The pressure
plate in turn slides axially against the stators and rotors to compress against the torque plate. When pressure is
released, brake adjuster springs pull the pressure plate back from contact with the rotor-stator stack.
The six brake adjusters automatically maintain the offset of the pressure plate from the rotor-stator stack. As the
brake disks wear, the adjustor compensates by allowing a pin and ball mechanism to deform a circular tube to
automatically hold the required offset.
Two wear pins are installed in each brake assembly which indicates when brakes require replacement. The pins
attach to the pressure plate and protrude through a brake housing bracket at grommet locations. The wear pins on
new brakes initially protrudes 1 3/8 inch past the grommets with brakes applied and when brakes need replacing
the protrusion will be zero.
A hydraulic hose disconnect fitting is attached at the top of the brake. The fitting provides a connection between
the brake hydraulic hose and the brake. The disconnect fitting will automatically close the hydraulic line on both
the hose and brake sides upon removal which will eliminate hydraulic line bleeding requirements after brake
replacement.
A bleed plug is installed at the top of the brake assembly, adjacent to the disconnect fitting. The plug can be used
to bleed air from the hydraulic brakes.
757 General Familiarization (7-2005)
ATA 32 - Page 72
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-50
757 General Familiarization (7-2005)
ATA 32 - Page 73
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Indication
A BRAKE SOURCE amber indication light is installed on the pilot's P1 panel. The BRAKE SOURCE indication
illuminates when both the right (normal) hydraulic system and left (alternate) hydraulic system are approximately
48 percent or less of normal system pressure.
The RESERVE BRAKES panel includes a selector switch with a white ON indicator light. The ON light indicates
when RESERVE BRAKES source is selected.
Two accumulator pressure gages are provided. One direct reading pressure gage is installed at the accumulator
charging valve location in the lower fuselage. The other indirect reading pressure gage is installed in the pilots P3
panel. The gages indicate the brake system nitrogen pre-charge pressure in the accumulator. The pressure signal
for the P3 panel gage originates at a transducer adjacent to accumulator and is sent to the brake pressure sensing
indicator and gage in the flight compartment.
In addition to the above indicators the brake system interfaces with the Engine Indicating and Crew Alerting
System (EICAS) to provide various caution messages. These messages include BRAKE SOURCE for low left and
right hydraulic system pressure.
IF 32-51
757 General Familiarization (7-2005)
ATA 32 - Page 74
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 75
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ANTISKID/AUTOBRAKE SYSTEM
Overview
The antiskid and the auto-brake systems are used in airplane braking. The antiskid system prevents wheel skids
by limiting pressure to the brakes. The auto-brake system provides automatic braking with the braking level
selected by the pilot.
Full power braking systems require precise metering of hydraulic pressure to the brakes. The antiskid system
works solely to prevent wheel skids. The auto-brake system works with the antiskid system to provide complete
automatic (feet off) braking.
Antiskid System
The system electronically compares actual airplane ground speed (from IRS) with wheel speed (from transducers)
for hydroplane and touchdown protection. This comparison provides a brake release signal to the antiskid valves.
The valves limit the pressure to the brakes. If the airplane speed drops below 7.5 knots, the system provides no
brake release signal.
Pressure reduction under other conditions is achieved by each wheel controlling its antiskid valve through the
control unit on the basis of wheel speed history. A rapid reduction in wheel speed is interpreted as a skid.
The system controls both normal and alternate brake systems through antiskid valves. Eight normal valves control
individual wheels. Four alternate valves control the lateral-pair wheels. The system also provides locked wheel and
hydroplane protection.
Four circuit breakers on the overhead circuit breaker panel P11 provide 28volts DC power to the system. When
power to the system fails, an amber ANTISKID fault light on pilots' overhead panel (P5) illuminates. The Engine
Indication and Crew Alerting System (EICAS) display on pilots' center instrument panel will show the ANTISKID
advisory message.
The system requires the following airplane interface inputs:
1. Inertial Reference System (IRS) speed data
2. Landing gear position (down and locked or not down and locked) signal
3. Parking brake valve position signal
757 General Familiarization (7-2005)
ATA 32 - Page 76
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-52
757 General Familiarization (7-2005)
ATA 32 - Page 77
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Autobrake system
The system applies brakes automatically at touchdown when the average wheel speed reaches 60 knots. It
maintains a constant pilot-selected deceleration level throughout the landing roll. The only pilot effort required is
to select the desired level prior to each landing. There is no interference with normal antiskid system operation.
Full manual braking is always available.
An AUTOBRAKES selector switch on pilots' center instrument panel (P1-3) provides pilots' choice of five levels (1,
2, 3, 4 and MAX AUTO) of automatic braking. The system uses the main wheels average speed signal for spin up
logic and the airplane speed data for deceleration control. The system compares the airplane speed with the
selected level of braking to generate a pressure command to the control servo valve. The valve provides the
commanded pressure via the normal antiskid valves to the brakes. The system can be armed only when there is no
faults in both antiskid and auto-brake systems.
The AUTOBRKS/ANTISKID TEST/IND 1 and AUTOBRKS/ANTISKID TEST/IND 2 circuit breakers on panel P11
simultaneously provide 28-volts DC power to the system. When power fails, an AUTOBRAKES light on pilots' center
instrument panel illuminates. The EICAS display will show the AUTOBRAKES advisory message. Both AUTOBK
ANTISKID TEST IND 1 and AUTOBK ANTISKID TEST IND 2 circuit breakers provide power for testing the system.
The system requires the following airplane interface inputs:
1. An operational (no fault) antiskid system.
2. IRS data
3. Landing gear air/ground signal
4. Thrust levers position (advanced or not advanced) signal
5. Spoiler handle position (fully extended or not fully extended) signal
An antiskid/auto-brake control unit provides all the monitoring and control, including arming, disarming and self
test of the system.
The antiskid/auto-brake system consists of the following components (quantity shown in bracket):
1. Transducer (8)
2. Antiskid module, normal system (2)
3. Antiskid module, alternate system (2)
4. ANTISKID fault light on P5 (1)
5. Antiskid shuttle valve module (2)
6. ANTISKID ON/OFF switch on P5 (1)
7. Antiskid/autobrake control unit, M102 (1)
8. Autobrake selector switch on P1-3 (1)
9. Autobrake module (1)
10. Autobrake shuttle valve assembly (2)
11. AUTOBRAKES light on P1-3 (1)
757 General Familiarization (7-2005)
ATA 32 - Page 78
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 32 - Page 79
IF 32-53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ANTISKID/AUTOBRAKE SYSTEM COMPONENTS
Antiskid/Autobrake Crew Control Panels and Annunciators
ANTISKID ON/OFF Switch
An ANTISKID ON/OFF switch on panel P5 provides 28-volts DC power to the control unit when it is in the ON
position. The switch removes power to the unit when it is in the OFF position.
An amber light with OFF legend on the switch comes on to warn the pilots that the antiskid system has been
turned off. With the ANTISKID ON/OFF switch in the OFF position, the ANTISKID OFF advisory message appears
on EICAS display.
The ANTISKID OFF message will inhibit the ANTISKID/AUTOBRK maintenance message unless
ANTISKID/AUTOBRK has existed 10 seconds prior to the display of ANTISKID OFF.
ANTISKID Fault Light
An amber ANTISKID fault light on panel P5 illuminates to signal antiskid system fault when faults exists. At the
same time the ANTISKID advisory message appears on the EICAS display.
757 General Familiarization (7-2005)
ATA 32 - Page 80
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-54
757 General Familiarization (7-2005)
ATA 32 - Page 81
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Autobrake Selector Switch
The switch, located on pilots center instrument panel P1-3, is a rotary, magnetic-latching seven position switch, or,
on airplanes with with RTO, an eight position switch. The switch performs the following functions:
1.
2.
3.
4.
Provide 28-volts DC power to the antiskid/autobrake control unit
Select 1, 2, 3, 4 or MAX AUTO airplane deceleration level
Arm or disarm the system
Turn on or turn off the AUTOBRAKES light
AUTOBRAKES Light
The AUTOBRAKES light is an amber light located near the selector switch. The light, controlled by the selector
switch, comes on when:
1. The switch is at DISARM position
2. The switch is at OFF position and the autobrake module solenoid valve output pressure switch shows
presence of high pressure
3. The switch is at 1, 2, 3, 4 or MAX AUTO position and system fault is detected
When the selector switch is at 1, 2, 3, 4 or MAX AUTO, the light illuminates for a moment as the switch is moved
through the DISARM. The light then goes out when the unit confirms that arming requirements are met. When the
system disarms, the light illuminates until the switch is placed to OFF or the system is rearmed.
IF 32-55
757 General Familiarization (7-2005)
ATA 32 - Page 82
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Antiskid/Autobrake Control Unit
The control unit (M102) compares each wheel speed with the IRS ground speed for touchdown and hydroplane
protection. A change in speed causes a change in control signal to increase or decrease hydraulic pressure to the
brakes.
The unit may be located in either the E6 rack of the aft equipment center, or the E5 rack of the main equipment
center.
The unit is an LRU and circuit cards are also LRUs. Either unit or cards can be replaced as required in
maintenance.
IF 32-56
757 General Familiarization (7-2005)
ATA 32 - Page 83
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Antiskid Module (Normal)
Two 4-valve antiskid modules are used in the normal brake system. The modules are located in the left and right
main wheel well fairing structure, near the wheel well ceiling, just aft of the left and right hydraulic system
reservoirs. Each module contains four identical antiskid valves, four hydraulic fuses, a shutoff valve, two inlet
filters, a check valve, a restrictor, and a housing with associated parts. The module provides individual wheel
control to each main gear. Each module is an LRU and the fuses, shutoff valve, and inlet filters are separate
component LRUs. The valves and filters can be removed for inspection without disconnecting hydraulic lines.
Antiskid Module (Alternate)
Two 2-valve antiskid modules are used in the alternate brake system. The modules are located at the keel beam
web on a mounting support bracket in left and right main wheel wells. Each module contains two identical antiskid
valves, two hydraulic fuses, one inlet filter, a check valve and a housing with associated parts. The module
provides laterally paired wheel control to each main gear. Except for configuration differences and quantity of
components and wheel control, the alternate 2-valve module functions similarly to the normal 4-valve module.
IF 32-57
757 General Familiarization (7-2005)
ATA 32 - Page 84
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Antiskid Shuttle Valves
The module contains four identical valves. Each valve is independent and pressure operated. Two modules are
located one each at left and right main wheel well ceiling, just forward of the main gear up-lock assembly. Each
valve shuttles pressure between normal and alternate systems.
The valve consists of an LRU valve assembly and an LRU filter. The valve has a manual override plug feature. The
filter prevents system from contamination.
In the event of shuttle valve failure, the slide plug on the face of the valve is removed. A flight dispatch plug (a flyaway ground maintenance tool consisting of a small threaded plug) is installed in its place. The installed plug
forces the slide in the valve to shift, thus blocking the normal port and opening the alternate port. This condition
remains with the plug installed. The plug has no moving parts and is equipped with a ring to allow an indicator tag
to be tied to it while being used.
IF 32-58
757 General Familiarization (7-2005)
ATA 32 - Page 85
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Antiskid Wheel Speed Transducers
The transducer, a speed sensing device on each main gear wheel, contains only one moving part, a rotor which
rotates inside a fixed stator. The stator attaches to a support inside the main wheel axle. The voltage created,
related to the speed of the wheel, provides the control unit with wheel speed data.
The rotor, thru a four-arm dog rigidly attached on the rotor shaft, couples to the transducer drive in the hubcap.
The drive, consisting of a bellows-type coupling and related mounting hardware inside the hubcap, turns the rotor
when the wheel rotates. The dog/bellows coupling allows removal of the wheel and hubcap without disassembly of
a bolted joint. Both transducer and drive are LRUs.
IF 32-59
757 General Familiarization (7-2005)
ATA 32 - Page 86
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Autobrake Module
The autobrake module is connected to the normal brake lines, and is located on the forward bulkhead of the left
main gear wheel well. The module contains an electric hydraulic pressure control servo valve (EHSV) (hereinafter
called the pressure control valve), an upstream three-way solenoid shutoff valve (hereinafter called the solenoid
valve), and two pressure switches, located one each at the outputs of the solenoid valve and the servo valve. The
module is an LRU as the valves and switches are LRUs. Solenoid valve, pressure control valve and pressure
switches can be replaced without removing the module from the airplane.
The module develops brake pressure in response to selected deceleration for all required autobrake functions. The
solenoid valve provides on-off control of hydraulic power to the valve module, and the pressure control valve
controls output pressure from the module as commanded by the control unit. Pressure switches on the module
monitor the pressure outputs from the solenoid valve and the pressure control valve and provide the logic to the
control unit.
IF 32-60
757 General Familiarization (7-2005)
ATA 32 - Page 87
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Autobrake Shuttle Valve Assembly
The assembly consists of a valve and a pressure sensing switch. The basic three-way valve allows brakes to work
by either a manual or an autobrake system. Two valve assemblies are located one each on left and right wheel well
forward bulkhead. Both valves and switches are LRUs. Pressure switch can be replaced without removing valve
from the airplane.
The pressure switch, connected to the normal input port, checks pressure downstream of the normal brake
metering valves. When manual braking effort exceeds 750 PSI on either the left or right pedal, the switch opens to
provide an input to the control unit to disarm the system.
IF 32-61
757 General Familiarization (7-2005)
ATA 32 - Page 88
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE WHEEL STEERING SYSTEM
Overview
The nose wheel steering system is hydraulically powered by the 'gear down' lines of the left hydraulic system. As a
backup hydraulic power source, the right system can input hydraulic power through a power transfer unit interface
when the left system is unavailable. Hydraulic steering pressure is removed when the landing gear handle is not in
the down position.
The nose wheel steering system is controlled by two means:
1. The captain's sidewall mounted tiller can be used to turn the nose wheels 65 degrees left or right
of center for ground maneuvering.
2. The rudder pedals provide limited steering control of 5 degrees left or right of center, when the
nose gear is compressed, for takeoff and landing roll steering.
The tiller is connected by control cables to a summing bar mechanism on the nose landing gear strut which
functions to input mechanical signals through a metering valve. The metering valve then provides hydraulic power
to two steering actuators which connect to a steering collar. The steering collar in turn connects to the lower
steerable position of the shock strut by torsion links.
Any relative movement of the control cables from either the tiller or rudder pedals will displace the summing bar
which mechanically signals a deviation between the commanded and actual wheel position. This deviation signal
moves the metering valve slide to pressurize the steering actuators in the proper direction to correct the condition.
When the lower steerable strut position matches the commanded (tiller or rudder pedal) position, the summing bar
is again centered which also centers the metering valve.
Rudder pedal steering is provided by an interconnect mechanism which working along with a piston position
system and its cables, drives an eccentric arm to engage or disengage rudder pedal input into the nose wheel
steering system. This mechanism also isolates the rudder pedals from the tiller controlled steering system when
the nose gear is unloaded and the shock strut is extended.
The two system centering springs keep the nose gear centered when there is no control input force. The springs
also return the steering mechanism to neutral control position whenever the control input force is removed from
the tiller.
A towing lever, on the steering metering valve module, permits towing without physically disconnecting the shock
strut torsion links.
757 General Familiarization (7-2005)
ATA 32 - Page 89
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overview (Continued):
The nose gear steering system is automatically deactivated when the nose gear is retracted by removing hydraulic
pressure from the system.
Cams internal to the nose gear shock strut are provided to hold the gear centered if hydraulic power is not
available. The cam is disengaged upon landing when the nose gear strut is compressed 3.0 inches or more.
IF 32-62
757 General Familiarization (7-2005)
ATA 32 - Page 90
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-63
757 General Familiarization (7-2005)
ATA 32 - Page 91
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOSE WHEEL STEEERING SYSTEM COMPONENTS
Tiller, Gearbox, and Torque Limiter
The steering tiller allows the captain to manually steer the airplane.
The tiller handle can be turned 360 degrees (maximum) in either direction from center. A geared nose wheel
position indicator is installed at the base of the tiller handle to show the angle of the nose wheels from center.
The cable loop is protected from high input forces from the pilot by a torque limiter that limits tiller lever input to
40 pounds maximum. The torque limiter transmits torque through spring loaded rollers that engage a cam
mounted to the forward cable drum. If the system cable load exceeds the torque limit, the rollers will come out of
the detents in the cam. The rollers will only re-engage when the tiller is restored to its proper relationship with the
forward cable drum.
IF 32-64
757 General Familiarization (7-2005)
ATA 32 - Page 92
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 93
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rudder Pedal Steering Interconnect Mechanism
The rudder pedal steering interconnect mechanism consists of a rudder pedal steering crank, a clutch arm, a cable
drum, an eccentric arm and a rudder pedal steering quadrant.
This mechanism connects the rudder pedals to the nose wheel steering system and is engaged by the piston
position system when the nose gear is compressed. Conversely it is disengaged when the nose gear shock strut is
extended at lift off.
The nose gear shock strut compression movement is transmitted by piston position linkage and a separate cable
loop to move the cable drum and reposition the eccentric arm to allow the clutch cam stops to contact the steering
crank. In this position, any movement of the rudder pedal is transmitted from the steering crank to the rudder
pedal steering quadrant through the centering spring. This quadrant is connected to the nose wheel steering cables
and is free to move the main steering system cables whenever nose wheel steering tiller is used, or drives the
cables when positioned by the rudder pedal steering mechanism.
When the nose gear shock strut is extended (on lift-off), the piston position system rotates the cable drum to move
the stops mounted on the eccentric arm to disengage the clutch arm from the steering arm stops. This disconnects
the rudder pedal system from the steering system.
757 General Familiarization (7-2005)
ATA 32 - Page 94
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-65
757 General Familiarization (7-2005)
ATA 32 - Page 95
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Spring Cartridge and Piston Position Quadrant
The spring cartridge attaches between the upper torsion link and the shock strutt mounted piston position quadrant.
The spring cartridge moves the piston position quadrant into either of two positions depending on whether the gear
is compressed or extended.
The piston position quadrant moves through an angular displacement of approximately 23 degrees between preset
steps in the quadrant attach bracket on the shock strut.
When the gear is extended the spring cartridge moves downward to rotate the quadrant clockwise 23 degrees. This
action disengages rudder pedal steering. When the gear is compressed the spring cartridge moves upward to rotate
23 degrees. This action engages the rudder pedal steering system through the interconnect mechanism.
IF 32-66
757 General Familiarization (7-2005)
ATA 32 - Page 96
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Summing Mechanism and Broken Cable Compensator
The summing mechanism consists of two lever assemblies, left and right, a valve input lever, a splined shaft and a
valve input rod.
Under normal conditions, any steering command is transmitted by the steering cable (NWS) to the lever assembly
and associated force link which drives the input roller and cam. The cam and valve input lever are splined on the
common shaft and as a result the valve input rod is driven by the input lever.
The broken cable compensator consists of two force links, left and right, and a roller and a cam which work in
conjunction with the summing mechanism.
The broken cable compensator is connected to the steering valve input lever and prevents cable tension to cause a
sustained input to the steering system if the left or right cable looses tension. As long as tension exists in both
cables the force links keep the roller engaged in the mating cam detent. When tension is lost in one cable the force
links disengage the roller from the cam detent, losing any sustained unbalanced steering force capability.
IF 32-67
757 General Familiarization (7-2005)
ATA 32 - Page 97
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Steering Metering Valve Module and Actuators
The steering metering valve is attached to the upper steering plate on the nose gear shock strut and its function is
to control the amount of hydraulic fluid provided to the two steering actuators.
The metering valve is mechanically controlled by linkage from the summing bar which detects any error angle
between the steering collar on the nose gear shock strut and tiller position.
Valve centering springs ensure that the valve spool remains centered when tension is lost in the control cables.
The swivel valves channel hydraulic fluid to the actuator. The swivel valves function to prevent one actuator from
working against the other when it passes over center at a steering angle of greater than 30 degrees.
A shimmy damping valve, called a dynamic load damper (DLD) is an integral part of the steering metering valve.
The valve is a combination of orifices, oil volumes, and spring which provides a dynamic filter circuit for spool
motion to reject low frequency actuator pressure variations, but to recognize high frequency pressure variations
which will occur if any shimmy conditions exists. The valve will then meter damping flow across the actuators.
Bleed orifices from supply to the dynamic load damper (DLD) provide continuous purging flow and fill the
compensator with fluid when the gear is extended.
A towing lever is provided on top of the valve package. Pulling the towing lever will first move the shut-off valve to
block supply pressure to the valve package and then shift the DLD spool to port fluid across the steering actuators.
The lever is spring loaded and must be held in towing position with a pin type tool.
757 General Familiarization (7-2005)
ATA 32 - Page 98
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-68
757 General Familiarization (7-2005)
ATA 32 - Page 99
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Steering Collar and Torsion Links
The nose gear steering collar is mounted on the shock strut outer cylinder and is retained by the actuator lower
trunnion support plate.
Both steering actuator rod ends are connected to the steering collar, which in turn is connected to the upper end of
the upper torsion link. When force is applied to the steering collar by either or both actuators, the collar rotates on
the shock strut and also imparts motion through the torsion links to turn the shock strut inner cylinder and nose
gear wheels.
IF 32-69
757 General Familiarization (7-2005)
ATA 32 - Page 100
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 32 - Page 101
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AIR/GROUND RELAY SYSTEM
Overview
The air/ground relays provide various airplane systems control in two independent air ground sensing systems
(system 1 and system 2). Air/ground sensing is provided by inputs from the nose gear not compressed sensors,
main gear truck tilt sensors, and truck positioner actuator hydraulic inlet pressure switches. The relays energize or
de-energize to provide equipment functions in ground or air mode. On most functions, either system 1 or 2 is
used. However, critical airplane function may use both systems for air/ground sensing.
When the airplane is on the ground (main gear trucks not tilted, nose gear shock strut compressed, and truck
positioner actuator hydraulic inlet pressure switches open) or in the air (main gear trucks tilted, nose gear shock
strut extended, and inlet pressure switches closed), the proximity sensors provide air/ground signal to a proximity
switch electronics unit (PSEU). The PSEU converts sensor input into a signal output to the relay. The relay
energizes or de-energizes to provide ground or air mode condition.
Air/ground system 1 or air/ground system 2 each consists of a number of relays, nose gear not compressed
sensor, and left and right main gear truck tilt sensors. All relays are 4-pole, double-throw, rated at 2 or 10
amperes, and are located in P33, P36 and P37 relay panels in the main E/E compartment.
757 General Familiarization (7-2005)
ATA 32 - Page 102
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AIR/GROUND RELAY SYSTEM COMPONENTS
Air/Ground Relays
All system 1 and system 2 air/ground control relays aree 4-pole, double-throw, 2 or 10 amp-rated, hermetically
sealed relays with pin-type terminals for electrical plug connections. The relays energize or de-energize to provide
a ground or air mode condition to the airplane. Except for relay K144 located in P33 panel, all system 1 relays are
located in P36 panel. System 2 relays are located in P37 panel.
IF 32-70
757 General Familiarization (7-2005)
ATA 32 - Page 103
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Nose Gear Not Compressed Sensors
A proximity sensor S10067 for system 1 and a proximity sensor S10068 for system 2 are located on the left and
right side respectively of the nose wheel steering collar, on a mounting plate. The sensors provide air/ground
signal to the PSEU. The PSEU converts the signal for energizing or de-energizing the relays. The same signal also
goes to the left and right EICAS computers for air/ground output comparison. If the nose gear not compressed
signal outputs from systems 1 and 2 do not agree, the computers provide a NOSE A/G DISAGREE message on the
EICAS status/maintenance display.
AIRPLANES WITH S242N701-1001 EICAS COMPUTERS INSTALLED; if either of the nose air/ground sensors
fail in the air mode while on the ground, then the message NOSE A/G SYS will show on the EICAS
status/maintenance display.
Main Gear Truck Tilt Sensors
Two proximity sensors (S10062, LH, S10060 RH) for system 1, and two proximity sensors (S10064 LH, S10059 RH)
for system 2, are located on the outboard side of the truck beam, aft of the main gear shock strut. The sensors
provide air/ground signal to the PSEU. The PSEU converts the signal for energizing or de-energizing the relays.
One of the on ground signals from system 1 and system 2 also goes to the left and right EICAS computers for
air/ground output comparison. If the on-ground signal outputs from systems 1 and 2 do not agree, the computers
provide an AIR/GND DISAGREE message on the EICAS status/maintenance display.
AIRPLANES WITH S242N701-1001 EICAS COMPUTERS INSTALLED; if either the system 1 or system 2
air/ground logic fails in the air mode while on the ground, then the message AIR/GND SYS will show on the
EICAS status/maintenance display.
757 General Familiarization (7-2005)
ATA 32 - Page 104
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-71
757 General Familiarization (7-2005)
ATA 32 - Page 105
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR POSITION INDICATING AND WARNING SYSTEM
Overview
The system shows status of landing gear and gear doors to the pilots. Two independent landing gear subsystems
are used to provide redundant control to indication lights and EICAS (Engine Indication and Crew Alerting System)
messages.
The system also provides inputs to the aural warning system, landing configuration warning module.
The system consists of indication lights on the pilot’s center instrument panel P3 and EICAS messages on the
upper EICAS display. The system receives gear position signals from the PSEU (Proximity Switch Electronic Unit)
to turn on or off the indicator lights and EICAS messages. Proximity sensors which provide landing gear and door
position signals to the PSEU are part of the proximity switch system.
757 General Familiarization (7-2005)
ATA 32 - Page 106
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LANDING GEAR POSITION INDICATING AND WARNING SYSTEM COMPONENTS
Indicator Lights
Five landing gear indication lights, one amber GEAR disagreement light, one amber DOORS light, and three green
NOSE, LEFT, and RIGHT lights are located on panel P3. Indication lights have no delay before being turned on.
Landing gear indicator lights have dual light bulbs with independent power sources to reduce the possibility of
indicator lights failing to illuminate as a result of power failure.
The amber GEAR light illuminates when the gear is in transit, or when the gear position does not agree with the
gear lever position. The GEAR light will be off when all gears are down and locked and control lever is down, or all
gears are up and locked and control lever is not down.
The amber DOORS light illuminates when any landing gear door is not completely closed and locked. When all the
landing gear doors are completely closed and locked the DOORS light will be off.
The green NOSE light illuminates when the nose gear is down and locked. When the nose landing gear is not down
and locked the NOSE light is off.
The green LEFT light illuminates when the left main gear is down and locked. When the left main landing gear is
not down and locked the LEFT light is off.
The green RIGHT light illuminates when the right main gear is down and locked. When the right main landing gear
is not down and locked the RIGHT light is off.
EICAS Display
The landing gear indication for EICAS is on the EICAS display on the pilot's center instrument panel P3. The EICAS
messages for landing gear position are as follows:
One amber GEAR DISAGREE
One amber GEAR DOORS
One white LDG GEAR MONITOR
One white PSEU BITE
During normal operation system 1 and 2 provide redundant control of EICAS. All amber EICAS messages have time
delay varying from 25 to 35 seconds. The white EICAS messages have time delays varying from 2 to 10 second
delays.
The amber GEAR DISAGREE message is displayed on EICAS when the gear is not in selected gear handle position
within 25 seconds.
757 General Familiarization (7-2005)
ATA 32 - Page 107
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EICAS Display (Continued):
The amber GEAR DOORS message is displayed on EICAS when any landing gear door is not closed and locked
within 35 seconds after moving the gear handle.
The white LDG GEAR MONITOR or PSEU BITE message appears on EICAS when separate and parallel inputs from
systems 1 and 2 to the EICAS computer disagree. LDG GEAR MONITOR message appears when disagreement is
indicated with gear commanded down, while PSEU BITE message appears when disagreement is indicated with
gear commanded up or door closed inputs do not agree between system 1 and system 2.
These maintenance level messages help fault isolation for the LDG GEAR MONITOR status level message. Each of
the messages are stored in EICAS non-volatile memory (NVM). The messages are as follows:
1. GEAR DISAGREE System 1 and 2 landing gear disagree indicators disagree in down position
2. NOSE GEAR DOWN System 1 and 2 nose gear down indications disagree
3. NOSE GEAR LOCKED System 1 and 2 nose gear locked indications disagree
4. ALL GEAR DOWN Landing gear system indicates all gear down and locked to the landing configuration
warning system when nose gear is up
5. GEAR LEVER System 1 and 2 gear lever indications disagree
6. L GEAR DOWN System 1 and 2 left main gear down and locked indications disagree
7. R GEAR DOWN EAL 501-515; INT 005, 014; MON 001-003; LESS EAL, INT, MON AIRPLANES POST-SB 3225; System 1 and 2 right main gear down and locked indications disagree
8. R GEAR DOWN EAL 516-999; INT 101-999; MON 004-999; PLUS EAL, INT, MON AIRPLANES POST-SB 3225; System 1 and 2 right main gear down and locked indications disagree or alternate extend power pack
fails to produce 2200 PSI within 30 seconds during alternate extension or on-ground door operation
9. GEAR DOORS System 1 and 2 gear door indications disagree
10. A LDG GEAR MONITOR status message will appear whenever one or more of the above maintenance
messages are enabled. The maintenance messages have time delays varying from 2 to 60 seconds
NOTE:
The LDG MONITOR status message is not in NVM but is displayed when one or more of
the NVM maintenance messages are displayed.
757 General Familiarization (7-2005)
ATA 32 - Page 108
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 32-72
757 General Familiarization (7-2005)
ATA 32 - Page 109
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Landing Gear Down and Locked Sensors
Sensors S10061 (left) and S10057 (right) for system 1, and S10074 (left) and S10070 (right) for system 2, are
located on the main gear down lock links at the apex. The sensors provide logic signals for the GEAR, LEFT, and
RIGHT lights and EICAS messages GEAR DISAGREE and LDG GEAR MONITOR.
IF 32-73
757 General Familiarization (7-2005)
ATA 32 - Page 110
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Landing Gear Up and Locked Sensor
Sensors S10240 (left) and S10239 (right) for system 1, and S10064 (left) and S10059 (right) for system 2, are
located on the up-lock assembly in the outboard fairing of the main wheel well. The sensors provide logic signals
for the GEAR light and EICAS messages GEAR DISAGREE and PSEU BITE.
IF 32-74
757 General Familiarization (7-2005)
ATA 32 - Page 111
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Main Landing Gear Door Closed Sensors
Sensors S10242 (left) and S10241 (right) for system 1, and S10076 (left) and S10072 (right) for system 2, are
located on the forward outboard main wheel well gear door opening. The sensors provide logic signals for EICAS
messages GEAR DOORS and PSEU BITE.
IF 32-75
757 General Familiarization (7-2005)
ATA 32 - Page 112
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Nose Gear Locked Sensors
Sensors S10065 for system 1 and S10078 for system 2 are located on the nose gear lock link at the apex. The
sensors provide logic signals for NOSE and GEAR lights and EICAS messages GEAR DISAGREE and PSEU BITE.
Nose Gear Up Position Sensors
Sensors S10238 for system 1 and S10077 for system 2 are located above the upper drag brace on the nose wheel
well ceiling. The sensors provide logic signals for GEAR light and EICAS messages GEAR DISAGREE and LDG GEAR
MONITOR.
Nose Gear Down Position Sensors
Sensors S10066 for system 1 and S10079 for system 2 are located above the upper drag brace on nose wheel well
ceiling. The sensors provide logic signals for the GEAR light and EICAS messages GEAR DISAGREE and LDG GEAR
MONITOR.
Nose Gear Door Closed Sensors
Sensors S10243 for system 1 and S10081 for system 2 are located near the center of the nose gear wheel well sill.
The sensors provide logic signals for the DOORS light and EICAS message GEAR DOORS
IF 32-76
757 General Familiarization (7-2005)
ATA 32 - Page 113
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PROXIMITY SWITCH SYSTEM
Overview
The proximity switch system consists of a proximity switch electronics unit (PSEU) and numerous proximity
sensors mounted throughout the airplane. The sensors provide position sensing for landing gear, doors and thrust
reversers, and basically involves the following subsystems:
1. Landing gear systems (systems 1 and 2)
2. Cargo door control system
3. Door system
4. Thrust reverser auto re-stow (left and right)
5. Thrust reverser indication (left and right)
The sensors sense the proximity of targets installed and provide position signals to the PSEU. Discrete operational
inputs to the PSEU include mechanical switches which represent system command or position sensing signal. The
PSEU converts all these inputs into output signals for operation and control of relays, lamps and other electronics.
The system incorporates built-in-test equipment (BITE) in the PSEU to provide both in-flight testing and on-ground
system trouble shooting. BITE tests on a malfunctioning PSEU produce a fault code which indicates the faulty
sensor or target state, card, or absence of section operating voltage.
757 General Familiarization (7-2005)
ATA 32 - Page 114
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Proximity Switch System Components
Proximity Switch Electronics Unit (PSEU)
The PSEU (M162), located on the E3-4 shelf in the main equipment center.
The PSEU chassis accommodates 12 active cards, 3 spare cards, and a BITE module.
All PSEU cards are line replaceable units (LRUs) and are accessed through hinged panels on the front of the PSEU.
It is not necessary to remove the PSEU from its shelf to access the cards.
IF 32-77
757 General Familiarization (7-2005)
ATA 32 - Page 115
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Proximity Sensors
The sensors provide the position sensing inputs to the PSEU. They are distributed to these systems:
1.
2.
3.
4.
Door System
Thrust Reverser System
Landing Gear System
Cargo Door Control
The sensor, a two-wire device connected to the PSEU proximity card, is a magnetic field-producing coil-core
combination. The sensor is contained in a non-magnetic stainless steel case. When a steel (magnet) target is
brought near or moved away from the sensor face, the sensor inductance increases or decreases respectively.
Two types of sensor are used, round or rectangular. The round sensor is used for applications that do not allow
the installation of the rectangular sensors due to physical constraints. Both sensors sense the proximity or distance
of a steel (magnetic) target to its sensitive surface. The sensor actuation gap varies with types of sensor and
sensor installation. Regardless of gap differences, all sensors provide same function and output to the proximity
card in the PSEU.
IF 32-78
757 General Familiarization (7-2005)
ATA 32 - Page 116
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 35
OXYGEN SYSTEM ....................................................................................................................................................... 3
Overview ................................................................................................................................................................. 3
CREW OXYGEN SYSTEM COMPONENTS..................................................................................................................... 4
Crew Oxygen Masks ................................................................................................................................................ 6
System Shutoff Valve .............................................................................................................................................. 7
PASSENGER OXYGEN SYSTEM ................................................................................................................................... 8
PASSENGER OXYGEN SYSTEM COMPONENTS .......................................................................................................... 9
Oxygen Generator .................................................................................................................................................. 10
Passenger Oxygen Masks...................................................................................................................................... 10
Door Latch Actuator .............................................................................................................................................. 10
PORTABLE OXYGEN SYSTEM ................................................................................................................................... 11
757 General Familiarization (7-2005)
ATA 35 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 35 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
OXYGEN SYSTEM
Overview
The oxygen systems supply gaseous oxygen at low pressure to the crew and passengers.
Oxygen for the flight crew is kept in a cylinder in the forward cargo compartment. An oxygen line connects the
cylinder to oxygen masks at each crew station and oxygen box assembly.
Chemical oxygen generators are in the passenger service units. The generators supply oxygen to the passengers,
flight attendants, and lavatories.
There is one portable oxygen cylinder in the flight compartment. The passenger compartment has several portable
oxygen generators.
757 General Familiarization (7-2005)
ATA 35 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
CREW OXYGEN SYSTEM COMPONENTS
Oxygen Cylinder
The crew oxygen cylinder is on the right side of the forward cargo compartment. The cylinder has a volume of 115
cubic feet and is normally pressurized to 1850 PSI (12755 KPA) at 70°F (21.1ºC). There is a cylinder shutoff valve,
an oxygen pressure indicator, a thermal relief port, and a pressure regulator connection on the neck of each
cylinder.
Pressure Regulator and Pressure Transducer
The pressure regulator attaches to a cross-fitting on the neck of the crew oxygen cylinder. A hinged bracket,
installed on the airplane structure, holds the cross fitting. The inlet port of the pressure regulator receives the high
pressure oxygen from the crew oxygen cylinder. The regulator decreases this pressure to about 85 PSI (586 KPA).
An oxygen line connects the pressure regulator to the crew oxygen masks. The pressure transducer attaches to the
pressure regulator. The pressure transducer sends electrical signals to the EICAS which shows cylinder pressure.
Overboard Vent
If the crew oxygen cylinder becomes over-pressurized, a thermal relief disk on the cylinder neck breaks. The high
pressure oxygen flows through an overboard vent line to the overboard vent. When pressure in the vent line
reaches about 500 PSI, a green disk in the overboard vent is blown out and the oxygen is vented to the air. The
disk is held in place by a snap-ring. The overboard vent is just aft of the forward cargo door on the lower right
side of the fuselage. The green disk is easily seen during the pre-flight check.
757 General Familiarization (7-2005)
ATA 35 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 35 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Crew Oxygen Masks
Each crew station has an oxygen mask with a self-adjusting pneumatic harness. The masks receive oxygen at 60 to
85 PSI (413 to 586 KPA) from the pressure regulator. The masks have a diluter demand regulator which decreases
the oxygen pressure to a small quantity above cabin pressure. There is a selector on each mask which supplies
100 percent oxygen, or oxygen mixed with cabin air, to each crew member. A check valve in the mask/regulator
will vent the oxygen to the cabin if pressure to the mask increases to 100 PSI (689 KPA).
757 General Familiarization (7-2005)
ATA 35 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
System Shutoff Valve
There is a system shutoff valve in the flight compartment which controls oxygen flow to each crew station.
757 General Familiarization (7-2005)
ATA 35 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PASSENGER OXYGEN SYSTEM
Overview
The passenger oxygen system supplies oxygen to the passengers if needed. The system has chemical oxygen
generators, two switches with electrical latch actuation, and oxygen masks. The generators and masks are kept in
modules in the passenger service units (PSUs). A manual switch and/or aneroid switch control the latches on the
module doors. When either switch is closed the latch mechanism releases. This permits the module doors to open.
When the doors open, the masks fall into the reach of the passengers.
757 General Familiarization (7-2005)
ATA 35 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PASSENGER OXYGEN SYSTEM COMPONENTS
Oxygen Modules
Oxygen modules are installed above all the passenger seats and attendant stations. Each lavatory also has an
oxygen module in the ceiling. Each module has an oxygen generator and several mask assemblies. The module
doors open to let the masks fall into the reach of the passengers. The doors open electrically by an aneroid switch
or a hand operated switch. The module doors also open manually.
757 General Familiarization (7-2005)
ATA 35 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Oxygen Generator
Chemical oxygen generators chemically change sodium chlorate and iron into salt, iron oxide, and gaseous oxygen.
Each generator has a release pin which holds the firing pin in position. A lanyard connects the masks to the
release pin. Behind the firing pin is a primer to start the chemical reaction in the generator. Each generator has a
filter through which the oxygen flows to get to the manifold. The manifold has one outlet to each oxygen mask. A
relief valve on the generator opens if the pressure in the generator is 90 PSI (620 KPA) or above. There is a color
band on the generator that is temperature sensitive. When the generator is fired this color band turns black and
this is an indication that the generator is used and must be replaced.
Passenger Oxygen Masks
Passenger oxygen masks are kept in each oxygen module above the passenger seats. A supply tube connects the
oxygen generator manifold to each mask. Or the supply tube connects a supply line manifold, attached to the wall
of the oxygen box, to each mask. Each mask has a harness which permits the passenger or flight attendant to
quickly put the mask over the mouth and nose.
Door Latch Actuator
Each oxygen module has a door latch actuator. The actuator has a solenoid switch with a plunger. The plunger
strikes the latch lever and releases the module door. The solenoid switch is powered by 28-volts DC from the
overhead circuit breaker panel, P11.
757 General Familiarization (7-2005)
ATA 35 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PORTABLE OXYGEN SYSTEM
Overview
Oxygen for first aid and sustaining purposes are kept in portable oxygen cylinders. These cylinders are installed at
easily reached locations throughout the airplane.
Portable Oxygen Cylinder
There are two types of portable oxygen cylinders, those with a demand regulator and those without a demand
regulator. A demand-type mask can be attached to cylinders with a demand regulator.
A pressure gage shows oxygen pressure in the cylinder and thereby the quantity of oxygen available. Cylinder
pressure should be 1800 PSIG at 70°F. The safety plug contains a fusible alloy which melts in case of too much
heat. This permits the cylinder to vent into the atmosphere. The ON-OFF valve controls the flow of high pressure
oxygen into the pressure regulator.
757 General Familiarization (7-2005)
ATA 35 - Page 11
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 36
PNEUMATIC SYSTEM ................................................................................................................................................. 2
Overview............................................................................................................................................................. 2
Air Supply Distribution System (Distribution Ducting) ....................................................................................... 2
Engine Air Supply ............................................................................................................................................... 2
APU Air Supply................................................................................................................................................... 2
Ground Air Supply .............................................................................................................................................. 2
Air Supply Indicating System ............................................................................................................................. 3
PNEUMATIC DISTRIBUTION SYSTEM ........................................................................................................................ 5
Overview............................................................................................................................................................. 5
Pneumatic Air Sources ....................................................................................................................................... 5
Pneumatic Air Requirements .............................................................................................................................. 5
Air Supply Control .............................................................................................................................................. 5
Engine Air Supply System .................................................................................................................................. 6
PNEUMATIC DISTRIBUTION COMPONENTS .......................................................................................................... 8
Pneumatic Ground Connector ............................................................................................................................. 8
Isolation Valve .................................................................................................................................................... 8
Intermediate Pressure Check Valve .................................................................................................................... 9
High Pressure Shutoff Valve ............................................................................................................................. 10
High Stage (HS) Pilot ....................................................................................................................................... 10
Air Supply Pressure Regulating and Shutoff Valve ........................................................................................... 12
Air Supply Over-temperature Limiting Sensor.................................................................................................. 12
APU Shutoff Valve ............................................................................................................................................ 14
APU Check Valve .............................................................................................................................................. 14
Air Supply Pre-Cooler ....................................................................................................................................... 15
Fan Air Modulating Valve ................................................................................................................................. 16
Fan Air Temperature Sensor ............................................................................................................................ 16
Air Supply Altitude Switch ............................................................................................................................... 18
ECS Bleed Configuration Card .......................................................................................................................... 19
Reverse Flow Check Controller ......................................................................................................................... 20
INDICATION SYSTEM – (COMPONENTS) ................................................................................................................ 21
Air Supply Pressure Indication ......................................................................................................................... 21
Air Supply Temperature Indicating System ...................................................................................................... 22
PROVIDE/REMOVE PNEUMATIC POWER .................................................................................................................. 23
Overview........................................................................................................................................................... 23
Provide Pneumatic Power Using Auxiliary Power Unit...................................................................................... 23
Provide Pneumatic Power Using Ground Air Source......................................................................................... 25
757 General Familiarization (7-2005)
ATA 36 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PNEUMATIC SYSTEM
Overview
The pneumatic or air supply system supplies pressure and temperature regulated air to various systems. The
supplied air is used as a working fluid, system power source or for system pressurization.
The pneumatic system supplies air for these systems: aircraft pressurization, air-conditioning, wing leading edge
and engine cowl thermal anti-icing, engine starting, hydraulic reservoir pressurization, total air temperature (TAT)
probe ambient air induction, rain repellent system pressurization and potable water tank pressurization.
The pneumatic system consists of the Air Supply Distribution System and the Air Supply Indication System.
Air Supply Distribution System (Distribution Ducting)
The pneumatic or air supply system is primarily an air distribution system, channeling air through ducts from one
of several air sources to user systems. Pneumatic air is channeled through ducts to the hydraulic reservoir and
potable water system, the total air temperature (TAT) probe, the air conditioning packs, wing thermal anti-ice
ducts, and engine starter.
Engine Air Supply
The airplane engines provide the main source of pneumatic air. This air source is normally used whenever the
engines are running. Air is supplied from two ports located at the engine compressor section. Through these ports
a small amount of high pressure air is bled from the engines. The pneumatic system regulates the pressure and
temperature of the air and channels it to the user systems.
APU Air Supply
The auxiliary power unit (APU) can be used to provide system air when the airplane is on the ground. The APU
also provides in-flight bleed air backup in case of engine failure or an engine pneumatic air system malfunction.
Ground Air Supply
As an alternate to the APU, a pneumatic ground cart can supply air to the airplane whenever it is parked.
Connection between the ground cart and the pneumatic system is through three ground air connectors.
757 General Familiarization (7-2005)
ATA 36 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Supply Indicating System
1. Air Supply Pressure Indication - Pneumatic system duct pressure is read by duct pressure transducers
located on the body crossover duct. The pressure is displayed by the Engine
gine Indicating and Crew Alerting
System (EICAS) and on a dual duct pressure indicator.
2. Air Supply Temperature Indication - Bleed air temperature is sensed by the pre-cooler out temperature
bulb and displayed by the Engine Indicating and Crew Alerting System (EICAS).
IF 36-1
757 General Familiarization (7-2005)
ATA 36 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 36 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PNEUMATIC DISTRIBUTION SYSTEM
Overview
The air supply distribution system regulates the temperature and pressure of supply air extracted from the
engines. It also transports supply air from the pneumatic sources (engines, auxiliary power unit (APU), and ground
air connectors) to the user systems within the airplane.
Pneumatic Air Sources
The primary source of air for the pneumatic system is engine bleed air. Bleed air is extracted from the left and
right engines at intermediate and high pressure ports.
Two secondary sources are the auxiliary power unit (APU) and three ground air carts. The APU is used with the
airplane on the ground and as a backup system in flight. The ground air supply carts connect with three ground air
connectors to supply compressed air to the airplane. The ground connectors are located on the wing to body fairing
(two on the left side, one on the right side).
Pneumatic Air Requirements
The pneumatic system supplies air for aircraft pressurization, air conditioning, wing leading edge and engine cowl
thermal anti-icing, engine starting, hydraulic reservoir pressurization, total air temperature (TAT) probe ambient
air induction, rain repellent system pressurization, and potable water tank pressurization.
Air Supply Control
The Bleed Air Control Panel (control panel) on the pilot's overhead P5 panel provides air supply system control
and indication. The control panel contains indication lights, duct pressure indicator, and four switch/lights which
control valve position.
The HI STAGE light illuminates when a system overpressure occurs upstream of the PRSOV.
The BLEED light illuminates when a system overheat occurs.
The DUCT LEAK illuminates when one of the system duct leak sensors detect hot air indicating a duct leak.
757 General Familiarization (7-2005)
ATA 36 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Air Supply System
Each engine air supply system consists of the following components: pneumatic ducting, intermediate pressure
check valve, high pressure shutoff valve, high stage (HS) pilot, air supply pre-cooler, reverse flow check controller,
fan air modulating valve, fan air temperature sensor, air supply pressure regulating and shutoff valve and bleed air
over-temperature sensor.
The engine air supply system controls the pressure, temperature and airflow of engine bleed air. Air is bled off the
engine at the intermediate (HP2) pressure bleed port and high (HP6) pressure bleed port. At high engine power
settings, HP2 bleed air flows through the intermediate check valve, the pre-cooler, the pressure regulating and
shutoff valve, and to system users. The high pressure shutoff valve remains closed during most high power
settings. If the HP6 bleed air does not exceed the HP6 pressure limit schedule (101-113 PSIG up to 31,000 ft and
88-98 PSIG above 31,000 ft), the high pressure valve opens. The intermediate pressure check valve closes
whenever downstream pressure exceeds upstream pressure preventing backflow into the intermediate (HP2) engine
compressor section.
It is normal for the observed duct pressure (on the P5 panel) to decrease when the pneumatic pressure supply
goes from the high stage port (HP6) to the low stage port (HP2). This could occur at different times for either the
left or right sides. During this condition, you can observe a split in duct pressure from the left to right side. This
means that one side will have a higher duct pressure than the other side. You can see a graph of duct pressure
versus EPR (engine power) in the "Low Duct Pressure" fault isolation procedure of the FIM 36-10-00/101. This
graph gives the expected pressures for different engine power settings.
Engine bleed air is ducted through the air supply pre-cooler. The pre-cooler is a cross-flow, air to air, heat
exchanger which uses engine fan air as its cooling medium. Fan air is routed to the pre-cooler through the fan air
modulating valve which is attached to the bottom of the pre-cooler. The fan air modulating valve regulates the air
flow to the pre-cooler based on control air pressure from the fan air temperature sensor.
Bleed air leaving the pre-cooler is regulated by the pressure regulating and shutoff valve. The valve regulates
bleed air between 40-53 PSIG. At engine settings when the bleed air pressure is below 40-53 PSIG, the valve is full
open.
A reverse flow check controller prevents pneumatic air from flowing back into the engine. When a backflow
condition is sensed, the controller electrically actuates the closing solenoid of the pressure regulating and shutoff
valve and the HS pilot.
757 General Familiarization (7-2005)
ATA 36 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 36-2
757 General Familiarization (7-2005)
ATA 36 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PNEUMATIC DISTRIBUTION COMPONENTS
Pneumatic Ground Connector
The pneumatic ground connector is a 3-inch diameter nipple check valve. The nipple mates with a standard ground
cart connector. Three ground connectors are located in the wing to body fairing area. Two connectors on the left
side and one connector on the right side of the fuselage are located aft of the ram air inlet doors.
The ground connector is a spring loaded closed nipple check valve. The valve allows air flow in only one direction
to the air supply ducting.
Isolation Valve
One isolation valve located on the air supply body crossover duct allows cross bleeding of pneumatic air. The valve
is 5 inches in diameter.
The valve consists of a 115-volt AC electrically driven, two position actuator assembly. The actuator includes a
manual override lever connected to the butterfly linkage inside the actuator assembly.
IF 36-3
757 General Familiarization (7-2005)
ATA 36 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Intermediate Pressure Check Valve
The intermediate pressure check valve is a six inch flapper type check valve that serves two purposes:
1. The first is to allow unrestricted airflow from the intermediate pressure (HP2) engine bleed port.
2. The second function is to prevent higher pressure air from the high pressure shutoff valve from flowing
back into the engine.
The valve flappers open when flow pressure in the direction of the valve body flow arrow is larger than the flow
pressure in the opposite direction. The stop tube prevents the flappers from over-travel. Airflow on the flappers in
the opposite direction pushes the flappers closed. When closed, the valve prevents airflow in the opposite
direction.
IF 36-4
757 General Familiarization (7-2005)
ATA 36 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
High Pressure Shutoff Valve
The high pressure shutoff valve (HPSOV) is a pneumatically operated butterfly type valve. The valve is in the open
area between the engine and the strut. Get access to the valve through the thrust reverser cowling.
The HPSOV is used to shutoff high pressure (HP6) air when intermediate pressure (HP2) air can meet air supply
system requirements.
When the manual override hex is turned clockwise, the cam actuates a pushrod mechanism which moves the ball
valve from its normal position to the opposite seat. Upstream pressure is shut off and the high stage pilot supply
air is vented to ambient. The HPSOV butterfly is mechanically locked closed when the manual override hex is
turned clockwise such that the lever is engaged in the closed lock detent. To unlock the butterfly requires that the
lever be pulled out of the detent and that the hex is turned counterclockwise until it returns to its normal position.
1. AIRPLANES WITHOUT S212N101-52 AND SUBSEQUENT HPSOV DASH NUMBERS; the manual
override hex cannot move the HPSOV butterfly to the open position.
2. AIRPLANES WITH S212N101-52 AND SUBSEQUENT HPSOV DASH NUMBERS; the manual override
hex can be turned counterclockwise to open the HPSOV. The HPSOV cannot be locked open.
The HPSOV butterfly will modulate in response to changes in upstream pressure to maintain a 55 PSIG
downstream duct pressure.
A pin connected to the butterfly shaft serves as a visual position indicator.
High Stage (HS) Pilot
The HS pilot acts in conjunction with the high pressure shutoff valve to control the flow of high pressure (HP6)
bleed air. The HS pilot is attached to the bottom of the strut above the engine. The HS pilot is cooled by air routed
from pre-cooler cooling air intake duct. Access is obtained by opening thrust reverser cowling.
The HS pilot controls pressure to the opening chamber of the high pressure shutoff valve.
The HS pilot contains three shutdown modes, which act by reducing control pressure to near zero, allowing the
high pressure shutoff valve to close.
757 General Familiarization (7-2005)
ATA 36 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 36-5
757 General Familiarization (7-2005)
ATA 36 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Supply Pressure Regulating and Shutoff Valve
The air supply pressure regulating and shutoff valve maintains bleed air pressure to 40-53 PSIG. It shuts off bleed
airflow in response to excessive downstream temperature or if commanded from the BLEED AIR L or R ENG switch
light on P5 panel. The valve is located in the strut compartment downstream of the pre-cooler. Access is through
the pressure regulating and shutoff valve access panel.
Air supply over-temperature limiting sensor modulates the valve to maintain bleed air temperature at 450°F.
When the manual override hex is turned clockwise, the butterfly will close. The butterfly is mechanically locked
closed when the manual override hex is turned clockwise such that the lever is engaged in the closed lock detent.
To unlock the butterfly requires that the lever be pulled out of the detent and that the hex is turned
counterclockwise until it returns to its normal position.
1. AIRPLANES WITHOUT S212N101-53 AND SUBSEQUENT PRSOV DASH NUMBERS; the manual
override hex cannot move the PRSOV butterfly to the open position.
2. AIRPLANES WITH S212N101-53 AND SUBSEQUENT PRSOV DASH NUMBERS; the manual override
hex can be turned counterclockwise to open the PRSOV. The PRSOV cannot be locked open.
The closed position switch provides remote indication of butterfly position. The closed position switch is connected
to the OFF indicator light on the BLEED AIR L and/or R ENG switch light.
Air Supply Over-temperature Limiting Sensor
The air supply over-temperature limiting sensor limits the temperature of the bleed air downstream of the PRSOV.
The sensor vents reference pressure from the PRSOV to modulate the PRSOV toward closed. The sensor is located
in the strut compartment aft of the PRSOV.
When sensed temperature is above 450°F the ball valve moves off the valve seat allowing venting of reference
pressure. At 495°F the ball valve is full open.
757 General Familiarization (7-2005)
ATA 36 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 36-6
757 General Familiarization (7-2005)
ATA 36 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Shutoff Valve
The APU shutoff valve controls bleed air supplied from the APU. The valve is located between the pressure
bulkhead and APU fire wall. The valve is 5 inches in diameter.
APU Check Valve
The APU check valve permits bleed air flow from the APU to body crossover ducting only.
IF 36-7
757 General Familiarization (7-2005)
ATA 36 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Supply Pre-Cooler
The air supply pre-cooler cools bleed air to a temperature of 380 ± 20°F. The pre-cooler is located in the strut
compartment.
The pre-cooler is a cross-flow, single-pass, air to air heat exchanger. The bleed air inlet is divided into high (HP6)
and intermediate (HP2) inlet ports. Bleed air leaves the pre-cooler and is directed to the PRSOV. Cooling air flow is
controlled by the fan air modulating valve upstream of the cooling air inlet. Cooling air outlet is vented into the
strut compartment and exhausted.
The pre-cooler weighs approximately 50 lb.
IF 36-8
757 General Familiarization (7-2005)
ATA 36 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fan Air Modulating Valve
The fan air modulating valve controls the amount of fan air supplied to the pre-cooler. This regulates the
temperature of the engine bleed air to 360-400 degrees F.
The valve is located between the strut floor and the engine case. It is attached to the air supply pre-cooler. Access
to the valve is gained by opening the thrust reverser cowling.
The fan air modulating valve is a pneumatically activated, temperature controlled, spring loaded open butterfly
modulating valve. The fan air temperature sensor operates almost the same way as the air supply overtemperature limiting sensor by controlling the reference pressure to the valve causing the butterfly to move
proportional to sensed temperature. A small amount of fan air is bled from the fan air modulating valve to cool the
reverse flow check controller and high stage pilot.
The valve can be manual closed or opened. The manual wrenching hex is provided with a spring-loaded manual
lock lever and a visual position indicator pin connected to the valve shaft assembly. The manual override lock is
provided to mechanically lock the butterfly plate assembly into the full lock or full open position.
Fan Air Temperature Sensor
The fan air temperature sensor is a pneumatic bleed-off type temperature unit, and functions in conjunction with
the fan air modulating valve to maintain bleed air temperature downstream of the pre-cooler within a specified
range. The sensor is located above the strut compartment.
757 General Familiarization (7-2005)
ATA 36 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 36-9
757 General Familiarization (7-2005)
ATA 36 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Supply Altitude Switch
The air supply altitude switch is a barometric switch whichh closes when an altitude of 31,000 feet is reached. The
switch is located in the right ECS bay forward of the a/c pack system. When the airplane reaches 31,000 feet the
switch closes. The closed switch changes the switchover schedule in the HS pilot.
IF 36-10
757 General Familiarization (7-2005)
ATA 36 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ECS Bleed Configuration Card
The ECS bleed configuration card receives input signals from
om the pressure regulating and shutoff valve position
switches, wing and engine anti-ice switches and the pack flow cards. These input signals are provided as analog
output signals, by the card, to the flight management computers, thrust management computer, and electronic
engine controllers. The computers and controllers use these signal inputs to match required engine output to
current airplane thrust and pneumatic demands.
The ECS bleed configuration card contains two BITE test circuits, OUTPUT TEST HIGH and OUTPUT TEST LOW.
The card BITE checks the cards' internal circuits for damage.
IF 36-11
757 General Familiarization (7-2005)
ATA 36 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Reverse Flow Check Controller
The controller is located in the strut forward of the pre-cooler. The reverse flow check controller is a pneumatically
actuated controller. The controller compares upstream and downstream duct pressure to actuate a normally closed
sensitive switch. The controller is cooled by air routed from the fan air modulating valve.
IF 36-12
757 General Familiarization (7-2005)
ATA 36 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
INDICATION SYSTEM – (COMPONENTS)
Air Supply Pressure Indication
Two duct pressure transducers provide signals to a dual duct pressure indicator for pressure indication on the
pilots' overhead P5 panel for the air supply system.
Two additional duct pressure transducers provide the Engine Indicating and Crew Alerting System (EICAS) with
signals for pressure indication in the flight compartment when on the ground only.
All transducers sense available system pressure on either side of the air supply isolation valve. All four
transducers are located just forward of the air supply body cross over ducting.
IF 36-13
757 General Familiarization (7-2005)
ATA 36 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Supply Temperature Indicating System
Air Supply system overheat protection is provided by two 490°F (254°C)
C) overheat switches. The overheat switches
sense an over-temperature condition on each side of the isolation valve. The L or R BLEED light on the P5 panel,
the EICAS caution Message L or R ENG BLEED VAL, and an aural warning (owl) provide indication of the sensed
overheat condition. The PRSOV and high stage valve are commanded closed in an overheat condition.
The pre-cooler discharge temperature sensor provides EICAS with temperature indication on the ground only. The
sensor is located downstream of the PRSOV on a duct section in the wing leading edge above the strut.
IF 36-14
757 General Familiarization (7-2005)
ATA 36 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PROVIDE/REMOVE PNEUMATIC POWER
Overview
This section covers the standard procedure for removing and providing pneumatic power using the auxiliary power
unit (APU), or a ground air source.
Pressurization of the pneumatic system is required to operate any air-driven system.
Provide Pneumatic Power Using Auxiliary Power Unit
Provide electrical power.
WARNING:
PROVIDING PNEUMATIC POWER WILL SUPPLY PNEUMATIC PRESSURE TO OPERATE THE
AIR CONDITIONING SYSTEM AND ENGINE START SYSTEM; TO PRESSURIZE THE
POTABLE WATER TANKS AND HYDRAULIC RESERVOIR; AND PROVIDE BLEED AIR TO TAT
PROBE, AND NACELLE AND WING ANTI-ICE DUCTS. CARE SHOULD BE TAKEN TO
ISOLATE THOSE SYSTEMS AND CONTROLS NOT INTENDED FOR OPERATION TO
PREVENT LOSS OF PRESSURE, TO PREVENT INADVERTENT ACTUATION OF EQUIPMENT,
DAMAGE TO AIRPLANE, AND INJURY TO PERSONNEL.
CAUTION:
ALWAYS APPLY ELECTRICAL POWER BEFORE APPLYING PNEUMATIC POWER AND
REMOVE PNEUMATIC POWER BEFORE REMOVING ELECTRICAL POWER TO PREVENT
POSSIBLE DAMAGE TO PNEUMATIC AIR SYSTEM USERS.
Start APU.
Verify that APU RUN light is lit on the pilots' overhead P5 panel.
757 General Familiarization (7-2005)
ATA 36 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Provide Pneumatic Power Using Auxiliary Power Unit (Continued):
Press BLEED AIR control APU VALVE switch/light on P5 panel to the open position. Verify that white bar comes on.
Verify that VALVE light comes on and then goes out.
If right air-conditioning pack or right side pneumatic ducting requires pneumatic power, press BLEED AIR control
ISOLATION VALVE switch/light on the P5 panel to open position. Verify that white flow bar light comes on. Verify
that VALVE light comes on and then goes out.
Observe BLEED AIR dual DUCT PRESS indicator on P5 panel and verify that both L and/or R pointer(s) indicate
positive pressure.
757 General Familiarization (7-2005)
ATA 36 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Provide Pneumatic Power Using Ground Air Source
You will need a pneumatic source capable of delivering clean air at the required pressure and flow values, with a
connection to mate with a 3-inch diameter check valve on the airplane ground pneumatic service connector.
Statically ground the airplane.
WARNING:
PROVIDING PNEUMATIC POWER WILL SUPPLY PNEUMATIC PRESSURE TO OPERATE THE
AIR CONDITIONING SYSTEM AND ENGINE START SYSTEM; TO PRESSURIZE THE
POTABLE WATER TANKS AND HYDRAULIC RESERVOIR; AND PROVIDE BLEED AIR TO TAT
PROBE, AND NACELLE AND WING ANTI-ICE DUCTS. CARE SHOULD BE TAKEN TO
ISOLATE THOSE SYSTEMS AND CONTROLS NOT INTENDED FOR OPERATION TO
PREVENT LOSS OF PRESSURE, TO PREVENT INADVERTENT ACTUATION OF EQUIPMENT,
DAMAGE TO AIRPLANE, AND INJURY TO PERSONNEL.
CAUTION:
ALWAYS APPLY ELECTRICAL POWER BEFORE APPLYING PNEUMATIC POWER AND
REMOVE PNEUMATIC POWER BEFORE REMOVING ELECTRICAL POWER TO PREVENT
POSSIBLE DAMAGE TO EQUIPMENT.
Open pneumatic ground service access door 193JL and 194FR, found on ECS access doors in the body to wing
fairing section. There are two pneumatic connectors found behind the left access door and there is one pneumatic
connector found behind the right access door. The number of connections made depends on system demands, two
ground carts may be necessary for engine starting.
Position ground pneumatic source equipment away from work area and connect pneumatic line(s) to ground
pneumatic service connector(s).
Provide electrical power.
757 General Familiarization (7-2005)
ATA 36 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Provide Pneumatic Power Using Ground Air Source (Continued):
WARNING:
DO NOT SUPPLY MORE THAN 45 PSIG OF PRESSURE TO THE PNEUMATIC SYSTEM. IF
YOU SUPPLY TOO MUCH PRESSURE, DAMAGE TO EQUIPMENT AND INJURY TO
PERSONNEL CAN OCCUR.
Start ground pneumatic source and do not supply more than 45 PSIG.
NOTE:
Engine starting requires high airflow and may require at least two ground service carts.
If pneumatic power is required for all system users and ground connection is made on left or right connector only,
depress bleed air control ISOLATION VALVE switch/light on the pilots' overhead P5 panel. Verify that white bar
light comes on. Verify that VALVE light comes on and then goes out.
Observe BLEED AIR dual DUCT PRESS indicator on P5 panel and verify L and/or R pointer(s) indicate positive
pressure. Pressure indication observed will depend upon the ground source used and system load.
757 General Familiarization (7-2005)
ATA 36 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 38
THIS PAGE INTENTIONALLY LEFT BLANK................................................................................................................. 2
POTABLE WATER SYSTEM......................................................................................................................................... 3
Overview............................................................................................................................................................. 3
Potable Water System Components Details ................................................................................................................. 6
Water Tank ......................................................................................................................................................... 6
Water Quantity Transmitter ................................................................................................................................ 7
Air Compressor .................................................................................................................................................. 8
THIS PAGE INTENTIONALLY LEFT BLANK................................................................................................................. 9
Air Filters ......................................................................................................................................................... 10
Pressure Switch ............................................................................................................................................... 10
Pressure Relief Valve........................................................................................................................................ 10
Pressure Regulator ........................................................................................................................................... 10
Pressurization Line Check Valves ..................................................................................................................... 12
Lavatory and Galley Shutoff and Drain Valves .................................................................................................. 13
Lavatory Wash Basin Faucet............................................................................................................................. 14
Water Heaters ................................................................................................................................................... 15
Service Panel.................................................................................................................................................... 16
Fill/Overflow Valve .......................................................................................................................................... 17
Water Quantity Indicators ................................................................................................................................. 18
THIS PAGE INTENTIONALLY LEFT BLANK............................................................................................................... 19
WASTE DISPOSAL SYSTEM ..................................................................................................................................... 20
Overview........................................................................................................................................................... 20
Toilet Waste System Operation ........................................................................................................................ 20
LAVATORY TOILET WASTE SYSTEM COMPONENTS ........................................................................................... 22
Toilet Tank ....................................................................................................................................................... 22
Toilet Bowl ....................................................................................................................................................... 22
Water Separator ............................................................................................................................................... 23
Motor-Pump-Filter Unit .................................................................................................................................... 24
Toilet Tank Drain Valve .................................................................................................................................... 25
Rinse/Fill Line Check Valve ............................................................................................................................. 26
Rinse/Fill Line Shutoff Valve (Not On All Aircraft) .......................................................................................... 27
Flush Timer ...................................................................................................................................................... 28
757 General Familiarization (7-2005)
ATA 38 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 38 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
POTABLE WATER SYSTEM
Overview
The potable water system has a water tank which stores fresh water. The system supplies water to the galley units
and lavatory wash basins.
The storage and distribution components of the potable water system, store, distribute, and drain fresh water. The
potable water tank is in the compartment aft of the bulk cargo compartment.
Each galley and lavatory has water shutoff valves above the floor. The valves allow any one unit or combination of
units, to be isolated while the rest of the units remain operating. The lavatory wash basins each have a faucet with
hot and cold water. The faucets are self venting. This allows automatic bleeding of air from the system, and allows
draining the system without opening the faucet.
Most of the water distribution system is flexible Teflon hose with a reinforced fiber covering. Metal fittings and
connectors are used at junctions and line replaceable units. Distribution lines route from below the water tank to
above the cabin ceiling. The lines enter galleys and lavatories from above at Doors 1, 2, and 3, and from below at
Door 4.
The water heaters provide hot water to the wash basins in each lavatory. Each lavatory has a water heater in the
supply line to the wash basin faucet.
Potable water quantity indicators show ground service personnel and the flight crew how much water is in the
potable water tank. The water quantity sensing system sends quantity signals to the water quantity indicators.
The potable water tank is pressurized to force water from the tank to the lavatories and galleys. The pressurization
system includes an air compressor, air filters, pressure relief valve, and pressure switch. Bleed air from the
pneumatic system provides additional pressurization.
757 General Familiarization (7-2005)
ATA 38 - Page 3
For Training Purposes Only
IF 38-1
757 General Familiarization (7-2005)
ATA 38 - Page 4
AeroEd LLC – Aviation Education Resource
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 38-1A
757 General Familiarization (7-2005)
ATA 38 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Potable Water System Components Details
Water Tank
The water tank is in the lower lobe, aft of the bulk cargo compartment. The tank is on the right side of the airplane
with its longitudinal axis running forward and aft. Vertical support rods attach the tank to floor beams. A diagonal
brace attaches to the aft end of the tank to prevent swaying.
The tank has a capacity of 60 gallons but is stand-piped to 50 gallons. The tank is nonmetallic monofilament spun
fiberglass. Two metal bands go around the tank for reinforcement and mounting support.
On the inboard side of the water tank is a mounting pad for the water quantity transmitter. A water quantity sensor
is molded inside the water tank wall. An air line enters the top of the tank for tank pressurization. The tank has
fittings on the upper side. One fitting connects to the water fill line; the other to the tank overflow line. A hose
attaches to the bottom of the tank for water distribution and draining.
IF 38-2
757 General Familiarization (7-2005)
ATA 38 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Water Quantity Transmitter
The water quantity transmitter attaches to the mounting pad on the side of the water tank. The transmitter plugs
into the mounting pad sensor outlets and fastens to the pad with screws. The water quantity transmitter receives a
signal from a sensor molded inside the water tank. The transmitter output is 0 to 10v dc. The transmitter receives
28-volts DC power from the APU/EXT power panel, P34, and the overhead circuit breaker panel, P11.
IF 38-3
757 General Familiarization (7-2005)
ATA 38 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Compressor
An electric motor-driven air compressor is mounted to the right of the potable water tank. The air compressor is
the main source of pressurization for the potable water tank. Compressor operation is controlled by a pressure
switch and a water pressure system relay.
IF 38-4
757 General Familiarization (7-2005)
ATA 38 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 38 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Filters
All air entering the potable water tank passes through an air filter. One filter is in the engine bleed air line to the
water tank. The other filter is in the inlet line to the air compressor. Both filters have removable elements for easy
servicing.
Pressure Switch
The pressure switch provides a ground to control the water pressure system relay. The switch closes, when
pressure in the water tank drops below the set point, allowing the air compressor to operate. At a pressure above
the set point the switch opens stopping compressor operation.
Pressure Relief Valve
A pressure relief valve protects the potable water tank from being over-pressurized. The valve is in the air
pressure line to the water tank so it protects against a malfunction in either the air compressor or the pneumatic
system.
Pressure Regulator
The pressure regulator receives air from the engine bleed air pneumatic system and reduces the pressure to a
maximum of 40±3 PSI.
757 General Familiarization (7-2005)
ATA 38 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 38-5
757 General Familiarization (7-2005)
ATA 38 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Pressurization Line Check Valves
Check valves are installed in the air pressure
sure lines to the water tank to prevent reverse flow of pressure from the
tank to the pneumatic system or air compressor.
ON AIRPLANES WITH SB 38-14; the check valve in the outlet line from the air compressor includes an un-loader
valve to vent pressure from the outlet line when the compressor is not operating. This reduces the starting load on
the air compressor.
IF 38-6
757 General Familiarization (7-2005)
ATA 38 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Lavatory and Galley Shutoff and Drain Valves
One two-port manually operated shutoff valve is installed in the fresh water supply line inside each lavatory and
each galley. The lavatory valve is about ten inches above the floor inside the lavatory sink cabinet. The location of
the valve in the galleys varies with galley location and are identified by placards. The valves allow any one unit, or
combination of units, to be isolated while the rest of the system remains in operation. The valve is a component of
the galley or lavatory unit and remains in the unit if the unit is removed.
Another two-port manually operated valve is installed in the fresh water drain line inside each lavatory and each
galley. The lavatory valve is located in the drain line at floor level inside the lavatory sink cabinet. The location of
the valves in the galleys varies according to galley location and identified by placards. The valves allow the faucets
and potable water supply lines to be drained after the water has been shutoff to the lavatory or galley. The valve is
a component of the galley or lavatory unit and remains in the unit if the unit is removed.
IF 38-7
757 General Familiarization (7-2005)
ATA 38 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Lavatory Wash Basin Faucet
Each lavatory wash basin has a water faucet. Each faucet has a hot and cold water valve with a common mixing
spigot. The faucets are self venting so they need not be opened to drain the potable water system. The wash basin
stopper is spring loaded closed. The stopper control lever must be held open until the basin is empty. The basin
overflow drain is not stoppered. A muffler in the overflow line quiets the sound of differential pressure escaping.
IF 38-8
757 General Familiarization (7-2005)
ATA 38 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Water Heaters
Each water heater is a three-pint tank with three 140 watt
tt heating elements. Each heater has two thermal switches;
an ON/OFF switch and indicator light; and a pressure relief valve.
IF 38-9
757 General Familiarization (7-2005)
ATA 38 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Service Panel
The potable water service panel is on the bottom centerline of the fuselage below the aft service door. The service
panel has both inlet and outlet ports for the fill/overflow valve. Also on the service panel is a water quantity
indicator and a control handle for the fill/overflow valve. The service panel has a drain handle to operate the
potable water drain valve.
IF 38-10
757 General Familiarization (7-2005)
ATA 38 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fill/Overflow Valve
The fill/overflow valve is a four-port valve used to fill the potable water tank. The fill/overflow valve attaches to a
floor beam to the right of the potable water tank. The handle to operate the fill/overflow valve is on the potable
water service panel. In the open or service position, the fill/overflow valve allows the potable water tank to be
filled. In the closed or flight ready position, the fill/overflow valve seals the potable water tank so it can be
pressurized.
IF 38-11
757 General Familiarization (7-2005)
ATA 38 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Water Quantity Indicators
There are two water quantity indicators. One is on the potable
table water service panel; the other on the forward galley
sidewall. The indicators are 1.35 inch diameter gages with a scale of 0 to 60 gallons. The indicators receive
electrical signals from the water quantity transmitter.
IF 38-12
757 General Familiarization (7-2005)
ATA 38 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 38 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
WASTE DISPOSAL SYSTEM
Overview
Each lavatory has an independent toilet waste system. Toilet waste is stored in a toilet tank in each lavatory.
During ground servicing the toilet tanks are drained, rinsed and a chemical pre-charge is added. The toilet flushing
medium is pre-charge mixed with toilet waste water which has been filtered and deodorized.
Toilet Waste System Operation
The toilet flushing cycle is started when the flush timer handle is depressed. This provides power to the flush
motor which operates the pump. Filtered water is pumped into the toilet bowl until the flush timer cuts power to
the flush motor (about ten seconds).
On Some aircraft when ground power is applied, the fill/shutoff valve opens. During servicing, the valve closes if
the liquid level in the tank reaches three inches from the top of the tank. The valve reopens when the liquid level
drops 0.5 ± .25 inches. The valve closes when ground power is removed.
Pressing the flush timer handle starts the toilet system. Operation is automatic and is controlled by time switches.
757 General Familiarization (7-2005)
ATA 38 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 38-13
757 General Familiarization (7-2005)
ATA 38 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
LAVATORY TOILET WASTE SYSTEM COMPONENTS
Toilet Tank
Each lavatory has a storage tank to hold toilet waste. The tanks are fire resistant plastic and are vented to the
lavatory vent system. Each tank is emptied, flushed and pre-charged through service panels on the bottom of the
fuselage.
Toilet Bowl
A stainless steel toilet bowl attaches to the top of each toilet tank. A separator between
een the tank and the toilet
bowl prevents passengers from seeing into the tank. The separator also prevents liquid in the tank from sloshing
up into the bowl. The separator is spring-loaded closed, but opens during the flush cycle. The tank and toilet bowl
are covered with a fiberglass shroud.
IF 38-14
757 General Familiarization (7-2005)
ATA 38 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Water Separator
A water separator is attached to each toilet system. The water separator is a plastic rectangular box with diagonal
metal condensation plates enclosed. The separator mounts behind the toilet back shroud. The water separator
connects the toilet tank vent to the airplane onboard ventilation system. The water separator prevents moisture
from the toilet tank from entering the onboard ventilation system. A vent near the upper edge of the toilet bowl
connects to a point just downstream of the water separator and helps to eliminate toilet bowl odors.
IF 38-15
757 General Familiarization (7-2005)
ATA 38 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Motor-Pump-Filter Unit
Each toilet waste tank has a motor-pump-filter unit. This unit attaches to the top of the tank and pumps filtered
flushing fluid into the toilet bowl. The motor operates on 115-volts AC power from the overhead circuit breaker
panel, P11. The filter is self cleaning and requires no special servicing.
IF 38-16
757 General Familiarization (7-2005)
ATA 38 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Toilet Tank Drain Valve
There is a drain valve in the bottom of each toilet tank. The drain valve is spring-loaded shut to prevent leakage
from the tank. The drain valve is removed through the top of the tank. A cable, from the ground service panel,
opens the valve allowing the tank to drain.
IF 38-17
757 General Familiarization (7-2005)
ATA 38 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rinse/Fill Line Check Valve
A two-port check valve is mounted on the top of each toilet tank. One port connects to the tank rinse/fill line and
the other port to a line from the anti-siphon valve. The valve
lve allows the rinse/fill line to drain after servicing, and
prevents liquid from siphoning out of the toilet tank.
AIRPLANES POST-SB 38-3; a three-port check valve is mounted on the top of each toilet tank. One port connects
to the tank rinse/fill line. The second port connects to a line from the anti-siphon valve, and the third port
connects to the vacuum-breaker line from the water separator. The valve allows the rinse/fill line to drain after
servicing, and prevents liquid from siphoning out of the toilet tank. The connection of the vacuum-breaker line to
the valve eliminates leakage at the service panel.
IF 38-18
757 General Familiarization (7-2005)
ATA 38 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Rinse/Fill Line Shutoff Valve (Not On All Aircraft)
A two-port fill/shutoff valve is mounted on the top of each toilet tank. The valve is installed in the tank rinse/fill
line. The valve prevents overfilling the toilet tank during servicing, allows the rinse/fill line to drain after
servicing, and prevents liquid from siphoning out of the toilet tank. The fill/shutoff valve has an electric motor
which is activated by a sensor which monitors liquid level inside the toilet tank.
IF 38-19
757 General Familiarization (7-2005)
ATA 38 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flush Timer
The flush timer is mounted on the lavatory
tory wall. The timer is actuated by the toilet flush handle, and provides 115volts AC power to the flush motor for ten seconds.
IF 38-20
757 General Familiarization (7-2005)
ATA 38 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 49
AUXILIARY POWER UNIT ............................................................................................................................................ 5
Overview............................................................................................................................................................. 5
Power Section..................................................................................................................................................... 6
Compressor ........................................................................................................................................................ 6
Combustor .......................................................................................................................................................... 6
Turbine ............................................................................................................................................................... 6
Load Compressor ............................................................................................................................................... 8
Bearings ............................................................................................................................................................. 8
Seals .................................................................................................................................................................. 8
Gearbox .............................................................................................................................................................. 9
APU AIR INTAKE SYSTEM & COMPONENTS............................................................................................................ 10
Overview........................................................................................................................................................... 10
Air Intake Door ................................................................................................................................................. 10
Air Intake Door Housing ................................................................................................................................... 10
Air Intake Door Seal ......................................................................................................................................... 10
Air Intake Door Actuator ................................................................................................................................... 10
Air Intake Ducts ............................................................................................................................................... 12
Air Intake Plenum............................................................................................................................................. 12
APU DRAINS AND VENTS ......................................................................................................................................... 14
Overview........................................................................................................................................................... 14
APU AND GENERATOR LUBRICATION SYSTEM........................................................................................................ 18
Overview........................................................................................................................................................... 18
APU AND GENERATOR LUBRICATION SYSTEM COMPONENTS ........................................................................... 20
Oil Reservoir..................................................................................................................................................... 20
Oil Pump Assembly .......................................................................................................................................... 21
De-Oil Solenoid Valve ....................................................................................................................................... 22
Switch for the Low Oil Temperature (APUs with the switch for the low oil temperature) ................................ 22
Gearbox Pressurization System ........................................................................................................................ 24
Oil Cooler ......................................................................................................................................................... 25
APUs WITH THE MAGNETIC CHIP DETECTORS FOR THE TURBINE BEARING AND THE COMPRESSOR BEARING . 26
Generator Scavenge Pump................................................................................................................................ 27
Filter Assembly for the Generator Scavenge Oil ............................................................................................... 27
APU AND GENERATOR LUBRICATION SYSTEM OIL SERVICING............................................................................... 28
Fill the APU Oil (Gravity Fill Method) ............................................................................................................... 30
Fill APU Oil (Pressure Fill Method) .................................................................................................................. 31
APU ENGINE FUEL SYSTEM ..................................................................................................................................... 34
Overview........................................................................................................................................................... 34
FUEL SYSTEM COMPONENTS .................................................................................................................................. 36
Fuel Control Unit .............................................................................................................................................. 36
757 General Familiarization (7-2005)
ATA 49 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Inlet Filter......................................................................................................................................................... 36
Fuel Pump ........................................................................................................................................................ 36
High-Pressure Relief Valve ............................................................................................................................... 36
High Pressure Filter ......................................................................................................................................... 36
Actuator Pressure Regulator ............................................................................................................................. 36
Fuel Metering Assembly ................................................................................................................................... 38
Pressurizing Valve ............................................................................................................................................ 38
Fuel Shutoff Solenoid Valve.............................................................................................................................. 38
Flow Divider ..................................................................................................................................................... 39
Nozzles ............................................................................................................................................................. 40
Manifolds ......................................................................................................................................................... 40
APU IGNITION/STARTING SYSTEM ......................................................................................................................... 41
Overview........................................................................................................................................................... 41
APU IGNITION/STARTING SYSTEM COMPONENTS ............................................................................................ 42
Ignition Unit...................................................................................................................................................... 42
Ignition Lead .................................................................................................................................................... 42
Igniter ............................................................................................................................................................... 42
Starter Motor .................................................................................................................................................... 44
APU COOLING AIR SYSTEM...................................................................................................................................... 45
Overview........................................................................................................................................................... 45
APU COOLING AIR SYSTEM COMPONENTS ......................................................................................................... 48
Fan Isolation Valve ........................................................................................................................................... 48
Cooling Fan ...................................................................................................................................................... 49
Oil Cooling Air Duct.......................................................................................................................................... 50
Compartment Cooling and Ventilation............................................................................................................... 50
APU EXHAUST GAS TEMPERATURE INDICATING SYSTEM & COMPONENTS.......................................................... 51
Overview........................................................................................................................................................... 51
Thermocouple Assemblies................................................................................................................................ 51
APU TIME TOTALIZER SYSTEM & COMPONENTS ................................................................................................... 52
Overview........................................................................................................................................................... 52
APU OIL INDICATING SYSTEM ................................................................................................................................. 53
Overview........................................................................................................................................................... 53
APU OIL INDICATING SYSTEM COMPONENTS .................................................................................................... 54
Low Oil Pressure Switch .................................................................................................................................. 54
Oil Temperature Sensor ................................................................................................................................... 54
Oil Quantity Transmitter................................................................................................................................... 55
Differential Pressure Switch for the Generator Oil Filter .................................................................................. 55
APU CONTROL SYSTEM ........................................................................................................................................... 56
Overview........................................................................................................................................................... 56
APU CONTROL SYSTEM COMPONENTS .............................................................................................................. 57
APU Control Panel - P5 Panel .......................................................................................................................... 57
Power Controls - P5 Panel ............................................................................................................................... 57
P62 Control Panel for the APU ......................................................................................................................... 58
APU Control Unit .............................................................................................................................................. 60
757 General Familiarization (7-2005)
ATA 49 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Monopole .......................................................................................................................................................... 62
APU Inlet Temperature Sensor ......................................................................................................................... 63
APU Inlet Pressure Sensor ............................................................................................................................... 63
APU OPERATION PROCEDURES START & SHUTDOWN ........................................................................................... 64
APU Start ......................................................................................................................................................... 64
APU Shutdown .................................................................................................................................................. 67
757 General Familiarization (7-2005)
ATA 49 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 49 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
AUXILIARY POWER UNIT
Overview
The auxiliary power unit is a gas turbine engine with one shaft, located in section 48, in the rear of the airplane.
The unit provides electrical and pneumatic power for in-flight and ground operations.
The APU provides electrical power to the airplane by an oil cooled generator. The APU provides pneumatic power
by an engine-driven load compressor. The supply of APU electrical power is given priority over APU pneumatic
power.
The APU control panel on the overhead panel P5 contains the APU start/shutdown (master control) switch and
some APU indicators. Other APU indication is displayed on the Engine Indication and Crew Alerting System
(EICAS) panel (P2). The aft control panel P8 of the pilot contains the APU fire handle and the bottle discharge
light. The APU control panel P62, on the nose landing gear, contains an APU shutdown switch and a discharge
switch for the fire bottle. The E6 rack of the aft equipment center holds the APU control unit, battery, and battery
charger.
IF 49-1
757 General Familiarization (7-2005)
ATA 49 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Power Section
The power section converts compressed air and jet fuel into shaft horsepower. It consists of a two-stage
centrifugal compressor, an annular combustor, and a three-stage axial turbine. All components are connected to a
single shaft. This shaft drives the load compressor, electric generator, and engine accessories.
Compressor
The compressor portion of the power section consists of inlet housing, two centrifugal compressors, and two steel
diffuser sections. The compressor rotors are titanium alloy forgings with integral rotor blades. Panels on the right
and left sides of the compressor inlet plenum allow borescope inspection of the first-stage compressor blades for
damage. Through couplings, the compressor rotors are mounted on and driven by the power section shaft. The
compressor housing provides the mounting base for the load compressor section.
Combustor
The combustion flame is contained within a reverse-flow annular combustor. Air enters through axial, de-swirl
vanes into a turbine plenum cavity area around the combustor. Orifices in the combustor shell allow air to cool the
chamber and regulate gas temperature. Near the aft end of the combustion chamber, holes are provided for the
igniter plug and 12 fuel atomizers.
Turbine
The turbine portion of the power section contains three axial-flow turbines with no cooling air, turbine plenum
housing, and an exhaust port. The first two stages contain cast alloy blades with fir tree attachment to the disk.
The third stage rotor is a one-piece forged configuration. The turbine rotors are mounted through couplings on the
power section shaft. Full containment is provided for all rotor blades. The stator vanes are mounted between an
inner and outer ring. The inner ring is riveted to a circular flange assembly around the inside of the stator ring.
The exhaust port is an extension of the turbine plenum housing.
Bearings
The power section shaft is supported by two antifriction type bearings. On the turbine end is a roller bearing with
a carbon ring seal. The load compressor end has a ball bearing with a divided inner ring and a redundant double
seal. The primary seal is a carbon face seal backed up by a secondary labyrinth seal. The annulus area between
the two seals is vented overboard. The bearings are mounted in an annular oil film surrounding the bearing outer
race. The oil is re-circulated around the bearing to cool it and absorb some vibration.
757 General Familiarization (7-2005)
ATA 49 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-2
757 General Familiarization (7-2005)
ATA 49 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Load Compressor
The load compressor provides compressed air for main engine starting and cabin air conditioning. The section
consists of a centrifugal compressor, a diffuser, and inlet housing. The load compressor is driven at engine shaft
speed. The inlet guide vanes regulate airflow through the compressor.
The load compressor is a single-stage centrifugal compressor of forged titanium-alloy construction. The
compressor rotor is mounted through couplings on its own shaft. An intermediate quill shaft connects the
compressor shaft to the power section shaft and rotates it at power section speed. Access ports on the right and
left sides of the intake plenum allow inspection of the load compressor blades for damage. Inlet guide vanes at the
compressor inlet control the amount of airflow through the compressor. The load compressor housing provides the
mounting base for the accessory drive gearbox.
Bearings
The load compressor shaft is supported by two bearings. The intake end has a ball bearing with a divided inner
ring. The discharge end has a roller bearing. Both bearings have hydraulic mounting.
Seals
The load compressor bearings use a redundant double seal. The primary seal is a carbon face seal with a positive
contact and a secondary labyrinth seal. The annulus between the seals is vented overboard.
IF 49-3
757 General Familiarization (7-2005)
ATA 49 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Gearbox
The accessory drive reduces the turbine speed to drive components necessary for engine operation and to drive the
electrical generator. A flexible quill shaft connects the gear
ar train to the load compressor shaft. The gear train
consists of eight spur gears that are straddle mounted. A sprag clutch under the starter gear engages the starter
motor. The gears are oil jet lubricated. An integral 6-quart capacity oil reservoir is located at the bottom of the
gearbox.
A sight gauge and oil level sensor are used to monitor the oil level. An air-oil separator in the gearbox removes
any entrapped oil from the air before it is vented overboard. Accessories mounted on the gearbox include the lube
pumps, fuel control unit, oil cooling fan, starter motor, and generator.
IF 49-4
757 General Familiarization (7-2005)
ATA 49 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU AIR INTAKE SYSTEM & COMPONENTS
Overview
The air intake system for the APU provides control and passage of air from the exterior of the airplane to the inlet
plenum for the APU compressor. The intake door, intake door seal, intake door housing, intake door actuator, and
intake ducts are located forward of the APU firewall on the right side above the horizontal stabilizer. The air intake
plenum is mounted on the aft side of the APU firewall and serves as the major support of the APU. The intake
door actuator opens the door outward to draw air through the ducts and plenum and into the inlet plenum for the
APU compressor. The actuator is powered by the APU battery bus with 28-volts DC.
Air Intake Door
The door opens outward to draw air into the APU. It must be fully opened for APU operation. An electromechanical
linear actuator operates the door. A switch mounted on the door housing indicates door open position to the APU
control unit.
Air Intake Door Housing
The housing surrounds the intake door and provides for the connection of the transition duct and the door actuator.
It is bolted to the outer skin of the airplane between the vertical and right horizontal stabilizers. The housing is a
composite of graphite-kevlar-fiberglass material.
Air Intake Door Seal
The seals are located inside the door housing around the door opening. The seal is in four sections bolted to the
housing along the front, rear, and sides.
Air Intake Door Actuator
The door actuator mounts on the intake door housing with the rod end bolted to the door. It consists of an electric
motor, actuator rod, jackscrew, reduction gear train, four limit switches, and a drive socket and clutch for manual
operation. Mechanical stops are located beyond the limit switches in case of limit switch failure. The actuator
square drive provides for manual operation of the actuator. The actuator is powered by the APU battery bus with
28-volts DC.
757 General Familiarization (7-2005)
ATA 49 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-5
757 General Familiarization (7-2005)
ATA 49 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air Intake Ducts
The duct forms an air passage from the intake housing to the intake plenum. The ducting consists of three
sections: transition duct, elbow duct, and diffuser duct. The transition duct bolts to the intake door housing on the
forward end and the elbow duct on the aft end. The elbow duct attaches to the pivot bulkhead and forms a slipjoint connection with the diffuser duct. The diffuser duct, in two sections, then bolts to the APU firewall.
Air Intake Plenum
The intake plenum serves as the supporting structure of the APU. It is attached to the APU firewall around the
forward end. The APU is hung from the plenum by tubular supports and vibration isolators. The plenum flange
seal is a compressible seal between the compressor inlet plenum of the APU and the intake plenum. A removable
access panel, located on the forward end, allows for inspection of the interior of the plenum.
IF 49-6
757 General Familiarization (7-2005)
ATA 49 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 49 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU DRAINS AND VENTS
Overview
This section describes the system of drains and vents for the removal of waste fluids that accumulate in the APU.
(Fig. 1)
APU Drains & Vents (Fig. 2)
Waste fuel, oil, and water are removed from the APU by gravity-fed drains and vents:
1. Rainwater seepage into the intake plenum drains overboard on the left side of the APU compartment
through the APU plenum drain.
2. Rainwater seepage from the air intake housing and duct drains overboard on the right side of the airplane.
3. Oil from the manual fill port drains into the APU compartment through the oil scupper drain.
4. Water accumulating in the inlet plenum drains into the APU compartment through the inlet plenum drain.
5. Five drains discharge through the drain mast mounted on the APU access door. Fuel and oil drain from the
fuel pump, oil pump, and IGV actuator. Excess fuel from aborted or wet starts drains through the turbine
plenum drain. The drain for the bearing seal cavity drains the redundant seals of the APU engine. Water
drains from the heat shield. Fuel from the fuel divider flows into the drain tank; then in-flight air pressure
purges the tank through the drain mast.
6. ON APUs WITH TELL TALE DRAINS; tell tale drains are installed on the fuel control unit/oil pump/IGV
actuator and drain lines for the bearing seal cavity. Tell tale drains are fluid traps which can be used to
isolate drain line leakage and to determine the leakage rates of these drains.
757 General Familiarization (7-2005)
ATA 49 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-7
757 General Familiarization (7-2005)
ATA 49 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-7A
757 General Familiarization (7-2005)
ATA 49 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-7B
757 General Familiarization (7-2005)
ATA 49 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU AND GENERATOR LUBRICATION SYSTEM
Overview
The APU and generator lubrication system lubricates, cools, and scavenges the electrical generator and all
bearings and gears in the engine. The system consists of the oil pump, generator scavenge pump, the scavenge
pump for the rear bearing, magnetic chip detectors, drain plug, oil cooler, oil filters, and several valves. The
lubrication system is monitored by the oil indicating system consisting of an oil temperature sensor, a switch for
the low oil pressure, an oil quantity transmitter, and on the APUs with Garrett SB 49-5438, a switch for the low oil
temperature.
The APU lubrication system includes an oil reservoir, an oil pump assembly, a de-oil solenoid valve, a switch for
the low oil temperature, a gearbox pressurization system, an oil cooler, and a scavenge system for the turbine
bearing.
757 General Familiarization (7-2005)
ATA 49 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-8
757 General Familiarization (7-2005)
ATA 49 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU AND GENERATOR LUBRICATION SYSTEM COMPONENTS
Oil Reservoir
The oil reservoir stores the oil supply for the system. It is located at the bottom of the gearbox and has a capacity
of approximately 6.2 quarts. A manual fill port, with a scupper drain, and pressure fill provisions are included. An
approximate oil reservoir level is read through a sight glass with markings indicating SAFE and ADD. An oil
quantity transmitter, located on the bottom of the gearbox, sends a signal to the Engine Indication and Crew
Alerting System (EICAS), and stores an oil quantity indication in the status and maintenance modes. The reservoir
drain plug is located on the bottom of the gearbox. The plug contains a magnetic chip detector which can be
inspected for possible gearbox damage without draining the reservoir.
APUs WITH A MAGNETIC CHIP DETECTOR FOR THE GENERATOR SCAVENGE
(APUs WITH GARRETT SB 49-5564); a magnetic chip detector for generator scavenge oil and a switch for the
low oil temperature are located on the right side of the gearbox.
IF 49-9
757 General Familiarization (7-2005)
ATA 49 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Oil Pump Assembly
The oil pump supplies oil under pressure to the main rotor bearings, to the bearings and gears in the gearbox, and
to the generator lubrication system. It also scavenges oil from the compressor bearings. The pump is pilot
mounted and bolted to the gearbox on the left side. The pressure and scavenge gear pumps are driven by a
common pump shaft from the accessory drive gearbox. The pressure pump has a pressure relief valve to prevent
over-pressurization beyond 200 ±5 PSI. The oil pump filter filters the oil flowing to the APU and generator
lubrication system. It is a disposable filter with pleated fiberglass in screw-on housing. An indicator for the filter
differential pressure pops out at 20 ±5 PSID to indicate when the filter is clogged. The pressure regulator valve
regulates the oil pressure to 65 ±5 PSIG. The oil pump housing provides a mounting base for the fuel control unit.
IF 49-10
757 General Familiarization (7-2005)
ATA 49 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
De-Oil Solenoid Valve
The de-oil solenoid is an electrically-actuated pneumatic valve bolted to the gearbox case. The valve de-oils the oil
pump to reduce viscous oil drag during cold starts.
APU WITH THE LOW OIL TEMPERATURE SWITCH; the switch for the low oil temperature activates the solenoid,
through the starter motor, when the oil temperature in the gearbox is approximately 25°F (-4°C). The APU control
unit also activates the solenoid during an APU shutdown or when the start relay is energized and the oil
temperature at the oil temperature sensor is less than 0°F (-18°C). The solenoid is deactivated when the starter
motor operation stops or when the APU speed is 50%.
APU WITHOUT THE LOW OIL TEMPERATURE SWITCH; the APU control unit activates the solenoid during an
APU shutdown or when the start relay is energized and the oil temperature at the oil temperature sensor is less
than 0°F (-18°C). The solenoid is deactivated when the starter motor operation stops or when the APU speed is
50%.
Switch for the Low Oil Temperature (APUs with the switch for the low oil temperature)
The switch for the low oil temperature is installed on the bottom of the gearbox. The switch closes, to de-oil the
system, when oil in the reservoir is below 25 ± °F, and opens at 40°F.
757 General Familiarization (7-2005)
ATA 49 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-11
757 General Familiarization (7-2005)
ATA 49 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Gearbox Pressurization System
The gearbox pressurization system regulates the air flow which is a buffer for the seal and pressurizes the gearbox
at varying altitudes to maintain oil pump performance. The air flow keeps oil from leaking past the primary seal
into the load compressor. The shuttle check valve works with the gearbox shutoff valve to control the air supply.
The air flow comes from first stage compression below 15,000 feet and second stage compression above 15,000
feet. Second stage compression supplies gearbox pressurization above 15,000 feet. The regulating valve for the
gearbox pressure maintains the pressure at 4.5 PSID. If the gearbox pressurization system fails to regulate at 4.5
PSID, the regulating valve will vent the gearbox to ambient at a maximum of 10 PSID.
Gearbox ventilation is included in the gearbox pressurization system. An air-oil separator centrifugally separates
the oil from the air. Gearbox pressure is vented to the APU exhaust through the regulating valve for the gearbox
pressure at all altitudes, and at 90-40% during every roll-down.
IF 49-12
757 General Familiarization (7-2005)
ATA 49 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Oil Cooler
The oil cooler is an oil-air heat exchanger consisting of a rectangular core of plate-fin design. Air from the inlet
plenum is blown through the oil cooler by a cooling fan that is operated by the gearbox. The bypass valve for the
oil cooler is a thermally actuated and bolted to the oil cooler header. When the oil is below 170°F the valve is open
allowing the oil to bypass the cooler.
IF 49-13
757 General Familiarization (7-2005)
ATA 49 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APUs WITH THE MAGNETIC CHIP DETECTORS FOR THE TURBINE BEARING AND THE
COMPRESSOR BEARING
The turbine bearing and compressor bearing scavenge return lines each has a chip detector which can be
inspected for possible bearing damage.
IF 49-14
757 General Familiarization (7-2005)
ATA 49 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Generator Scavenge Pump
The generator scavenge pump feeds generator cooling oil through the generator scavenge filter before the oil
enters the oil reservoir. The pump is bolted to the gearbox case underneath the generator drive pad.
Filter Assembly for the Generator Scavenge Oil
The filter for the generator scavenge oil is a disposable fiberglass filter of the same design as the oil pump filter.
It has a differential pressure indicator which activates at 20 ±5 PSID. An electrical switch for the differential
pressure initiates an APU automatic shutdown when the differential pressure across the filter is 35 ±5 PSID and
the oil temperature is greater than 115 ±25°F (46 ±14°C).
IF 49-15
757 General Familiarization (7-2005)
ATA 49 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU AND GENERATOR LUBRICATION SYSTEM OIL SERVICING
The oil reservoir is on the bottom of the gearbox. The reservoir has a quantity of 6.2 quarts. A manual fill port
(with a scupper drain) is on the left side of the gearbox. There are also pressure fill fittings that are aft of the
manual fill port.
You can use EICAS or the sight gage to do a check of the APU oil level. On some APUs, the sight gage can be used
with the APU off or with the APU in operation. The sight gage is aft of the oil fill port on the left side of the APU.
There is an oil quantity transmitter on the APU that shows the oil quantity on EICAS. The transmitter sends signals
to EICAS, which shows FULL, 0.75, 0.50, 0.25, or ADD on the PERF/APU display. When the oil level is 4.25 qts., or
less for 60 seconds (with the engine stopped), the transmitter sends a signal to EICAS and the APU OIL QTY
message shows.
Access Location Zones:
162 Bulk Cargo Compartment (Right)
315 APU Compartment (Left)
316 APU Compartment (Right)
NOTE:
Drain the APU oil when the engine is hot. If it is necessary, operate the APU (AMM 49-1100/201) until the APU is hot. Then perform the APU shutdown before changing APU oil.
Make sure the APU control switch is in the OFF position and attach a DO-NOT-OPERATE tag.
Open this circuit breaker on the overhead circuit breaker panel, P11, and attach a DO-NOT-CLOSE tag: 11B34, APU
MN BAT CONT or APU ALTN CONT
Open this circuit breaker on the E6 rack in the aft equipment center and attach a DO-NOT-CLOSE tag: APU CONT
Open the APU access doors 315AL & 316AR
Engage the support rods for the APU access doors.
757 General Familiarization (7-2005)
ATA 49 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Do a Check of the APU Oil Level.
APU and Generator Lubrication Oil Servicing (Continued):
Use one of these procedures to do a check of the oil level in the APU:
Use the EICAS display to do a check of the oil level. Supply electrical power. Look for the APU OIL QTY message
on the EICAS STATUS display.
If the APU OIL QTY message is on EICAS, fill the APU oil.
ON AN OIL SIGHT GAGE WITHOUT A FULL MARK FOR THE APU ON; use the sight gage when the APU is off
to do a check of the oil level.
Look at the oil level in the sight gage. If the oil level is below the ADD on the sight gage for the APU OFF, fill the
APU oil.
757 General Familiarization (7-2005)
ATA 49 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU and Generator Lubrication Oil Servicing (Continued):
WARNING:
DO NOT REMOVE THE OIL FILL CAP WHEN THE APU IS IN OPERATION. IF THE OIL FILL
CAP IS REMOVED WHEN THE APU OPERATES, INJURIES TO PERSONS OR DAMAGE TO
EQUIPMENT CAN OCCUR.
ON AN OIL SIGHT GAGE WITH A FULL MARK FOR THE APU ON; use the sight gage when the APU operates to
do a check of the oil level.
NOTE:
You can only do a check of the APU oil level when the engine operates if the sight gage
has a mark for the APU ON.
Fill the APU Oil (Gravity Fill Method)
If possible, stop the APU for four hours to let the temperature decrease before you fill the oil.
Make sure the APU control switch is in OFF position and attach a DO-NOT-OPERATE tag.
Open this circuit breaker on the overhead panel P11 and attach a DO-NOT-CLOSE tag: 11B34, APU MN BAT CONT
or APU ALTN CONT
Open this circuit breaker on the E6 rack in the aft equipment center and attach a DO-NOT-CLOSE tag: APU CONT
Clean the oil fill cap before it is removed.
757 General Familiarization (7-2005)
ATA 49 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU and Generator Lubrication Oil Servicing (Continued):
WARNING:
DO NOT REMOVE THE OIL FILL CAP IF THE APU IS WARM, AND THE OIL LEVEL IS AT OR
ABOVE THE FULL MARK. THE HOT OIL CAN CAUSE INJURY.
Remove the oil fill cap.
WARNING:
DO NOT KEEP THE OIL ON YOUR SKIN. IF YOU DO NOT CLEAN THE OIL OFF YOUR SKIN,
THE OIL CAN CAUSE INJURY.
CAUTION:
DO NOT MIX TWO TYPES OF OIL (MIL-L-7808 TYPE I AND MIL-L-23699 TYPE II) WHEN
YOU ADD THE OIL IN THE APU. IT IS PERMITTED TO MIX DIFFERENT BRANDS OF OIL
WITH THE SAME TYPE OF OIL WHEN YOU ADD THE OIL IN THE APU. A MIXTURE OF THE
TWO TYPES OF OIL IN THE APU CAN CAUSE DAMAGE TO THE APU. IF YOU DO NOT
CLEAN THE OIL OFF, THE OIL CAN CAUSE A STAIN ON YOUR CLOTHES AND PAINT CAN
BECOME SOFT. DO NOT PUT TOO MUCH OIL IN THE RESERVOIR OR YOU CAN CAUSE
THE APU TO HAVE A SHUTDOWN FROM LOW OIL PRESSURE.
Slowly add oil until the oil flows into the scupper drain.
Install the oil fill cap. Make sure the oil fill cap is tightened.
Fill APU Oil (Pressure Fill Method)
Prepare APU for oil service as stated in the gravity fill method.
Remove the caps from the oil pressure fittings. Clean the oil pressure fittings.
757 General Familiarization (7-2005)
ATA 49 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU and Generator Lubrication Oil Servicing (Continued):
WARNING:
DO NOT KEEP THE OIL ON YOUR SKIN FOR A LONG TIME. IF YOU DO NOT CLEAN THE
OIL OFF YOUR SKIN, THE OIL CAN CAUSE INJURY.
CAUTION:
DO NOT MIX THE TWO TYPES OF OIL (MIL-L-7808 TYPE I AND MIL-L-23699 TYPE II)
WHEN YOU ADD THE OIL IN THE APU. IT IS PERMITTED TO MIX DIFFERENT BRANDS OF
OIL WITH THE SAME TYPE OF OIL WHEN YOU ADD THE OIL IN THE APU. A MIXTURE OF
THE TWO TYPES OF OIL IN THE APU CAN CAUSE DAMAGE TO THE APU. IF YOU DO NOT
CLEAN THE OIL OFF, THE OIL CAN CAUSE A STAIN ON YOUR CLOTHES AND PAINT CAN
BECOME SOFT. DO NOT PUT TOO MUCH OIL IN THE RESERVOIR OR YOU CAN CAUSE
THE APU TO HAVE A SHUTDOWN FROM LOW OIL PRESSURE.
Connect the oil supply and the oil overflow hoses to the pressure fittings.
Slowly add oil until you see oil in the overflow hose.
When the oil from the overflow hose is at a slow drip, remove the pressure fill hoses.
Install the caps on the pressure fittings.
Put the Airplane Back to Its Usual Condition.
If the APU OIL QTY message is shown on EICAS, do the Maintenance Message Erase Procedure.
Close the APU access doors.
Remove the DO-NOT-CLOSE tag and close this circuit breaker on the E6 rack in the aft equipment center: APU
CONT.
Remove the DO-NOT-CLOSE tag and close this circuit breaker on the overhead panel P11: 11B34, APU MN BAT
CONT or APU ALTN CONT.
Remove the DO-NOT-OPERATE tag from the APU master control switch.
757 General Familiarization (7-2005)
ATA 49 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 49 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU ENGINE FUEL SYSTEM
Overview
The engine fuel system of the APU provides pressurized metered fuel to the combustion chamber and pressurized
fuel to the actuator for the inlet guide vanes. The main fuel tank on the left side of the airplane supplies the fuel.
The system consists of a fuel control unit, a flow divider, nozzles and manifolds. Electrical power for the system is
28-volts DC.
Fuel is pumped to the fuel control unit of the APU by a primary or secondary fuel pump in the left airplane wing.
The ac fuel boost pump on the left forward side is used when bus power for ac ground service is available. When
no ac ground power is available, the dc fuel pump of the APU is used. The fuel shutoff valve for the APU, located
in the left airplane wing, opens with an APU ON signal. Fuel flows back to the fuel control unit where it is
pressurized and metered. The fuel passes through the fuel flow divider which directs the fuel to the primary and
secondary fuel nozzles.
757 General Familiarization (7-2005)
ATA 49 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-16
757 General Familiarization (7-2005)
ATA 49 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL SYSTEM COMPONENTS
Fuel Control Unit
The fuel control unit pressurizes and meters fuel going to the fuel flow divider. It also provides pressurized fuel to
the actuator for the inlet guide vanes. The fuel control unit consists of the following in fuel flow order: inlet filter,
fuel pump, high-pressure relief valve, high pressure filter, actuator pressure regulator, fuel metering assembly,
pressurizing valve, and shutoff solenoid valve for the fuel. The unit is mounted to the front of the oil pump
assembly by a QAD clamp. Fuel is drained from the unit through a drain tube connected to its underside.
Inlet Filter
The inlet filter is located at the inlet of the fuel pump bolted to the fuel control unit. It is a 10 micron disposable
filter element. A delta pressure indicator pops out at 5 ±0.5 PSID to indicate when the fuel filter is clogged. A filter
bypass valve activates to allow fuel to flow around a plugged inlet filter.
Fuel Pump
The fuel pump provides pressurized fuel to the metering valve. The pump is driven by an oil lubricated splined
coupling from the oil pump assembly. Double carbon face seals prevent contamination between oil and fuel. A
drain between the seals allows a quick check of seal conditions.
High-Pressure Relief Valve
The high-pressure relief valve is located at the fuel pump outlet. It is a spring-loaded ball design to protect against
over-pressurization of the system.
High Pressure Filter
The high pressure filter is located at the outlet port of the fuel pump. It protects the fuel metering valve and the
actuator for the inlet guide vanes from minute particles due to normal pump wear. The filter is a stainless steel
screen, which you can clean, and is attached to a removable plug.
Actuator Pressure Regulator
Power to the actuator for the inlet guide vanes is supplied by pressurized fuel from the fuel control unit. The
actuator pressure regulator provides hydraulic pressurized fuel for the IGV actuator.
757 General Familiarization (7-2005)
ATA 49 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-17
757 General Familiarization (7-2005)
ATA 49 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fuel Metering Assembly
The fuel metering assembly meters the output of the fuel control unit to the engine for acceleration and governed
speed operation. It consists of a torque-motor metering valve and a differential pressure regulator. These
components are contained within the housing for the fuel control unit. The APU control unit electronically controls
the torque-motor metering valve which adjusts the fuel flow rate for the required governed speed. The differential
pressure regulator maintains a constant differential pressure across the metering valve.
Pressurizing Valve
The pressurizing valve maintains the out-going fuel at a constant pressure during the early phase of the start
cycle. It is located at the inlet to the shutoff solenoid valve.
Fuel Shutoff Solenoid Valve
The shutoff solenoid valve for the fuel shuts off fuel to the flow divider. It is located at the outlet of the fuel
control unit. It is a normally spring-loaded closed, direct acting, three-way valve. A 28-volt DC source energizes the
valve open with a signal from the APU control unit. Fuel flow cools the valve.
757 General Familiarization (7-2005)
ATA 49 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flow Divider
The flow divider distributes fuel from the fuel control unit to the primary and secondary nozzles. It consists of a
screen filter, run and start sequence valves, solenoid valve, primary drain valve, and secondary drain valve. The
flow divider is mounted on the APU housing behind the surge control valve. At the inlet is a self-bypassing filter
screen. The primary path for the fuel flow is open throughout the APU start and run operation. The secondary fuel
path, however, is restricted by start and run sequence valves which open at preset fuel pressures.
The run sequence valve allows additional fuel to enter the secondary manifold after the solenoid valve opens at
95% speed. The run sequence valve has a lower setting than the start sequence valve which allows proper fuel
atomization at lower supply pressure. Primary and secondary drain valves allow fuel to drain from the hoses and
nozzles after fuel flow to the divider has stopped.
IF 49-18
757 General Familiarization (7-2005)
ATA 49 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Nozzles
The nozzles atomize and inject fuel into the engine combustor. There are six primary nozzles for starting and APU
operation and six secondary nozzles for running and high altitude operation. The nozzles are connected to
manifolds which receive fuel from the fuel metering valve. They are equally spaced around the engine periphery.
The primary nozzles are designed for high atomizing ability during the low pressures at low flow rate. This helps
with engine light-off during the start cycle. The secondary nozzles are brought into use when a high flow rate
exists.
Manifolds
The manifolds provide a fuel flow path from the flow divider to the fuel nozzles. The manifolds are insulated
flexible lines.
IF 49-19
757 General Familiarization (7-2005)
ATA 49 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU IGNITION/STARTING SYSTEM
Overview
The APU ignition and starting provides the means of accelerating the APU engine to starting speed and igniting the
fuel-air mixture in the combustor. The system consists of a starter motor, ignition unit, igniter plug, and an igniter
lead. The Transformer Rectifier Unit (TRU) provides 28-volts DC to the system. The APU battery serves as a
backup power source. The ignition and starting system is automatically controlled by the APU control unit.
757 General Familiarization (7-2005)
ATA 49 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU IGNITION/STARTING SYSTEM COMPONENTS
Ignition Unit
The ignition unit converts battery power into high voltage current required to produce a capacitive spark at the
igniter plug. It has a 28-volts DC input and an 18 KV output. The unit consists of a transformer, electronic vibrator,
rectifier, booster coil, and series of capacitors all enclosed in a single case. It is located below the mounting
bracket on the right rear side.
Ignition Lead
The ignition lead conducts the output current of the ignition unit to the igniter plug. The lead consists of an
insulated electrical conductor encased in a braided copper conduit with insulated, threaded connectors at each end.
The lead and connectors are shielded to prevent radio interference.
Igniter
The igniter provides a high energy spark for igniting the fuel-air mixture in the engine combustor. It consists of an
outer casing, a center electrode made of tungsten alloy, a ceramic insulator, and a hastelloy X tip. It is located on
the right side of the combustor case.
757 General Familiarization (7-2005)
ATA 49 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-20
757 General Familiarization (7-2005)
ATA 49 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Starter Motor
The starter motor has four brushes and four poles. The starter
arter is powered by 28-volts DC from the TRU or the APU
battery. An indicator peg for brush wear is located under a transparent window. The starter motor is mounted with
a V-band clamp on the upper left side of the accessory gearbox.
IF 49-21
757 General Familiarization (7-2005)
ATA 49 - Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU COOLING AIR SYSTEM
Overview
APUs PRE-ALLIEDSIGNAL-SB 48-7391; the cooling air system provides forced air cooling to the APU engine
compartment and engine lubricating oil of the APU. The cooling air system consists of an isolation valve for the
cooling fan, cooling fan, associated ducts, and an air/oil heat exchanger (oil cooler).
APU POST-ALLIEDSIGNAL-SB 48-7391; the cooling air system provides forced air cooling to the APU engine
compartment and engine lubricating oil of the APU. The cooling air system consists of a cooling fan, associated
ducts, and an air/oil heat exchanger (oil cooler).
APU PRE-ALLIEDSIGNAL-SB 49-7391; cooling air is drawn from the APU inlet plenum through the fan isolation
valve and the fan inlet duct by the cooling fan. The fan circulates some air through the oil cooler and vents the rest
to the APU engine compartment. The oil cooler air is expelled overboard through the discharge duct for the oil
cooling air.
APU POST-ALLIEDSIGNAL-SB 49-7391; cooling air is drawn from the APU inlet plenum through the fan inlet
duct by the cooling fan. The fan circulates some air through the oil cooler and vents the rest to the APU engine
compartment. The oil cooler air is expelled overboard through the discharge duct for the oil cooling air.
757 General Familiarization (7-2005)
ATA 49 - Page 45
For Training Purposes Only
IF 49-22
757 General Familiarization (7-2005)
ATA 49 - Page 46
AeroEd LLC – Aviation Education Resource
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-22A
757 General Familiarization (7-2005)
ATA 49 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU COOLING AIR SYSTEM COMPONENTS
Fan Isolation Valve
APU PRE-ALLIEDSIGNAL-SB 49-7391; the isolation valve for the cooling fan is a butterfly valve that is
pneumatically powered, spring-loaded closed, and opens whenn the APU is started. The valve is installed between
the cooling air flange of the air inlet plenum and the cooling air fan. It consists of a shaft-mounted butterfly, a
pneumatic actuator, and an open-position indicator switch enclosed in a housing.
APU WITH THE FAN ISOLATION VALVE CONNECTED TO THE SURGE VALVE; the valve is pneumatically
actuated when the pressure of the compressor discharge air (PCD) rises above 7.5 PSIG. The position indicator
switch supplies valve-open indication to the APU control unit. The pneumatic actuator consists of a cylinder,
piston, piston rod, diaphragm, and spring. The actuator piston is spring-loaded to the retracted (valve closed)
position.
APU WITH THE FAN ISOLATION VALVE CONNECTED TO THE FIRST STAGE COMPRESSOR; the valve is
pneumatically actuated when the pressure of the compressor discharge air (PCD) rises above 5.0 PSIG. The
position indicator switch supplies valve-open indication to the APU control unit. The pneumatic actuator consists of
a cylinder, piston, piston rod, diaphragm, and spring. The actuator piston is spring-loaded to the retracted (valve
closed) position.
IF 49-23
757 General Familiarization (7-2005)
ATA 49 - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Cooling Fan
The cooling air fan provides cooling air to the oil cooler and the APU compartment. It consists of a fan inlet
housing, fan rotor and a fan stator assembly. It has a 37 blade axial rotor with 31 stator vanes. It is powered by
the APU gearbox and provides 90 lb/min of cooling air. It has forward ball bearings
gs and aft roller bearings.
IF 49-24
757 General Familiarization (7-2005)
ATA 49 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Oil Cooling Air Duct
The engine oil is cooled in an air/oil heat exchanger. The duct for the oil cooling air exhausts the oil cooler air
overboard through an opening on the left side of the airplane.
Compartment Cooling and Ventilation
Compartment cooling air of the APU is discharged through a port located between
een the cooling fan and oil cooler. To
prevent over-pressurization of the compartment, a louvered vent is located on the left side of the airplane skin.
The vent door blows open at 2 PSID.
IF 49-25
757 General Familiarization (7-2005)
ATA 49 - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU EXHAUST GAS TEMPERATURE INDICATING SYSTEM & COMPONENTS
Overview
The indicating system for the exhaust gas temperature (EGT)
GT) of the APU measures the temperature of the exhaust
gas and provides EGT indication on the flight deck. The system consists of thermocouple probes, APU control unit,
and EICAS indication. Power for the system is 28 volts dc from the main battery bus.
Thermocouple Assemblies
APU exhaust gas temperature is sensed by four thermocouple assemblies equally spaced aft of the power section.
A pair of thermocouples, enclosed in an inconel support tube with a stainless steel header, makes up each
assembly. A pair of assemblies make up a rake.
IF 49-26
757 General Familiarization (7-2005)
ATA 49 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU TIME TOTALIZER SYSTEM & COMPONENTS
Overview
The time totalizer (hourmeter) and the cyclemeter measure the amount of APU operation. The meters are located
in the aft equipment center on the E6 rack. Power is supplied to both indicators from the APU battery bus.
IF 49-27
757 General Familiarization (7-2005)
ATA 49 - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU OIL INDICATING SYSTEM
Overview
The oil indicating system of the APU monitors the engine lubrication system. The system consists of the oil
temperature sensor, the switch for the low oil pressure, the differential pressure switch for the generator oil filter,
and oil quantity transmitter. The oil indicating circuit obtains 28-volts DC for operation from the battery bus. The
high oil temperature, low oil pressure, and the differential pressure circuits for the generator oil filter are part of
the automatic shutdown circuit.
757 General Familiarization (7-2005)
ATA 49 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU OIL INDICATING SYSTEM COMPONENTS
Low Oil Pressure Switch
The switch consists of a diaphragm, a belleville spring, a shunt disc, and an electrical connection. It is installed in
a common housing with the oil temperature sensor locatedd on the case for the APU load compressor. The APU
control unit shuts down the APU when the switch signal is below 35 | 5 PSIG for more than 15 seconds at speeds
greater than 95 percent. Automatic shutdowns due to low oil pressure are stored in BITE memory.
Oil Temperature Sensor
The sensor consists of a variable resistance bulb and an electrical connector. The switch is installed in a common
housing with the switch for the low oil pressure located on the load compressor case. At speeds greater than 95
percent, an oil temperature signal to the APU control unit greater than 310° (± 10°) F for more than 10 seconds
activates protective shutdown. The cause of shutdown is stored in BITE memory.
IF 49-28
757 General Familiarization (7-2005)
ATA 49 - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Oil Quantity Transmitter
The transmitter consists of a magnet/float that triggers a series of magnetic proximity switches which gives a
signal to EICAS. EICAS messages read "FULL", "0.75",
.75", "0.50", "0.25", and "ADD". The transmitter is installed
through the bottom of the gearbox. When the quantity reaches "ADD" an additional EICAS "APU OIL QTY" message
also appears.
Differential Pressure Switch for the Generator Oil Filter
The switch is located within the housing for the generator scavenge filter. It consists of a shunt disc, belleville
spring, and a housing with electrical connection. The switch senses a differential pressure across the filter, which
indicates a clogged filter. The control unit shuts down the APU when a signal is received for more than 0.5 second
at speeds above 95 percent. A shutdown due to a clogged filter is stored in BITE memory.
IF 49-29
757 General Familiarization (7-2005)
ATA 49 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU CONTROL SYSTEM
Overview
The APU controls consist of manual and automatic controls for starting, stopping, and maintaining the APU within
safe limits during operation. Once an APU start is initiated, further control of the unit is fully automatic. The
control system consists of the APU mounted sensors which feed inputs into the Electronic Control Unit for the APU.
The controller, in turn, feeds signals to APU mounted torque-motors and information to the APU control panel in
the flight compartment and EICAS. Electrical power for the system is 28-volts DC.
If Airesearch SB 2117342-49-2195 has been done to the electronic control unit (ECU), the APU will have altitude
start enhancements and start protection for low oil pressure. With the changed hardware, the starter will disengage
at 42% RPM on the ground and 55% RPM at altitudes above 36,000 feet (10973 meters). Also, when the APU is
initially started with low oil pressure, the APU will shutdown 15.5 seconds after it gets to 95% RPM (or one
second if the APU oil temperature is greater than 20°F). The subsequent start with low oil pressure will cause a
shutdown one second after the APU gets to 95% RPM. The APU will then not start if there is low oil pressure.
757 General Familiarization (7-2005)
ATA 49 - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU CONTROL SYSTEM COMPONENTS
APU Control Panel - P5 Panel
The APU control panel is located on the overhead panel and contains the master control switch for the APU FAULT
light and RUN light. The master control switch is a three-position switch with a momentary START position. The
amber FAULT light illuminates with a failure to start or a fault causing APU shutdown. FAULT also illuminates
momentarily while the fuel valve is in transit. The white RUN light illuminates when APU reaches 95% speed.
Power Controls - P5 Panel
The bleed air switch for the APU is located on overhead panel P5. It opens the shutoff valve for the APU bleed air
to allow APU bleed air to flow to the airplane cabin. The generator control switch for the APU is also located on
overhead panel P5. It excites the APU generator field and supplies power to the generator control unit for the APU.
The APU must be at 95% speed for either pneumatic or electrical loading.
The APU Indications show on EICAS as follows:
1. The APU FAULT message is shown as a level C alert message when the APU has a protection shutdown.
2. The APU DOOR message is shown on the status and the maintenance displays of EICAS when the door is
not in the commanded position.
3. AIRPLANES WITH THE APU CONTROL UNIT -18 AND BEFORE; the APU BITE message shows on the
EICAS maintenance display when one or more of these faulty units is in the memory of the APU control
unit (ECU): #1 SPD SENSOR, #2 SPD SENSOR, LOP SWITCH, #1 T/C RAKE, #2 T/C RAKE, DEOIL SOL,
FLOW DIV SOL, HOT SENSOR, P2 SENSOR, FAN VALVE, FILTER SWITCH.
4. AIRPLANES WITH THE APU CONTROL UNIT -19 AND SUBSEQUENT; the APU BITE message shows
on the EICAS maintenance display when one or more of these faulty units is in the memory of the APU
control unit (ECU): #1 SPD SENSOR, #2 SPD SENSOR, LOP SWITCH, EGT #1 CIRCUIT, EGT #2 CIRCUIT,
DEOIL SOL, FLOW DIV SOL, HOT SENSOR, P2 SENSOR, FAN VALVE, FILTER SW (GEN).
5. The OIL Q shows the APU oil quantity on the status display as FULL, 0.75, 0.50, 0.25, or ADD. When the oil
quantity shows add, the APU OIL QTY message will show on the status and the maintenance displays of
EICAS. The APU OIL QTY message shows when the APU will operate for approximately 30 hours at the oil
consumption rate that is usual.
6. The APU RPM shows the engine speed on the EICAS maintenance display.
7. The APU EGT shows the exhaust gas temperature on the status and the maintenance displays of EICAS.
757 General Familiarization (7-2005)
ATA 49 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
P62 Control Panel for the APU
The APU control panel is on the right side of the outer cylinder for the nose landing gear. This is the only location
from which you can do an APU shutdown other than the flight compartment. This panel can do an APU shutdown,
arm the fire protection system, and operate the fire bottle. The APU shutdown switch is to be used during
emergencies only.
IF 49-30
757 General Familiarization (7-2005)
ATA 49 - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-30A
757 General Familiarization (7-2005)
ATA 49 - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Control Unit
The electronic control unit (ECU) is a digital control system based on a microprocessor. It governs the engine
starting sequence, acceleration, governed speed operation, operation within temperature limits, the control of the
inlet guide vanes, surge valve control, and the engine shutdown. The input signals to the control unit are utilized
by it to supply control signals to various accessories. The input signals from the APU are engine speed, exhaust
gas temperature, air inlet temperature and pressure, high oil temperature, low oil pressure, the position of the fan
isolation valve, the position of the air intake door, the actuator position of the inlet guide vanes, differential
pressure switch for the generator filter, and transducer signals for the differential and total pressures. Signals
from the aircraft are the shutoff valve for the bleed air, environmental control system (ECS), air/ground, and main
engine start (MES). The control unit has two separate systems for automatic shutdown protection: normal
shutdown logic (hardware shutdown) and redundant shutdown logic (software shutdown). The APU control unit is
located on the aft equipment center in the E6 rack and panel type mounting.
AIRPLANES WITH THE APU CONTROL UNIT -18 AND BEFORE; the ECU supplies an integral BITE (Built-In
Test Equipment) function to help the line maintenance technician do APU trouble-shooting. The ECU continuously
monitors and stores data about APU shutdowns and failures of Line Replaceable Units (LRUs). The BITE memory
can only be interrogated after APU shutdown on the ground. It recalls the stored data, shows the causes for up to
five protective shutdowns and gives LRU failure indications.
AIRPLANES WITH THE APU CONTROL UNIT -19 AND SUBSEQUENT; the ECU supplies an integral BITE
(Built-In Test Equipment) function to help the line maintenance technician do APU trouble-shooting. The ECU
continuously monitors and stores data about APU shutdowns and failures of Line Replaceable Units (LRUs). The
BITE memory can be interrogated while the APU operates or after shutdown. It recalls the stored data, shows the
causes for up to four protective shutdowns and gives LRU failure indications. The APU must be shutdown for the
ECU to perform the self test.
757 General Familiarization (7-2005)
ATA 49 - Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-31
757 General Familiarization (7-2005)
ATA 49 - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Monopole
The monopole is a variable reluctance transducer which is magnetic and does not touch the parts. It converts
mechanical motion into an electrical signal. A ferromagnetic nut mounted on the APU drive shaft produces an
electrical pulse in the monopole every time it passes. Monopole assemblies on the left and right sides of the inlet
plenum transmit electrical signals to the APU control unit.
IF 49-32
757 General Familiarization (7-2005)
ATA 49 - Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Inlet Temperature Sensor
The inlet temperature of the load compressor (LCIT) sensor is a chromel-alumel unit for rapid response to
temperature changes. The thermocouple senses temperature increases associated with ambient changes, reverse
flow conditions, and surges in the load compressor. The APU control unit will shut down the engine if the inlet
temperature exceeds 210°F or increases more than 60°F in 1.1 seconds. The sensor is located forward of the inlet
plenum by the actuator upper mount for the inlet guide vanes.
APU Inlet Pressure Sensor
The inlet pressure (P2) sensor is a solid state transducer which is piezo-resistive and has a sensing element which
is a silicon strain gage. It provides a 0-5 volt DC signal proportional to the inlet pressure. The APU control unit
uses its signal to modify the engine fuel schedule. The sensor is mounted on the left side near the top of the inlet
plenum.
IF 49-33
757 General Familiarization (7-2005)
ATA 49 - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU OPERATION PROCEDURES START & SHUTDOWN
APU Start
When you operate the APU, make sure it is in the specified limits listed in the maintenance manual.
Make sure there are no free objects near the APU air inlet.
Set the main battery switch on the P5 Overhead Panel to ON.
Set the APU BLEED AIR VALVE switch on the P5 Overhead Panel to ON.
Do a test of the fire detection system for the APU.
Start and operate the APU:
WARNING:
APU BEFORE SERIAL NUMBER P-1172 PRE-GARRETT-SB 3862160-49-5716; DO NOT GO
INTO THE APU COMPARTMENT WHEN THE APU IS IN OPERATION. THE COOLING FAN
CAN HAVE A FAILURE AND CAUSE INJURY TO PERSONS.
757 General Familiarization (7-2005)
ATA 49 - Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Start (Continued):
CAUTION:
MAKE SURE THE APU ACCESS PANELS ARE CLOSED. A FIRE DETECTOR IS INSTALLED
ON THE ACCESS PANEL AND A FIRE CAN OCCUR WITHOUT AN INDICATION WHEN THE
ACCESS PANELS ARE OPEN. IF THE ACCESS PANELS MUST BE OPEN, MAKE SURE A
PERSON IS ON THE GROUND TO LOOK FOR A FIRE. THE PERSON ON THE GROUND CAN
TELL THE PERSONS IN THE CONTROL CABIN WHEN A FIRE OCCURS. THE PERSON ON
THE GROUND COULD ALSO USE THE EMERGENCY SHUTDOWN ON THE NOSE LANDING
GEAR PANEL P62.
Turn the APU control switch on the P5 Overhead Panel to START and release it to ON. Make sure the RUN light
flashes twice.
NOTE:
The RUN light flashes when the self-test for the APU control unit is complete.
Make sure the FAULT light comes on and goes off.
NOTE:
The FAULT comes on when the APU fuel valve opens.
Monitor the EGT and the RPM on the EICAS MAINTENANCE display.
NOTE:
AC power must be supplied to get the EICAS indications.
Make sure the EGT increases at 7% rpm when the igniters energize.
Make sure the RUN light comes on when the APU is at 95% rpm.
NOTE:
You can supply pneumatic and electrical loads when the APU rpm is more than 95%.
If the APU acceleration times are not in the specified limits, the APU will have a shutdown and the APU FAULT
light will come on.
757 General Familiarization (7-2005)
ATA 49 - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Start (Continued):
If the APU does not start, do these steps:
CAUTION:
MAKE SURE YOU OBEY THE STARTER DUTY CYCLE IN THE APU OPERATION LIMITS. IF
YOU DO NOT OBEY THESE INSTRUCTIONS, YOU CAN CAUSE DAMAGE TO THE APU.
1. Turn the APU control switch on the P5 panel to OFF.
2. Stop for 1 minute to let the fuel drain.
3. Turn the APU control switch to START and release it to ON.
4. If the does not start a third time, look for the APU BITE message on the EICAS MAINTENANCE display.
5. If the APU BITE message was on EICAS, do the APU Control Unit BITE procedure (AMM 49-61-05/201).
NOTE:
If the APU has a shutdown a third time 15 seconds after it gets to 95% rpm, look for the
LOP light during the BITE procedure.
ECU WITH LOW OIL PRESSURE PROTECTION; the start will be prevented if there was low oil pressure after two
starts. If the LOP light was on, make sure the LOP failure is corrected before you try to start the APU again.
757 General Familiarization (7-2005)
ATA 49 - Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Shutdown
There are four different types of APU shutdowns. These shutdowns are as follows:
1. A usual APU shutdown which uses the APU control switch.
2. An emergency APU shutdown which uses the pilots fire switch on the P8 control stand panel.
3. An emergency APU shutdown which uses the APU shutdown switch on P62 panel on the nose landing gear.
4. A protection shutdown which is done by the APU control unit.
This task gives the steps to do the first three types of shutdowns. The APU control unit automatically does the
protection shutdown.
1.
Perform the usual APU shutdown with the APU control switch:
Turn the APU control switch on the P5 panel to OFF.
a) If the APU has a pneumatic load, the shutdown will occur after 60 seconds.
b) If the APU does not have a pneumatic load, the shutdown will occur immediately.
c) When the APU is at 15% rpm, the inlet door and the fuel valve will close.
NOTE: The FAULT light will come on while the fuel valve closes.
757 General Familiarization (7-2005)
ATA 49 - Page 67
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Shutdown (Continued):
2.
Do an emergency shutdown with the APU fire switch on the P8 control stand panel:
CAUTION:
USE THIS EMERGENCY SHUTDOWN PROCEDURE CAREFULLY. DO NOT TURN THE FIRE
SWITCH UNLESS THERE IS A FIRE. IF YOU TURN THE FIRE SWITCH, THE FIRE BOTTLE
CONTENTS WILL BE RELEASED.
Pull the fire switch out to do the APU shutdown.
If there is a fire, turn the fire switch to release the fire bottle contents.
When the APU stops, turn the APU control switch on the P5 panel to OFF.
Turn and push the fire switch to put the switch back in its usual position.
757 General Familiarization (7-2005)
ATA 49 - Page 68
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Shutdown (Continued):
3.
Do an emergency shutdown with the APU auto-shutdown switch on the P62 panel which is on the nose
landing gear:
Push the APU shutdown switch on the P62 panel to do the shutdown.
NOTE:
The ARM light on the P62 panel will come on. This shows that the fire bottles are
prepared to release their contents.
If there is a fire, push the APU BOTTLE DISCHARGE switch or switches on the P62 panel to release the fire bottle
contents.
NOTE:
AIRPLANES WITH DUAL FIRE BOTTLES; There are two switches on the P62 panel.
When the APU stops, turn the APU control switch on the P5 panel to OFF.
Set the main battery switch to OFF and then to ON.
NOTE:
This will set the emergency shutdown system again. You can also open and close the APU
REMOTE FIRE IND (11B33) on the P11 overhead panel to set the shutdown system.
757 General Familiarization (7-2005)
ATA 49 - Page 69
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
757 GENERAL FAMILIARIZATION
ATA 52
DOORS ........................................................................................................................................................................ 3
NO. 1, 2 OR 4 PASSENGER DOORS OVERVIEW ....................................................................................................... 3
NO. 1, 2 OR 4 PASSENGER DOOR .......................................................................................................................... 4
Lining ................................................................................................................................................................. 4
Door Torque Tube .............................................................................................................................................. 6
Hinge Arms ........................................................................................................................................................ 6
Guide Arm .......................................................................................................................................................... 6
Body Torque Tube .............................................................................................................................................. 8
Rotary Snubber ................................................................................................................................................... 8
Latch Mechanism ............................................................................................................................................. 10
Latch Torque Tube ........................................................................................................................................... 10
Handle Mechanism ........................................................................................................................................... 12
Girt Bar Mechanism ......................................................................................................................................... 14
Door Emergency Powered Opening System ...................................................................................................... 16
Emergency Power Reservoir ............................................................................................................................. 16
Emergency Power Actuator and Emergency Power Chain ................................................................................. 16
NORMAL DOOR OPENING FROM OUTSIDE .......................................................................................................... 18
NORMAL DOOR OPENING FROM INSIDE .............................................................................................................. 20
NORMAL DOOR CLOSING FROM INSIDE .............................................................................................................. 22
NORMAL DOOR CLOSING FROM OUTSIDE ........................................................................................................... 23
EMERGENCY DOOR OPERATION........................................................................................................................... 24
NO. 3 EMERGENCY EXIT DOOR ................................................................................................................................ 26
Overview........................................................................................................................................................... 26
NO. 3 EMERGENCY EXIT DOOR ............................................................................................................................ 26
No. 3 Emergency Exit Mechanism .................................................................................................................... 26
NO. 1 AND NO. 2 CARGO DOORS.............................................................................................................................. 30
Overview........................................................................................................................................................... 30
NO. 1 AND NO. 2 CARGO DOORS COMPONENTS ..................................................................................................... 32
No. 1 and No. 2 Cargo Door Latch Mechanism ................................................................................................. 32
No. 1 and 2 Cargo Door Hinge Linkage............................................................................................................. 35
Hinge Drive Mechanism.................................................................................................................................... 35
Hinge Drive Power Unit .................................................................................................................................... 35
Manual Drive Gearbox/Flexible Drive Shaft ..................................................................................................... 36
Cargo Door Actuation and Door Warning Proximity Sensors ............................................................................ 39
Cargo Door Motor Interlock Switch................................................................................................................... 39
NORMAL DOOR OPENING & CLOSING ...................................................................................................................... 42
Electrical door opening ......................................................................................................................................... 42
Electrical Door Closing ..................................................................................................................................... 43
EQUIPMENT COMPARTMENT DOORS ...................................................................................................................... 44
757 General Familiarization (7-2005)
ATA 52 - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Overview........................................................................................................................................................... 44
EQUIPMENT COMPARTMENT DOORS COMPONENTS .............................................................................................. 45
Forward Access Door........................................................................................................................................ 45
Electrical/Electronics Access Door................................................................................................................... 46
Air-Conditioning Access Doors ......................................................................................................................... 47
Service Access Door ......................................................................................................................................... 48
Elevator Controls Access Door.......................................................................................................................... 49
APU Access Doors ............................................................................................................................................ 50
Vertical Fin Access Door (Airplanes with SB 53-38) ........................................................................................ 51
Flight Compartment Door ................................................................................................................................. 52
DOOR WARNING SYSTEM ......................................................................................................................................... 55
Overview........................................................................................................................................................... 55
DOOR WARNING SYSTEM COMPONENTS ................................................................................................................ 56
Door Warning Sensors ...................................................................................................................................... 56
Door Warning Annunciator Lights ..................................................................................................................... 56
EICAS Messages ............................................................................................................................................... 58
757 General Familiarization (7-2005)
ATA 52 - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DOORS
NO. 1, 2 OR 4 PASSENGER DOORS OVERVIEW
There are three passenger doors on each side of the fuselage. Usually, the doors on the left side of the fuselage
are for passenger and crew entry. The doors on the right side of the fuselage are for use by service personnel.
IF 52-1
757 General Familiarization (7-2005)
ATA 52 - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NO. 1, 2 OR 4 PASSENGER DOOR
The passenger doors are hinged, outward-opening, and pressure-sealing. The doors operate from inside or outside
the airplane.
The doors have upper and lower gates which unfold when the doors are closed and latched. When the door is
unlatched, the gates fold inward to allow the door to fit through the door opening. The gate pushrods connect the
gate to the latch torque tube. The gates operate when either handle is rotated to latch/unlatch the door.
The doors have a door torque tube mounted on the door. A pushrod connects the door torque tube to the handle
box. The door torque tube opens the door to the cocked position when either handle is rotated to unlatch the door.
The door torque tube connects to the body torque tube by two hinge arms and a guide arm. The door pivots about
the body torque tube on the hinge arms. The guide arm on the body torque tube controls the arc through which the
door rotates.
The body torque tube is mounted on fuselage structure forward of each door. After the door is cocked, the door
rotates about the body torque tube to the fully open position. A rotary snubber on the body torque tube slows the
rotation of the door during emergency door opening.
The doors are latched closed by four latch roller cranks. The latch roller cranks are mounted on the latch torque
tubes. The latch torque tubes are connected to the handle box by latch pushrods. The latch roller cranks rotate to
engage or disengage the latch rollers in the latch cams mounted on the door sill. Either handle operates the latch
torque tube to latch/unlatch the door.
The interior or exterior handle operates the latch mechanism on the door. Either handle operates the handle box
mechanism, which controls the latch mechanism.
The emergency mechanism activates the doors for emergency operation (powered door opening and escape system
deployment). The emergency mechanism is activated or deactivated from inside the airplane using the mode
selector lever. When the door is opened from outside the airplane, the door emergency mechanism is automatically
deactivated. The inadvertent slide deployment mechanism prevents accidental escape system deployment.
Lining
The door has a lining to cover its mechanisms and the escape system. An assist handle is mounted on the lining to
assist in closing the door from the inside. A viewing window in the lower lining allows visual access to the
inflation bottle pressure gage. A prismatic window in the upper lining allows the crew to see outside.
The lower lining is hinged at its upper edge. It hinges open during emergency operation so the escape system can
deploy.
The upper lining is hinged at its lower edge. It hinges open to allow access for maintenance.
757 General Familiarization (7-2005)
ATA 52 - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-2
757 General Familiarization (7-2005)
ATA 52 - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Torque Tube
The door torque tube is mounted on the forward side of the door. The door torque tube assembly consists of the
upper hinge arm, cocking bell crank, coupler, torque shaft, and lower hinge arm. The door torque tube assembly is
mounted in five bearings.
The cocking bell crank has internal splines on each end. The upper hinge arm has an external spline which is
inserted into the upper end of the cocking bell crank. The upper coupler shaft has an external spline which is
inserted into the lower end of the cocking bell crank.
The torque shaft has internal splines on each end. The lower coupler shaft has an external spline which is inserted
into the upper end of the torque shaft. The lower hinge arm has an external spline which is inserted into the lower
end of the torque shaft.
The door torque tube rotates in five bearings. When the door handle is rotated to unlatch the door, the door torque
tube rotates the door to the cocked position.
Hinge Arms
Two hinge arms connect the door torque tube to the body torque tube. The hinge arms are splined to the door
hinge torque tube. The upper hinge arm is connected to the upper hinge link on the body torque tube. The lower
hinge arm is connected to the lower hinge bell crank on the body torque tube.
The hinge arms support the door when it is unlatched or open. When the door is being opened or closed, it rotates
on the hinge arms about the body torque tube.
Guide Arm
The guide arm controls the motion of the door when it is opened or closed. The guide arm connects to a clevis
attached to the lower hinge fitting. The guide arm is attached to the lower hinge bell crank and is connected to
fuselage structure by two pins trapped between two cam plates.
The guide arm elbow pin and follower pin move between cam plates attached to fuselage structure beneath the
body torque tube. When the door is being opened or closed, the pins follow the tracks in the cam plates to control
the rotation of the door.
757 General Familiarization (7-2005)
ATA 52 - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-3
757 General Familiarization (7-2005)
ATA 52 - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Body Torque Tube
The body torque tube is mounted on fuselage structure forward of the door. The body torque tube assembly
consists of the upper hinge link, upper torque shaft, hold-open lock, rotary snubber, sprocket assembly,
counterbalance assembly, and lower hinge bell crank. The body torque tube is mounted in six bearings and on
bushings.
The upper hinge link has an external spline which is inserted into the upper end of the upper torque shaft. The
upper torque shaft is inserted through the hold-open lock. The upper end of the snubber shaft has an external
spline which is inserted into the lower end of the upper torque shaft.
The lower hinge bell crank has an external spline which is inserted into the lower end of the counterbalance
assembly. The lower end of the snubber shaft is inserted through the sprocket assembly. The external spline of the
lower snubber shaft is inserted into the upper end of the counterbalance assembly.
AIRPLANES WITH SPIRAL SPRING COUNTERBALANCE; the springs are mounted to a spring support fitting on
the fuselage structure forward of the body torque tube. The upper and lower springs wrap around the splined
sleeve in opposite directions.
AIRPLANES WITH SPRING CYLINDER COUNTERBALANCE; the counterbalance consists of a camshaft, roller
which is attached to the piston and plug, and spring cylinder. The spring consists of two springs in compression
(one inside the other), piston and plug, and yoke.
The body torque tube rotates in six bearings. The body torque tube is the axis about which the door rotates open
or closed.
The sprocket assembly is part of the emergency 0pening system. An emergency power chain engages with the
sprocket. During emergency opening, the chain rotates the body torque tube to power the door open.
The hold-open lock holds the door fully open. The hold-open lock is manually released to close the door.
Rotary Snubber
The rotary snubber is part of the body torque tube assembly. The rotary snubber has an upper shaft with an
external spline which is inserted into the upper torque shaft. The rotary snubber has a lower shaft which is
inserted through the sprocket assembly. The lower snubber shaft has an external spline which is inserted into the
counterbalance assembly.
The rotary snubber slows the rotation of the door during emergency operation to prevent damage to the door, door
hinges, or fuselage.
757 General Familiarization (7-2005)
ATA 52 - Page 8
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-4
757 General Familiarization (7-2005)
ATA 52 - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Latch Mechanism
The latch mechanism latches or unlatches the door and folds the upper and lower gates inward so the door can fit
through the door opening. The door is latched when the latch rollers engage the latch cams mounted on the body.
The latch rollers engage the latch cams when either handle is rotated to latch the door.
The latch mechanism is mounted on the door. The interior or exterior handle rotates a cam in the handle box. Cam
follower cranks operate latch pushrods which are connected to the latch torque tubes. The latch torque tubes rotate
to engage the latch rollers with the latch cams mounted on the frame to latch the door.
The upper and lower gates are operated by gate pushrods. The gate pushrods are connected to the latch torque
tubes by a gate crank. When either handle is rotated to unlatch the door, the gate crank rotates and folds the gate
inward.
The cocking pushrod from the handle box to the door torque tube rotates the door torque tube to move the door to
the cocked position. The cocking pushrod is operated by a cam follower crank in the handle box. The cam is
turned when either handle is rotated to latch or unlatch the door.
Latch Torque Tube
There are two latch torque tubes on each door. Each latch torque tube has a latch roller crank on each end. Each
latch torque tube has a gate crank and a torque tube crank.
When the latch torque tube rotates, the latch roller crank rotates to engage the latch roller into the latch cam
mounted on the frame to latch the door.
The gate crank is connected to the gate by the gate pushrod. When the latch torque tube rotates, the gate crank
rotates to operate the gate.
The upper torque tube crank is connected to the handle box by a latch pushrod. The lower torque tube crank is
connected to the handle box by latch pushrods through an idler crank. When either handle is rotated to latch or
unlatch the door, the latch pushrods from the handle box cause the latch torque tubes to rotate.
757 General Familiarization (7-2005)
ATA 52 - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-5
757 General Familiarization (7-2005)
ATA 52 - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Handle Mechanism
The handle mechanism is mounted on the door. The handle mechanism consists of the handle box assembly and
the interior and exterior operating handles.
The handle box is attached to door structure. The handle box mechanisms control the latch mechanism and the
door cocking motion. When the door is opened using the exterior handle, the handle box mechanism automatically
disarms the system.
The latch crank operates the latch pushrods of the latch mechanism. The latching crank is connected to the latch
roller crank cam follower external spline. The rotation of the latch roller crank cam follower is controlled by the
control cam.
The cocking crank operates the latch pushrod to the door torque tube. The cocking crank is connected to the
cocking roller crank cam follower external spline. The rotation of the cocking roller crank cam follower is
controlled by the control cam.
The control cam rotates when either handle is used to latch or unlatch the door. As the control cam rotates, the
latch crank and the cocking crank rotate to unlatch and cock the door.
The interior handle is connected directly to the control cam by the stub shaft. Operation of the interior handle will
not deactivate the escape system. From the inside, the escape system must be deactivated using the mode selector
lever.
The exterior handle does not operate when the door is opened using the interior handle. The exterior handle is
bolted to the outer shaft and the sliding center shaft. The outer shaft rotates in the basket cam. The basket cam
has two slots which allow the exterior handle to retract flush in the vertical position (door latched or unlatched)
only.
When the exterior handle is flared (butterflied), the escape system is automatically deactivated. The center shaft
slides inward through the outer shaft as the handle is flared. A yoke attached to the center shaft operates the mode
selector lever to deactivate the escape system.
When the exterior handle is flared, the center shaft slides inward over the inner shaft and engages the clutch
collar, which is attached to the inner shaft. When the exterior handle is rotated, the center shaft rotates, and
causes the inner shaft and the interior handle to rotate with the exterior handle.
757 General Familiarization (7-2005)
ATA 52 - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-6
757 General Familiarization (7-2005)
ATA 52 - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Girt Bar Mechanism
The girt bar mechanism operates with the mode selector lever. When the mode selector lever is placed to ENGAGE,
the girt bar sliders remain engaged in the floor fittings, which hold the girt bar in place on the floor. When the
mode selector lever is placed to DETACH, the sliders move to free the girt bar from the floor fittings. The sliders
engage with the girt bar lifters which are connected to the lower gate. As the lower gate folds inward during door
unlatching, the girt bar is picked up.
A lockout mechanism prevents changing the mode selector lever position after the door is opened. The lockout
mechanism prevents damage to the emergency mechanism. A lockout cam on the lower door latch torque tube
prevents movement of the emergency mechanism unless the door is closed and latched. When the door is closed
and latched, the lockout cam rotates out of the way of the crank stop on the girt bar mechanism, allowing
movement of the emergency mechanism. When the door is unlatched, the lockout cam rotates to block movement
of the crank on the girt bar mechanism.
757 General Familiarization (7-2005)
ATA 52 - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-7
757 General Familiarization (7-2005)
ATA 52 - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Door Emergency Powered Opening System
The door emergency powered opening system is actuated when the door is opened from inside the airplane with
the mode selector lever in ENGAGE. The door is powered open from slightly past the cocked position by pneumatic
actuation.
The door emergency powered opening system components are mounted on fuselage structure forward of the door.
A drive chain is engaged with a sprocket on the body torque tube. The chain is driven by a pneumatic actuator.
Pneumatic pressure for the actuator comes from the pneumatic reservoir. Reservoir pressure is released when a
diaphragm on the reservoir is ruptured. The trigger mechanism actuates a knife to rupture the diaphragm.
The trigger mechanism consists of an emergency power lever, cable and pulley system, and spring cylinder,
mounted on the body; and a spring cartridge, roller arm, and trigger, mounted on the upper hinge arm. When the
door is opened from inside in the ENGAGE mode, the trigger (held in place by the trigger pin) moves the roller arm
to actuate the emergency power lever. The emergency power lever pulls on the cable which actuates the emergency
power reservoir to release pneumatic pressure to the actuator. The actuator powers the drive chain to rotate the
body torque tube and open the door.
Emergency Power Reservoir
The emergency power reservoir supplies pneumatic pressure to the emergency power actuator through tubing. The
reservoir is mounted in a cradle over the door, and secured to the cradle with clamps. The trigger cable actuates
the toggle lever on the reservoir. The toggle lever depresses the reservoir knife to rupture a diaphragm and release
the pressure to the actuator.
Emergency Power Actuator and Emergency Power Chain
The emergency power actuator pulls the emergency power chain to rotate the body torque tube and power the door
open. Pneumatic pressure from the emergency power reservoir drives the actuator.
757 General Familiarization (7-2005)
ATA 52 - Page 16
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-8
757 General Familiarization (7-2005)
ATA 52 - Page 17
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NORMAL DOOR OPENING FROM OUTSIDE
Push the handle latch and hold the ends of the two exterior handles.
Pull the external handles out of the handle recess to the completely flared (butterflied) position.
NOTE:
The door emergency systems are automatically deactivated when the exterior handles are
flared.
WARNING:
MAKE SURE THE INTERIOR HANDLE WILL NOT HIT PERSONS WHEN IT MOVES. THE
INTERIOR HANDLE FOLLOWS THE MOVEMENT OF THE EXTERIOR HANDLES, AND CAN
CAUSE INJURY TO PERSONS IN THE AIRPLANE.
CAUTION:
IF THE EXTERIOR HANDLES ARE ROTATED PRIOR TO COMPLETE FLARING, THE BEARING
MOUNT ASSEMBLY CAN BE DAMAGED OR SEVERED FROM THE HANDLE.
Slowly turn the exterior handle no more than 90 degrees toward the open position.
WARNING:
MAKE SURE THE GIRT BAR IS NOT IN VIEW BELOW THE RETRACTED LOWER GATE.
FAILURE TO OBEY CAN CAUSE THE ESCAPE SYSTEM TO INFLATE, WHICH CAN CAUSE
INJURY OR DAMAGE.
Make sure the girt bar is not in view below the retracted lower gate. If the girt bar is in view, turn the exterior
handles to the close position. Do not continue to open the door until the defect in the girt bar operation is
corrected.
757 General Familiarization (7-2005)
ATA 52 - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Normal Door Opening from Outside (Continued):
Continue to slowly turn the exterior handles no more than 30 degrees more (total of 120 degrees from the closed
position).
WARNING:
MAKE SURE THE EMERGENCY OPERATION TRIGGER MECHANISM IS NOT IN THE
POSITION TO OPERATE. FAILURE TO OBEY CAN CAUSE THE OPERATION OF THE
EMERGENCY POWER SYSTEM, WHICH CAN CAUSE INJURY OR DAMAGE.
When the door is open sufficiently to see the emergency operation trigger mechanism, make sure the trigger
mechanism is not in the position to operate. If the trigger mechanism is in the position to operate, turn the
external handle to the close position. Do not continue to open the door until the defect in the trigger mechanism
operation is corrected.
Continue to slowly turn the exterior handles 60 degrees more (total of 180 degrees) to move the door to the cocked
position.
NOTE:
The cocked position is the position where the door is open approximately 30 degrees from
the closed position.
Release the exterior handles and push them back against the door.
CAUTION:
MAKE SURE YOU OPERATE THE DOOR SMOOTHLY. FAILURE TO OPERATE THE DOOR
SMOOTHLY CAN CAUSE DAMAGE TO THE DOOR GUIDE ARM ADJUSTER RODS.
Use the assist handle on the aft door frame to push the door out through the doorway to the fully open position.
Make sure the hold-open lock is engaged.
757 General Familiarization (7-2005)
ATA 52 - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NORMAL DOOR OPENING FROM INSIDE
Before opening the door, the door emergency systems must be deactivated. This is done by moving the mode
selector lever from ENGAGE to DETACH. This action disengages the girt bar from the floor retainers to deactivate
the escape system and retracts the trigger pin to deactivate the emergency powered door opening system. The
retractable door above the interior handle will retract into the door lining. The girt bar engaged warning light will
turn off.
Initial rotation of the interior handle controls the door latching system. The upper and lower door latch torque
tubes rotate to disengage the latch rollers from the latch cams and operate the gate pushrods to fold the upper and
lower gates inward.
757 General Familiarization (7-2005)
ATA 52 - Page 20
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Normal Door Opening from Inside (Continued):
Further rotation of the interior handle controls the door cocking motion. The handle rotation causes the door torque
tube to rotate. As the door torque tube rotates, the forward edge of the door angles inboard; the aft edge of the
door is pulled slightly inboard and forward.
From the cocked position, the door is swung through the door opening. The door arc is controlled by the guide arm
pins tracking in the cam plates as the hinge arms pivot about the door torque tube.
AIRPLANES WITH SPIRAL SPRING COUNTERBALANCE; the upper spring provides a force to help rotate the
door from the cocked position to the open position.
AIRPLANES WITH SPRING CYLINDER COUNTERBALANCE; the spring cylinder against the camshaft provides a
force to help rotate the door from the cocked position to the open position.
The rotary snubber on the body torque tube slows the rotation of the door to prevent damage. When the door is
fully open, the hold-open lock holds the door open and prevents the door from contacting the fuselage.
757 General Familiarization (7-2005)
ATA 52 - Page 21
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NORMAL DOOR CLOSING FROM INSIDE
Before the door is closed, the hold-open lock must be disengaged. The door is swung through the door opening to
the cocked position. The spring cylinder counterbalance (or the lower spring on airplanes with two spiral
counterbalance springs) provides a force to help rotate the door from the open position to the closed position. An
assist handle on the door can be grasped to pull the door into the cocked position.
757 General Familiarization (7-2005)
ATA 52 - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Normal Door Closing from Inside (Continued):
From the cocked position, initial rotation of the interior handle causes the door torque tube to rotate. This moves
the door into the door opening.
Further rotation of the interior handle operates the upper and lower latch torque tubes, which unfold the upper and
lower gates, and engage the latch rollers in the latch cams.
The door emergency system must now be activated. This is done by moving the mode selector lever from DETACH
to ENGAGE. This action engages the girt bar with the floor retainers to activate the escape system and extends the
trigger pin to activate the emergency powered door opening system. The retractable door above the interior handle
will extend to prevent inadvertent handle operation. The girt bar engaged warning light will turn on.
NORMAL DOOR CLOSING FROM OUTSIDE
Before the door is closed, the hold-open lock must be disengaged.
To close the door from the outside, the door is moved to the cocked position. From the cocked position, the
external handles are flared from the recess and rotated 180 degrees. This closes and latches the door. The exterior
handles must then be stowed in the recess.
Exterior handle rotation results in the same actions of door mechanisms as the interior handle rotation. As the
exterior handle is rotated to the door latched position, the interior handle also rotates to the door latched position.
757 General Familiarization (7-2005)
ATA 52 - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EMERGENCY DOOR OPERATION
Emergency door operation can be initiated only from inside the airplane since exterior operation automatically
deactivates the emergency systems.
When the door is opened with the mode selector lever in ENGAGE, the door is powered open and the escape
system is deployed. The trigger pin actuates the emergency power reservoir to supply pressure to the actuator to
power the door open. As the door opens, the escape system deploys because the girt bar was left attached to the
floor. The girt pulls the escape pack out of the lower bustle. As the escape pack falls, the inflation begins.
757 General Familiarization (7-2005)
ATA 52 - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 52 - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NO. 3 EMERGENCY EXIT DOOR
Overview
There is one No. 3 emergency exit door on each side of the fuselage, just aft of the wing. The emergency exit doors
are intended for use only during an emergency evacuation of the passenger compartment. The emergency exit
doors are always armed to deploy the escape slides when an emergency exit door is opened.
NO. 3 EMERGENCY EXIT DOOR
The No. 3 emergency exit door is a plug-type door, hinged at the bottom with two hinges. The emergency exit door
opens out and down to a position where it is hanging from its hinges, upside-down, below the door cutout.
An escape slide and pack board are mounted in a compartment on the interior of the emergency exit door. The
escape slide is automatically deployed when the door is opened - from inside or outside. When the door opens, the
escape slide is released from the door and inflated.
A lining on the interior of the emergency exit door covers the escape slide and the door mechanisms. There is a
slide inflation bottle view door on the slide compartment cover. The interior handle is flush with the lining, in a
recess. A prismatic window is on the door, with a cutout in the lining to allow outside viewing. Two release buttons
are at the lower edge of the lining, on each side of the slide cover.
The exterior handle or interior handle can be used to open the emergency exit door. Operation of either handle
causes the same actions of the door mechanisms. As the handle is operated, the door torque tube rotates. The
rotation of the torque tube opens the pressure relief door immediately and then drives the latch cranks to lift the
emergency exit. The emergency exit door is lifted 1.5 inches for the guide pins to move clear of the guide fittings.
This lifting action frees the door to rotate outboard.
No. 3 Emergency Exit Mechanism
The interior or exterior handle operates a pushrod and crank linkage to rotate the torque tube. The torque tube has
a latch crank at each end, and a pressure relief door actuating cam toward the forward end. The pressure relief
door actuating cam connects to the pressure relief door by a control rod.
Two hinge assemblies are on the lower door sill structure. The emergency exit door is attached to the hinge
assembly by the snubber links and the snubber fitting.
The snubber is connected to the snubber fitting and the upper idler at the upper end of the snubber, and
connected to the snubber links at the lower end of the snubber.
757 General Familiarization (7-2005)
ATA 52 - Page 26
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
No. 3 Emergency Exit Mechanism (Continued):
The spring-loaded release positions a pawl to hold a pin against the cam surface of the lower idler. The lower idler
has a notch in its cam surface. When the door has been lifted 1.3 inches, the pin will drop into the notch in the
cam surface. This prevents the door from dropping down to its original position after it has been lifted. The
spring-loaded release button can be pressed to release the pin from the notch. This allows the emergency exit door
to drop down to its closed position.
757 General Familiarization (7-2005)
ATA 52 - Page 27
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
No. 3 Emergency Exit Mechanism (Continued):
The interior or exterior handle operates a pushrod and crank linkage to rotate the torque tube. The torque tube has
a latch crank at each end, and a pressure relief door actuating cam toward the forward end. The pressure relief
door actuating cam connects to the pressure relief door by a control rod.
Two hinge assemblies are on the lower door sill structure. The emergency exit door is attached to the hinge
assembly by the snubber links and the snubber fitting.
The snubber is connected to the snubber fitting and the upper idler at the upper end of the snubber, and
connected to the snubber links at the lower end of the snubber.
The spring-loaded release positions a pawl to hold a pin against the cam surface of the lower idler. The lower idler
has a notch in its cam surface. When the door has been lifted 1.3 inches, the pin will drop into the notch in the
cam surface. This prevents the door from dropping down to its original position after it has been lifted. The
spring-loaded release button can be pressed to release the pin from the notch. This allows the emergency exit door
to drop down to its closed position.
757 General Familiarization (7-2005)
ATA 52 - Page 28
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 52 - Page 29
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NO. 1 AND NO. 2 CARGO DOORS
Overview
The No. 1 and No. 2 cargo doors are similar in design and share many common components.
The No. 1 cargo door is located on the right side of the lower fuselage at body station 590. It provides a clear
opening 55.0 inches wide by 42.5 inches high, and is approximately 97 to 105 inches above the ground.
The No. 2 cargo door is located on the right side of the lower fuselage at body station 1410. It provides a clear
opening 55.0 inches wide by 44.0 inches high, and is approximately 91 to 99 inches above the ground.
No. 1 and 2 cargo doors are conventional riveted/bolted sheet metal construction. Each door is a structural box
with access to mechanisms and electrical wiring through openings in the inner skin.
The cargo door is a hinged, plug-type door. It is unlatched manually, from outside or inside the airplane and
opened electrically or manually from outside or inside the airplane. It can be latched only from outside the
airplane. Cabin pressure load on the door is transmitted to body structure by six stops at the forward and aft
edges, and two stops at the upper sill.
During unlatching, the latch cranks react with latch tracks on the fuselage structure cutout to lift the door 1-1/2
inches. Lifting the door unlatches the two upper door stop fittings and raises the other door stop fittings to clear
the stop pads on the fuselage structure.
Each door is attached to the fuselage by two hinge arms and two interconnected hinge linkages. The hinge arms
are driven by the electrically powered hinge drive mechanism, which includes a power unit, two mechanical rotary
actuators, and drive shafts. Guide rollers on the forward and aft edges guide the door to the latched position.
A pressure seal is installed around the fuselage cutout structure. Seal depressors on the door contact the seal
when the door is closed, to prevent cabin pressure leakage and moisture penetration. H. Lights are installed on the
door to aid cargo loading operations.
The door may be left open in any position in winds up to 65 knots without structural damage. In winds exceeding
65 knots initial damage could occur to the door hinge linkage. The hinge drive power unit is designed to operate in
winds up to 40 knots. In winds exceeding 40 knots, the power unit may not have sufficient power to operate the
door. This would be indicated by hinge drive power unit clutch slippage.
757 General Familiarization (7-2005)
ATA 52 - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
757 General Familiarization (7-2005)
ATA 52 - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NO. 1 AND NO. 2 CARGO DOORS COMPONENTS
No. 1 and No. 2 Cargo Door Latch Mechanism
The door latch mechanism includes a door handle mechanism which controls the door lift-latch mechanism from
either inside or outside the airplane. The inside handle can only be used to unlatch and lift the door. Operation of
the handle, through pushrods, bell cranks and a torque tube, rotates two latch cranks, opens or closes two
pressure vent doors, rotates two lift cams and rotates two upper stop fittings.
757 General Familiarization (7-2005)
ATA 52 - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
No. 1 and No. 2 Cargo Door Latch Mechanism (Continued):
The latch cranks are located on the torque tube at the forward and aft door edges. Lockout pawls adjacent to the
latch cranks prevent closing of the latching mechanism when the door is unlatched and partially open. The torque
tube runs fore and aft through the door assembly and slightly above the door center line. The pressure vent doors
also mechanically attach to the torque tube inside of the latch cranks. The lift cams are located adjacent to the
pressure vent doors and on the torque tube. The upper stop fittings are connected through bell cranks and
pushrods to the torque tube. The upper stop fittings are located above the door upper edge beam. Over center
springs are provided on both the handle mechanism and torque tube.
757 General Familiarization (7-2005)
ATA 52 - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 52 - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
No. 1 and 2 Cargo Door Hinge Linkage
Each hinge linkage mechanism (two per door) consists of a hinge arm, scissors arm, two hinge links, upper and
lower lift links, a program rod, adjustable rod and adjustment arm.
Hinge Drive Mechanism
The hinge drive mechanism consists of the hinge drive power unit, two drive shafts and two rotary actuators. The
hinge drive power unit, which has an auxiliary manual drive, operates the drive shafts connected to the two hinge
drive rotary actuators. The drive shaft engages the rotary actuator input shaft, and the rotary actuator output shaft
engages the hinge arm. The rotary actuator housing is mounted on the hinge support structure. The rotary
actuators (planetary gear reduction units) are mounted on the hinge support structure. The auxiliary manual drive
consists of a manual drive socket on the forward side of the hinge drive power unit, a flexible shaft, and a manual
drive gearbox connected to fuselage skin. The flexible shaft connects the manual drive gearbox to the hinge drive
power unit.
Hinge Drive Power Unit
The hinge drive power unit provides the motive force to open and close the cargo door. One power unit is installed
on each cargo door hinge drive. The power unit consists of a 3-phase, 115/200-volts, 400-Hz electrical motor and
a reduction gearbox. The gearbox transmits torque to two splined drive shafts which connect to rotary actuators.
The power unit receives electrical power form either ground power or the onboard airplane APU. The flexible drive
from the manual drive gearbox is attached to the power unit so that the cargo door can be opened and closed
manually or by the use of power-driven tools in the event of electrical motor failure or loss of electrical power. The
unit turns in either direction upon either manual or electrical command. Driving the cargo door from fully closed to
fully open, or from fully open to fully closed, requires 59.0 revolutions of the output shaft (850 revolutions of
manual drive input).
During normal operating conditions (total output torque of 180 inch-pounds maximum) the unit can electrically
open the door from fully closed to fully opened within 75 seconds. During maximum operating conditions (output
torque of 300 inch-pounds), the unit can electrically open the door from fully closed to fully open within 30
seconds. The power unit contains a device to limit the output torque from a minimum of 300 inch-pounds to a
maximum of 400 inch-pounds. The electric motor incorporates a thermal protection device which prevents overcurrent damage to the motor and windings. It prevents steady state or transient temperatures of the motor core
from exceeding 390°F (199°C). The electric motor will not start when supplied with two phases of power only, but
will remain running if one phase is lost during operation. The power unit incorporates a holding brake which is
activated by the loss of at least two phases of power to the electrical connector. The holding brake is deactivated
by application of three phases of power to the electrical connector.
Two rotary actuators are installed on each cargo door hinge mechanism. The rotary actuators transfer the torque
output from the hinge drive power unit to the cargo door hinge linkage. The rotary actuator is essentially a gearbox
with an input-to-output gear ratio of 185 to 1. The actuator will turn in either direction depending upon input
rotation. Torque applied to the output end of the actuator will not drive the actuator. Thus the cargo doors will not
fall and close due to their own weight after electrical power loss or power unit failure.
757 General Familiarization (7-2005)
ATA 52 - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Manual Drive Gearbox/Flexible Drive Shaft
A manual drive gearbox is attached to fuselage skin just forward of each cargo door. The gearbox is attached to a
flexible drive shaft which in turn is connected to the hinge drive power unit. The gearbox/flexible drive shaft
provides the means for opening and closing the cargo doors manually or by power-driven tools in the event of
electrical power loss or hinge drive power unit failure. The input-to-output gear ratio is 1 to 1.
The gearbox contains a 3/8-inch square receptacle for accepting a manual or power-driven tool. The flexible drive
shaft connected between the manual drive gearbox and hinge drive power unit gearbox consists of a corrosion
resistant steel flexible inner core and end fittings. Approximately 850 revolutions of the flexible cable are required
to either open or close a cargo door. Maximum torque required for door opening is approximately 80 inch-pounds.
Maximum torque required for door closing is approximately 60 pound-inches.
757 General Familiarization (7-2005)
ATA 52 - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-9
757 General Familiarization (7-2005)
ATA 52 - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 52 - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Cargo Door Actuation and Door Warning Proximity Sensors
Proximity sensors are included at the door periphery which provides door warning signal input to the flight
compartment Engine Indicating and Crew Alerting System (EICAS) and signal input to cut off power to the hinge
drive system when the door reaches the desired up or down position.
NOTE:
Prox sensors, S10083 and S10088 are part of the door warning system.
For airplanes with -17 and on PSEU, prox sensors S10353 and S10360 are inoperative and do not require any
adjustment.
Cargo Door Motor Interlock Switch
AIRCRAFT WITHOUT SB 52-31; a cargo door motor interlock micro switch in the door open control circuit is
mounted on the fuselage upper sill. The switch is held in the open position by a switch striker on door moveable
upper aft stop fitting when the door is latched, in order to prevent operation of door open motor in the hinge drive
power unit. When door is unlatched, the switch striker moves allowing the switch to close, and complete the circuit
between door control switch and door open motor to permit door opening operation with door control switch.
AIRCRAFT WITH SB 52-31; a cargo door motor interlock switch is mounted in the door, just below the latch crank
torque tube. When the door is latched, a switch striker on the torque tube holds the switch open, preventing door
opening operation. When the door is unlatched, the switch striker rotates away from the switch, causing the switch
to close and thereby allowing door opening operation.
757 General Familiarization (7-2005)
ATA 52 - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 49-10
757 General Familiarization (7-2005)
ATA 52 - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-10A
757 General Familiarization (7-2005)
ATA 52 - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NORMAL DOOR OPENING & CLOSING
Electrical door opening
CAUTION:
DO NOT LEAVE DOOR OPEN IN WINDS EXCEEDING 65 KNOTS. STRUCTURAL DAMAGE
MAY OCCUR.
NOTE:
The cargo door can be manually unlatched, and powered open from inside of the airplane,
but the door can only be closed and latched from outside of airplane.
Provide electrical power.
Pull internal or external latch handle to unlatch door (external handle moves downward approximately 105 degrees)
CAUTION:
TO AVOID MOTOR OVERHEAT DAMAGE, LIMIT DOOR OPERATION TO TWO COMPLETE
OPEN/CLOSE CYCLES IN 5 MINUTES.
Place and hold inside or outside door control switch in the OPEN position until door is fully open.
NOTE:
Door motion can be stopped or reversed at any point by placing switch in OFF or CLOSE
position.
Install a safety barrier across the cargo door opening if work will be done near the open door.
757 General Familiarization (7-2005)
ATA 52 - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Electrical Door Closing
CAUTION:
MAKE SURE THERE IS NO UNWANTED MATERIAL IN THE CARGO DOOR SILL AREA. IF
YOU CLOSE THE CARGO DOOR WHEN THERE IS UNWANTED MATERIAL IN THE SILL
AREA, DAMAGE TO THE CARGO DOOR OR SILL COULD OCCUR.
Make sure there is no unwanted material in the cargo door sill area.
CAUTION:
MAKE SURE THE CARGO NETS ARE SECURE BEFORE YOU CLOSE THE CARGO DOOR. IF
THE CARGO NETS ARE NOT SECURE BEFORE YOU CLOSE THE CARGO DOOR, DAMAGE TO
THE CARGO DOOR OR SILL COULD OCCUR.
Make sure the cargo nets are secure.
Place and hold outside door control switch in the CLOSE position until door is fully closed.
NOTE:
Door motion can be stopped or reversed at any point by placing control switch in OFF or
OPEN position.
Raise and stow external latch handle in the recessed position. Check that the appropriate cargo door message is
extinguished from the Engine Indication and Crew Alerting System (EICAS). Check that pressure vent doors are
open prior to closing handle and are closed after moving handle to closed position. Check that exterior handle
operation is free from binding.
CAUTION:
DO NOT FORCE HANDLE. MAXIMUM NORMAL EXTERNAL HANDLE LOAD IS 70.0 POUNDS
APPLIED 1.0 INCH FROM END OF HANDLE. EXCESS FORCE MAY DAMAGE HANDLE
MECHANISM
757 General Familiarization (7-2005)
ATA 52 - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EQUIPMENT COMPARTMENT DOORS
Overview
The equipment compartment exterior access doors are doors which provide access to the airplane equipment
compartments for maintenance purposes.
Two of these doors lead into pressurized compartments (forward access door, electrical/electronics access door)
and therefore include pressure seals. These doors also have warning sensors which signal the flight crew of an
unsafe condition through the Door Warning Indication System.
All of the equipment doors are normally opened from the outside except the Electrical/Electronics access door
which may also be opened from the inside.
IF 52-11
757 General Familiarization (7-2005)
ATA 52 - Page 44
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EQUIPMENT COMPARTMENT DOORS COMPONENTS
Forward Access Door
This door provides maintenance and inspection access from outside the airplane to the lower fuselage
compartment, located aft of the fwd pressure bulkhead, forward of the nose wheel well and beneath the flight deck
floor.
The forward access door is a hinged, plug-type door that is 18 inches wide by 19 inches long.
The door latch mechanism is manually operated and consists of a flush mounted handle assembly which is
connected by intermediate latch links to the latch pin. A trigger mechanism releases the handle from the flush
stowed position and CCW rotation of the handle will pull the latch pin free of the fuselage latch fitting.
When the door is open, an ACCESS DOOR warning indication will be visible to the flight crew on the P11 Overhead
Panel and on the Engine Indication and Crew Alerting System (EICAS) P2 panel. Closing and latching the door will
extinguish the warning indication.
IF 52-12
757 General Familiarization (7-2005)
ATA 52 - Page 45
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Electrical/Electronics Access Door
This door provides maintenance and inspection access from outside the airplane to the main E/E equipment
center that is located aft of the nose wheel well and forward of the forward cargo compartment.
The E/E access door is a hinged, plug-type door approximately 19-1/2 inches wide by 28 inches long that can be
opened or closed from either inside or outside the airplane.
A pressure seal is attached about the periphery of the door assembly and directly contacts the inside face of the
adjacent body skin panel. A pressure sealing flapper-type drain valve is attached on the inside surface of the door
skin.
The door latch mechanism is manually operated and consists of a handle assembly with a protruding inside handle
and a flush mounted external handle which is connected by intermediate latch links to the latch pin. An over center
spring holds the mechanism in the closed or open position. A trigger mechanism releases the outside handle from
the flush stowed position and CCW rotation of the handle will pull the latch pin free of the fuselage latch fitting.
Using the inside handle requires a CW rotation of the handle.
When door is open an E/E ACCESS DR warning indication will be visible to the flight crew on the P11 Overhead
Panel and the Engine Indication and Crew Alerting System (EICAS) P2 panel. Closing and latching the door will
extinguish the warning indication.
IF 52-13
757 General Familiarization (7-2005)
ATA 52 - Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Air-Conditioning Access Doors
Two air-conditioning access doors provide access to the air conditioning packs from outside (underneath) the
airplane. The air conditioning access doors are located in the lower wing-to-body fairing area, left and right of
airplane centerline, below the wing center section and forward of the main landing gear doors.
Each door is approximately 57 inches wide by 137 inches long and is fabricated from honeycomb composite,
fiberglass and graphite reinforced epoxy. Both doors are manually operated and are hinged along the inboard edge
at four locations. Quick-release latches are provided along the other three edges. Two plunger-type latches are
located along each door forward and aft edge and four pin-type latches are located along the outboard edge of each
door. These doors also have hold-open struts to prevent doors from swinging on hinges when open.
Each large access door has a small, integral, hinged ground air supply access door near the forward end. These
doors are approximately 12 inches wide by 24 inches long and are constructed of composites similar to the larger
door. Two hinges are located along the inboard edges of these doors with quick-release plunger-type latches along
the forward and aft edges (2 per edge).
IF 52-14
757 General Familiarization (7-2005)
ATA 52 - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Service Access Door
The service access door provides maintenance and service access from outside the airplane
ne to the aft fuselage 48
section compartment that houses the stabilizer trim actuator and stabilizer trim jackscrew mechanism.
The door is approximately 20 inches square and is located just aft of the aft pressure bulkhead on the left side of
the lower fuselage.
The door is manually operated, outward opening and is hinged along the forward edge at two places. It has a
spring-loaded latch mechanism that engages latch fittings with two rollers.
A hold-open strut is provided to prevent the door from swinging on its hinges when in the open position.
IF 52-15
757 General Familiarization (7-2005)
ATA 52 - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Elevator Controls Access Door
The elevator controls access door is similar in design to the service access door and provides maintenance and
service access from outside of airplane to the aft fuselage 48 section compartment that houses the elevator and
rudder controls.
The door is approximately 18 inches long by 28 inches wide and is located just forward of the APU access doors at
the airplane centerline on the lower fuselage.
The door is manually operated, outward opening and is hinged along the forward edge at two places. It has a
spring-loaded latch mechanism that engages fuselage latch fittings with two rollers.
A hold-open strut is provided to prevent the door from swinging on its hinges when in the open position.
IF 52-16
757 General Familiarization (7-2005)
ATA 52 - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
APU Access Doors
Access provisions for the APU compartment consist of two clam shell-type doors which open by rotating downward
and outward at hinges along the outboard edge of both doors. The doors are located at the extreme aft end of
lower fuselage tail section and extend (approximately) from STA 1885 to STA 1963.
Five tension-type latches are installed at the inboard interface joint between the two doors. A spring-loaded latch
hook mechanism installed on one door engages a U-type retainer fitting on the other door at five locations to hold
the doors tightly closed. Operation of these latches can only be accomplished with a standard ten-inch screwdriver
blade which is inserted into the latch release mechanism. The latch is closed by placing the doors in the closed
position, engaging the hook latch over the retainer fitting and then pushing the latch handle closed.
Each door has a forward and aft mounted telescoping hold-open strut which will keep the door from swinging on
its hinges when the door is in the open position.
A small spring-loaded pressure vent door is installed on the left APU main access door. The door relieves excess
pressure within the APU compartment.
A drain mast is installed on the right APU main access door. The drain mast provides a drain path for excess APU
combustible fluids when in flight.
IF 52-17
757 General Familiarization (7-2005)
ATA 52 - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Vertical Fin Access Door (Airplanes with SB 53-38)
The vertical fin access door provides access to the vertical stabilizer inspar structure. It is located on the upper
fuselage centerline aft of the pressure bulkhead. The door opens downward and is hinged on the left edge and
bolted shut on the right edge.
IF 52-18
757 General Familiarization (7-2005)
ATA 52 - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Flight Compartment Door
The flight compartment door provides selective entry to the flight compartment.
The door is locked from outside the flight compartment when the door is closed, there is 28-volts DC power on the
airplane, and the FLT DK DOOR switch on the pilots' overhead panel P5 is deactivated.
The door can be opened from inside the flight compartment by turning the door lock.
The door can be opened with a key from outside the flight compartment.
The door can be opened from either side if the FLT DK DOOR switch is activated or if there is not 28-volts DC
power on the airplane.
The flight compartment door electric strike is mounted on the door jamb.
The door electric strike has a solenoid, a fuse pin, and a door strike. The door electric strike operates on 28-volts
DC.
The door opens about hinges mounted on the forward lavatory module. The hinges are attached to the lavatory
with screws. Hinge pins on the door are spring-loaded for easy removal/installation of the door.
The flight compartment door has an upper and a lower section. During normal operation, the two sections open
and close together. The upper section is connected to the lower section by a fuse plate.
The door is fused to allow it to open in the event of rapid decompression. This prevents explosive failure of the
door.
In the event of rapid decompression in the flight compartment, the fuse plate, on the upper door section, and the
fused door stop, on the doorjamb, fail allowing the lower door section to open forward.
In the event of rapid decompression in the passenger compartment, the electric strike fuse pin fails allowing the
door to open aft.
Light shields on the top and bottom edges of the door prevent incidental light from entering the flight
compartment. This is to allow low level lighting on instrument panels.
757 General Familiarization (7-2005)
ATA 52 - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-19
757 General Familiarization (7-2005)
ATA 52 - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 52 - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DOOR WARNING SYSTEM
Overview
The system provides flight crew with visual warning if doors to the pressurized areas are open or unlatched. The
system consists of proximity sensors, door warning annunciator lights, EICAS (Engine Indication and Crew Alerting
System) messages, and logic through the PSEU (Proximity Switch Electronics Unit).
All doors providing access to pressurized areas are included in the system. Each door has at least one proximity
sensor, a discrete EICAS message on the pilots' center instrument panel P2, and a general warning light on the
pilots' overhead panel P5. The EICAS message is displayed and the annunciator light illuminates to indicate that
door is unlatched or open.
757 General Familiarization (7-2005)
ATA 52 - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DOOR WARNING SYSTEM COMPONENTS
Door Warning Sensors
Sensors that are located on the door or the door sill sense door position.
The sensors are two wire, magnetic field-producing coils, contained in non magnetic stainless steel cases. The
sensors are part of the proximity switch system.
When a sensor face is brought near, or moved away from a steel (magnetic) target, the sensor inductance
increases or decreases. This inductance change provides a low or high output signal to the PSEU.
The PSEU processes the sensor signals in three stages.
1. The proximity card supplies a pulse signal to the sensor, and inputs the inductance of the pulse, coming
back from the sensor, to the logic card.
2. The logic card converts the signals, from the proximity cards, into a high or low output. The logic card
output then drives a particular gate in the driver card.
3. The driver card then provides an open circuit or ground for the indication lights, and a logic signal to the
EICAS computer. The EICAS computer then displays the correct EICAS message.
Door Warning Annunciator Lights
There are four amber door warning annunciator lights on the pilots' overhead panel P5 as follows:
1. The ENTRY DOORS light illuminates when any passenger door is not completely closed and locked. When
all passenger doors are closed and locked, the ENTRY DOORS light is not illuminated.
2. The EMER DOORS light illuminates when any emergency door is not completely closed and locked. When
all emergency doors are closed and locked, the EMERGENCY DOORS light is not illuminated.
3. The CARGO DOORS light illuminates when either the No. 1 or No. 2 cargo door is not completely closed
and locked. When both cargo doors are closed and locked, the CARGO DOORS light is not illuminated.
4. The ACCESS DOORS light illuminates when either the forward or E/E access door is not completely closed
and locked. When both access doors are closed and locked, the ACCESS DOORS light is not illuminated.
757 General Familiarization (7-2005)
ATA 52 - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF 52-20
757 General Familiarization (7-2005)
ATA 52 - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EICAS Messages
The EICAS messages are outlined in Table 1. All messages are displayed without a time delay. This means when
the door is opened, the EICAS message is immediately displayed. As soon as all the doors are closed, the
individual or combination message(s) will be removed from the EICAS display.
All individual EICAS messages in table 1 are level C messages.
*[1] If both of these doors are open, individual messages are replaced with a single message: ACCESS DOORS.
*[2] If both of these doors are open, individual messages are replaced with a single message: CARGO DOORS.
*[3] If two or more of these doors are open, individual messages are replaced with a single message: L ENTRY
DOORS.
*[4] If two or more of these doors are open, individual messages are replaced with a single message: R ENTRY
DOORS.
*[5] If both of these doors are open, individual messages are replaced with a single message: EMER DOORS.
757 General Familiarization (7-2005)
ATA 52 - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
NOTES:
757 General Familiarization (7-2005)
ATA 52 - Page 59
For Training Purposes Only
This Page Intentionally Blank
AeroEd LLC – Aviation Education Resource
TABLE OF CONTENTS
B757 GENERAL FAMILIARIZATION
ATA-71-80 PW2000 SPECIFIC
PW2000 SERIES ENGINES ................................................................................................................................... 6
Overview ..................................................................................................................................................... 6
Engine Specifications:................................................................................................................................. 8
ENGINE COWLING ............................................................................................................................................... 9
Overview ..................................................................................................................................................... 9
ENGINE MOUNTS .............................................................................................................................................. 12
Overview ................................................................................................................................................... 12
Forward Lower Engine Mount ................................................................................................................... 13
Aft Lower Engine Mounts ......................................................................................................................... 14
ENGINE VENTS AND DRAINS ............................................................................................................................ 16
Overview ................................................................................................................................................... 16
Drains ....................................................................................................................................................... 18
Vents......................................................................................................................................................... 18
ENGINE CONSTRUCTION................................................................................................................................... 20
Overview ................................................................................................................................................... 20
Designation of Engine Stations ................................................................................................................. 22
Engine Flange Identification ..................................................................................................................... 23
COMPRESSOR SECTION ................................................................................................................................... 25
Overview ................................................................................................................................................... 25
Low Pressure Compressor (LPC) Section ................................................................................................. 25
DIFFUSER AND COMBUSTION SECTION .......................................................................................................... 30
Overview ................................................................................................................................................... 30
TURBINE AND EXHAUST SECTION ................................................................................................................... 34
Overview ................................................................................................................................................... 34
Turbine Nozzle Section ............................................................................................................................. 34
High Pressure Turbine (HPT) ................................................................................................................... 36
Low Pressure Turbine (LPT) .................................................................................................................... 38
Turbine Exhaust Case ............................................................................................................................... 40
ACCESSORY DRIVES ......................................................................................................................................... 42
Overview ................................................................................................................................................... 42
Angle Gearbox ........................................................................................................................................... 42
Main Gearbox ............................................................................................................................................ 42
ENGINE FUEL AND CONTROL ........................................................................................................................... 44
Overview ................................................................................................................................................... 44
FUEL DISTRIBUTION SYSTEM .......................................................................................................................... 46
Overview ................................................................................................................................................... 46
Fuel Pump ................................................................................................................................................ 46
Fuel Distribution Valve ............................................................................................................................. 47
Fuel Injectors ............................................................................................................................................ 48
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 1
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL CONTROL SYSTEM .................................................................................................................................. 49
Overview................................................................................................................................................... 49
Fuel Control (FC) ...................................................................................................................................... 49
Probe for the Inlet Total Pressure/Temperature (PT2/TT2) ................................................................... 50
Electronic Engine Control (EEC) ............................................................................................................... 51
EEC Alternator .......................................................................................................................................... 52
N1 Speed Transducer ............................................................................................................................... 53
Data Entry Plug ........................................................................................................................................ 54
Thermocouple Probe for the Fuel Temperature ........................................................................................ 55
Thermocouple Probe for the Oil Temperature .......................................................................................... 56
Resolver for the Thrust Lever Angle ......................................................................................................... 57
Propulsion Discrete Interface Unit (PDIU) ................................................................................................ 58
Thermocouple Probe for the T 3.5 Temperature ...................................................................................... 59
FUEL FLOW INDICATING SYSTEM .................................................................................................................... 60
Overview................................................................................................................................................... 60
Fuel Flow Transmitter .............................................................................................................................. 60
FUEL FILTER BYPASS WARNING SYSTEM ....................................................................................................... 62
Overview................................................................................................................................................... 62
Differential Pressure Switch for the Fuel Pump Filter .............................................................................. 62
IGNITION SYSTEM ............................................................................................................................................ 63
Overview................................................................................................................................................... 63
IGNITION POWER SUPPLY............................................................................................................................ 64
Overview................................................................................................................................................... 64
Ignition Exciter ......................................................................................................................................... 64
HIGH TENSION DISTRIBUTION SYSTEM ..................................................................................................... 65
Overview................................................................................................................................................... 65
Igniter Plug............................................................................................................................................... 66
ENGINE IGNITION CONTROL ........................................................................................................................ 68
Overview................................................................................................................................................... 68
Engine Start/Ram Air Turbine Control Module ........................................................................................ 68
Fuel Control Module ................................................................................................................................. 68
ENGINE AIR SYSTEM ........................................................................................................................................ 71
Overview................................................................................................................................................... 71
COMPRESSOR AND TURBINE COOLING (CTC) SYSTEM.............................................................................. 72
Overview................................................................................................................................................... 72
Low Pressure Turbine (LPT) Case Cooling Air Valve and HP Compressor (HPC) and HP Turbine (HPT)
Case Cooling Air Valve. ............................................................................................................................ 72
Turbine Vane Cooling Air Tubes ............................................................................................................... 74
Turbine Cooling Air (TCA) System ........................................................................................................... 76
COMPRESSOR AIR CONTROL SYSTEM ............................................................................................................ 79
Overview................................................................................................................................................... 79
COMPRESSOR STATOR VANE CONTROL SYSTEM ...................................................................................... 80
Overview................................................................................................................................................... 80
Stator Vane Actuator ................................................................................................................................ 80
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 2
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Variable Stator Linkage ............................................................................................................................ 80
COMPRESSOR BLEED CONTROL SYSTEM ................................................................................................... 82
Overview ................................................................................................................................................... 82
14th-Stage Bleed Valve ............................................................................................................................. 84
14th-Stage Bleed Converter Valve............................................................................................................. 85
Intercompressor (2.5) Bleed Actuator....................................................................................................... 86
ENGINE CONTROL SYSTEM .............................................................................................................................. 88
Overview ................................................................................................................................................... 88
Thrust Levers ........................................................................................................................................... 90
Autothrottle Servomechanism ................................................................................................................... 91
Thrust Lever Interlock .............................................................................................................................. 92
ENGINE FIRE EMERGENCY SHUTDOWN........................................................................................................ 94
Overview ................................................................................................................................................... 94
ENGINE INDICATING SYSTEM .......................................................................................................................... 97
Overview ................................................................................................................................................... 97
ENGINE PRESSURE RATIO (EPR) INDICATING SYSTEM .............................................................................. 98
Overview ................................................................................................................................................... 98
EPR Indication .......................................................................................................................................... 98
Probe for the Inlet Total Pressure/Temperature (PT2/TT2) ................................................................. 100
Exhaust Pressure (PT4.9) Probes ........................................................................................................... 100
Electronic Engine Control........................................................................................................................ 100
ENGINE TACHOMETER SYSTEM................................................................................................................. 102
Overview ................................................................................................................................................. 102
Engine Speed Card ................................................................................................................................. 104
N1 Speed Transducer ............................................................................................................................. 105
EEC Alternator ........................................................................................................................................ 106
EXHAUST GAS TEMPERATURE (EGT) INDICATING SYSTEM ..................................................................... 107
Overview ................................................................................................................................................. 107
EGT Display ............................................................................................................................................ 107
EGT (T4.9) Thermocouple Probes .......................................................................................................... 108
EGT Thermocouple Cables and Box ........................................................................................................ 109
AIRBORNE VIBRATION MONITORING (AVM) SYSTEM ............................................................................... 110
Overview ................................................................................................................................................. 110
Airborne Vibration Monitoring (AVM) Indication ................................................................................... 110
Engine Accelerometers ........................................................................................................................... 112
AVM Signal Conditioner .......................................................................................................................... 113
ELECTRONIC PROPULSION CONTROL SYSTEM (EPCS) ............................................................................ 115
Overview ................................................................................................................................................. 115
STANDBY ENGINE INDICATION SYSTEM ................................................................................................... 118
Overview ................................................................................................................................................. 118
EXHAUST ........................................................................................................................................................ 119
Overview ................................................................................................................................................. 119
TURBINE EXHAUST SYSTEM ..................................................................................................................... 120
Overview ................................................................................................................................................. 120
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 3
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Turbine Exhaust Sleeve .......................................................................................................................... 120
Turbine Exhaust Plug ............................................................................................................................. 120
THRUST REVERSER SYSTEM ..................................................................................................................... 121
Overview................................................................................................................................................. 121
Thrust Reverser ...................................................................................................................................... 124
V-Groove Latch Bands and Hinges ......................................................................................................... 126
Tension Latches ..................................................................................................................................... 128
Access Doors .......................................................................................................................................... 128
Thrust Reverser Hydraulic Cowl Opening Actuators and Flex Hoses ...................................................... 130
Thrust Reverser Sleeve .......................................................................................................................... 131
Thrust Reverser Blocker Doors and Drag Links ..................................................................................... 132
Thrust Reverser Cascade Segments ....................................................................................................... 133
Thrust Reverser Hydraulic Actuators ...................................................................................................... 134
Thrust Reverser Rotary Flex Shafts and Tubing ..................................................................................... 134
Opening Thrust Reverser ........................................................................................................................ 136
Closing Thrust Reversers ....................................................................................................................... 140
THRUST REVERSER CONTROL SYSTEM .................................................................................................... 142
Overview................................................................................................................................................. 142
Thrust Reverser Directional Control Valve .............................................................................................. 142
Reverse Thrust Lever ............................................................................................................................. 144
Thrust Lever Interlock ............................................................................................................................ 144
Thrust Reverser Linear Variable Differential Transformer ..................................................................... 145
Thrust Reverser Isolation Valve .............................................................................................................. 146
Thrust Reverser Stow Proximity Switch ................................................................................................. 147
THRUST REVERSER POSITION INDICATING SYSTEM ................................................................................ 148
Overview................................................................................................................................................. 148
Thrust Reverser Position Indicating ....................................................................................................... 150
Thrust Reverser Isolation Valve Fault Indicating .................................................................................... 151
OIL SYSTEM ................................................................................................................................................... 152
Overview................................................................................................................................................. 152
OIL STORAGE SYSTEM .............................................................................................................................. 154
Overview................................................................................................................................................. 154
Engine Oil Tank ...................................................................................................................................... 154
Engine Oil Tank Filler Cap ..................................................................................................................... 154
Engine Oil Tank Pressurization Valve ..................................................................................................... 154
ENGINE OIL DISTRIBUTION SYSTEM......................................................................................................... 156
Overview................................................................................................................................................. 156
Lubrication and Scavenge Oil Pump....................................................................................................... 156
Overpressure Relief Valve....................................................................................................................... 157
Main Oil Filter ........................................................................................................................................ 158
Fuel/Oil Heat Exchanger ........................................................................................................................ 160
Fuel/Oil Heat Exchanger Bypass Valve .................................................................................................. 160
Air/Oil Heat Exchanger .......................................................................................................................... 162
No. 4 Bearing Scavenge Valve ................................................................................................................ 163
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 4
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Magnetic Chip Detectors ......................................................................................................................... 164
OIL QUANTITY INDICATING SYSTEM ........................................................................................................ 166
Overview ................................................................................................................................................. 166
Oil Quantity Indication on EICAS ............................................................................................................ 166
Oil Quantity Transmitter ......................................................................................................................... 166
OIL PRESSURE INDICATING SYSTEM ........................................................................................................ 168
Overview ................................................................................................................................................. 168
Main Oil Pressure Transmitter ............................................................................................................... 168
No. 4 Bearing Scavenge Oil Pressure Transmitter ................................................................................. 168
Main Oil Pressure Indication on EICAS .................................................................................................. 168
No. 4 Bearing Scavenge Oil Pressure Indication on EICAS ..................................................................... 170
LOW OIL PRESSURE WARNING SYSTEM .................................................................................................... 172
Overview ................................................................................................................................................. 172
Low Oil Pressure Warning Switch .......................................................................................................... 172
OIL TEMPERATURE INDICATING SYSTEM ................................................................................................. 174
Overview ................................................................................................................................................. 174
Oil Temperature Thermocouple Probe .................................................................................................... 174
Oil Temperature Indication ..................................................................................................................... 174
OIL FILTER BYPASS WARNING SYSTEM.................................................................................................... 176
Overview ................................................................................................................................................. 176
Oil Filter Differential Pressure Switch .................................................................................................... 176
ENGINE STARTING SYSTEM ........................................................................................................................... 179
Overview ................................................................................................................................................. 179
Pneumatic Starter ................................................................................................................................... 179
Starter Control Valve .............................................................................................................................. 179
ENGINE STARTER SERVICING ......................................................................................................................... 182
Add Oil to the Engine Starter ................................................................................................................. 182
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 5
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
PW2000 SERIES ENGINES
Overview
The 757 airplanes are supplied with power by two PW2037 or two PW2040
turbofan engines.
The power plants are held in the strut mounted nacelles. One power plant is installed below each wing.
Each power plant supplies the airplane with hydraulic, pneumatic, and electrical power and also propulsion.
An engine which is fully assembled, and includes the inlet cowl, weighs approximately 8452 pounds.
The PW2037 and the PW2040 engines are a two-spool, axial-flow turbofan engine with a high bypass ratio.
The shaft center of the two rotor assemblies is the same. The low pressure (N1) compressor has five stages
and is driven by a five-stage turbine. The high pressure (N2) compressor has twelve stages and is driven by a
two-stage turbine.
All engine parameters are controlled by the electronic engine control (EEC). The EEC gives a signal to the fuel
control to adjust the fuel flow and servo fuel pressure.
The ignition of the fuel is done by two different systems. The systems will automatically supply and control
ignition as necessary.
Engine cooling air is used to decrease the clearance to a minimum between the rotor blades and the engine
case. This is done to increase fuel efficiency.
This engine includes a modulating 2.5 bleed valve.
Power levels set by the flight crew are sent to the engine from a transducer in the flight compartment.
Many engine parameter transmitters are installed on the engine and send data to the Engine Indicating and
Crew Alerting System (EICAS).
A thrust reverser with a translating sleeve uses fan air to slow down the airplane.
An oil system on the engine is used for lubrication and to decrease the internal engine components.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 6
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
The engine starting system is used to turn the N2 rotor .
An integrated drive generator is installed on the main gearbox and supplies electrical power to many airplane
systems.
Fire detection systems are added to the basic engine to supply the flight crew with an indication of an engine
fire.
A hydraulic pump is installed on the main gearbox and supplies hydraulic power to the airplane.
Bleed air from the engine is used to supply pneumatic power and to operate the air conditioning packs.
IF PW-1
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 7
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Specifications:
ITEM
SPECIFICATION
Axial-Flow, gas turbine turbofan
One
Annular
Two spool, 17-stage compressor, consisting of a 5stage low pressure front compressor (includes 1ststage fan) and a 12-stage high pressure compressor.
7-stage, split, having 1st- and 2nd-stage high
pressure turbine and 3rd-, 4th-, 5th-, 6th-, and 7thstage low pressure turbine.
Type of Engine
Number of Combustion Chambers
Type of Combustion Chamber
Type of Compressor
Type of Turbine
ITEM
Basic Engine
Turbine Exhaust Sleeve
Turbine Exhaust Plug
Inlet Cowl
Demountable Powerplant
757 General Familiarization (7-2005)
WEIGHT
7129 lb (3150 kg)
140 lb. (63 kg)
25 lb. (11.3 kg)
411 lb. (186.4 kg)
8500 lb. (3833 kg)
LENGTH
149 in. (3734 mm)
51 in. (1295 mm)
31 in. (787 mm)
35 in. (889 mm)
235 in. (5969 mm)
ATA 71-80 PW - Page 8
DIAMETER
85 in. (2159 mm)
29 in. (737 mm)
23 in. (584 mm)
97 in. (2464 mm)
106 in. (2692 mm)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE COWLING
Overview
The engine cowling is an aerodynamically smooth surface on the engine and prevents damage to the
components on the engine.
You can get access to the engine components through the fan cowl panels, the core cowl panels and the fan
duct cowl and thrust reverser. Also, you can get access to some components through the access doors in the
cowling.
The fan duct cowl and thrust reverser is part of the nacelle
Inlet Cowl
The inlet cowl is installed on the forward flange of the fan case and is an entry for air to go into the engine
with minimum drag. The inlet cowl is an aluminum structure with a honeycomb core acoustical lining and
Kevlar-graphite external panels.
The inlet cowl weighs approximately 411 pounds.
The thermal anti-icing (TAI) duct supplies warm engine bleed air to a spray duct manifold to prevent ice on
the inlet cowl. The manifold is installed fully around the lip of the inlet cowl.
The total pressure/total temperature (PT2/TT2) probe is installed on the inner face of the inlet cowl. You
can get access through an access door in the outer skin of the inlet cowl.
A pressure relief door is installed on the aft lower section of the inlet cowl.
Fan Cowl Panels
The fan cowl panels are installed on the strut with three hinges on each panel. The fan cowl panels are
latched together near the center at the bottom left with three tension latches.
The fan cowl panels are a composite structure of nomex honeycomb and kevlar-graphite panels with an
aluminum frame. The left fan cowl panel weighs approximately 96 pounds and the right fan cowl panel weighs
approximately 81 pounds.
Two hold-open rods, kept on the inner side of each fan cowl panel, hold the fan cowl panels in the open
position. You can release the lower end of the hold-open rod from the stowed position and attach it to a
receiver bracket. The two hold-open rods must be fully extended to hold each fan cowl panel which is locked
in the open position.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 9
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
A pressure relief door on the left fan cowl panel opens to release pressure if failure of the thermal anti-ice or
start duct occurs.
The right fan cowl panel has an access door in it to manually close the thermal anti-ice (TAI) valve.
A starter vent and IDG service access door is installed near the bottom of the right fan cowl panel.
There is an opening on the bottom edge of the left fan cowl panel to examine the IDG oil level and not open
the cowling.
A port in the lower section of the right fan cowl panel gives access to the starter control valve for manual
override without the cowling open.
Fan Duct Cowl and Thrust Reverser
The right fan duct cowl and thrust reverser has an opening for the engine deoiler exhaust vent.
Core Cowl Panels
The core cowl panels are installed on the strut with three hinges on each panel. The core cowl panels are
latched together at the bottom with three tension latches.
The core cowl panels are made of aluminum. Each core cowl panel weighs approximately 76 pounds.
A hold-open rod, kept on the inner side of each panel, holds the core cowl panels in the open position. You
can release the lower end of the hold-open rod from the stowed position and attach it to a receiver bracket.
One hold-open rod must be fully extended to hold each core cowl panel which is locked in the open position.
There is a pressure relief door on each core cowl panel. These pressure relief doors open to release pressure
if failure of the pneumatic duct occurs.
The left core cowl panel has an exhaust vent for the engine 14th-stage bleed.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 10
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-2
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 11
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE MOUNTS
Overview
The forward and aft lower engine mounts hold the engine and are attached
ched at six locations. The forward
engine mount is attached to the rear flange of the intermediate case at three locations. The aft engine mount
is attached to the double flange on the top of the turbine exhaust case and at points approximately 35 degrees
on the side of the centerline. The lower engine mounts are made to remove the force of the engine thrust and
the vertical and side loads. The lower engine mounts can become larger in the axial and radial directions
when the temperature increases.
IF PW-3
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 12
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Forward Lower Engine Mount
The forward lower engine mount holds the engine at three locations:
1. A shear-pin fitting
2. Two thrust links.
The shear pin is installed into a mating hole in the rear flange of the
he intermediate case and held by four
bolts. The two thrust links are attached to the rear flange of the intermediate case on one side of the shear
pin fitting and to the aft end of the mount assembly. The forward lower engine mount attaches to the strut
with four vertical tension bolts.
IF PW-4
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 13
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Aft Lower Engine Mounts
The aft lower engine mount holds the engine at three points on the double flange of the turbine exhaust case.
The aft lower engine mount includes two tangential links and a center link which attaches to the same mount
fitting. The mount fitting attaches to the strut with four vertical tension bolts.
IF PW-5
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 14
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 15
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE VENTS AND DRAINS
Overview
The engine drain system collects drain fluid and lets the drain fluids fall overboard through a port in the
access door on the lower fan cowl panel.
The engine vents are used to bleed air overboard.
There are nine drains and two vents. The vents are the IDG thermal relief vent and the deoiler overboard
ejector tube vent. The drain outlets are all in one group.
There are six fittings with caps on them in the drain system which are used as fluid traps. You can use these
fluid traps to help you isolate the source of leakage when fluids from the drain mast are found.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 16
For Training Purposes Only
ATA 71-80 PW - Page 17
AeroEd LLC – Aviation Education Resource
IF PW-6
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Drains
(1)
(2)
(3)
(4)
(5)
(6)
(7)
(8)
(9)
Electronic Engine Control (EEC) Alternator Drain
Hydraulic Pump Pad Seal Drain
Fuel Pump Pad Seal Drain
Starter Pad Seal Drain
Oil Tank Scupper and Fuel/Oil Heat Exchanger Bypass Valve Drain
(a) Fuel/Oil Heat Exchanger Bypass Valve section of this drain has a capped fitting which you
can used as a fluid trap to isolate the source of leakage.
Fuel/Oil Heat Exchanger Drain
Integrated Drive Generator (IDG) Pad Seal Drain
Core Accessories Drain
(a) This drain line collects fluids from the components that follow:
1) Intercompressor (2.5) Bleed Actuator
2) HP Compressor and HP Turbine Case Cooling Air Valve
3) LP Turbine Case Cooling Air Valve
4) Stator Vane Actuator
5) Air/Oil Heat Exchanger Valve
6) 14TH-Stage Bleed Converter Valve.
(b) All of the above components have a different fitting with a cap on it, in the drain line but the
stator vane actuator.
(c) When fluid drains from the core accessory drain line, you can use fluid traps to help you
isolate the source of the leakage.
Strut Drain
Vents
(1) The Integrated Drive Generator (IDG) thermal relief vent releases hot oil through the lower access door in
the fan cowl panels.
(2) The deoiler breather vent releases clean air near the drain outlets.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 18
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-7
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 19
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE CONSTRUCTION
Overview
This section includes the general description of the following PW2000 engine sections:
(1)
(2)
(3)
(4)
Compressor Section
Diffuser and Combustion Section
Turbine and Exhaust Section
Accessory Drives
The PW2000 series engines have incorporated technology which provides significant advantages in fuel
consumption and engine weight. The improved engine performance of the basic engine is achieved with a
single-stage fan, five-stage low compressor, 12-stage high pressure compressor, a burner that employs low
emissions features, a two-stage high pressure turbine and a five-stage low pressure turbine. Improved
components include a wide-chord, single-shroud fan blade, single-crystal high pressure turbine blades,
carbon seals in the Nos. 2 and 3 bearing compartments and a full authority electronic engine control system.
Engine General
The PW2000 engine is a two-spool axial-flow turbofan engine of high compression ratio and high bypass ratio
having a 17-stage split compressor, an annular combustion chamber, a 2-stage high pressure turbine and a 5stage low pressure turbine rotor which is mechanically independent of the high pressure system.
The 1st-stage compressor rotor of the front compressor section is much larger in diameter than the other
stages and is called the fan.
The fan provides two separate airstreams. The primary (or inner) airstream traveling through the engine
operates internal devices to generate pressures and gases in the exhaust nozzle and thereby provide a
propulsive force.
The secondary (or outer) airstream is mechanically compressed by the fan when entering the engine and is
ducted to the outside of the engine a short distance from the fan at the fan discharge duct. This secondary
airstream adds to the propulsive force and increases the efficiency of the engine.
The engine cases are bolted together, and support the inner parts of the engine through struts and bearings.
The high pressure system consists of the rear compressor rotor and 1st and 2nd-stage rear compressor drive
turbine rotor. The engine front mount is located at the 12 o'clock position of the intermediate case and carries
thrust, vertical and side loads. The rear mount is located at the 12 o'clock position of the turbine exhaust case
and carries vertical, side and torsional loads. Ground handling provision is provided by pads at six locations
on the fan exit case plus locations on the turbine exhaust case.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 20
For Training Purposes Only
ATA 71-80 PW - Page 21
AeroEd LLC – Aviation Education Resource
IF PW-8
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Designation of Engine Stations
Certain specific points along the engine axial profile are identified by station name to provide ease of
reference for items such as component locations, test taps, and sensor locations.
IF PW-9
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 22
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Engine Flange Identification
Engine flanges are identified in alphabetical sequence from front to rear.
IF PW-10
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 23
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 24
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COMPRESSOR SECTION
Overview
The compressor section consists of a low pressure compressor (LPC) section, intermediate case section, and
a high pressure compressor (HPC) section. The compressor sections each contain a rotor and stator
assembly. The rotor in each of these sections is driven by a separate turbine, and each rotor is mechanically
independent of the other. A rotor contained in the fan section is driven integrally with the front compressor
rotor.
Low Pressure Compressor (LPC) Section
The low pressure compressor (LPC) consists of a five-stage rotor, with four stages of stators. The rotor, is
mated to the LPC shaft by a splined joint, and driven by the low pressure turbine. The first stage of the LPC
is the fan stage, consisting of 36 blades installed in the hub. An aerodynamic inlet cone is mounted on the
face of the hub.
Compressor Inlet Cone
The inlet cone, constructed of Kevlar - epoxy composite, provides a smooth aerodynamic fairing at the inlet
area of the LPC. The inlet cone consists of two segments; the inlet cone segment assembly which is bolted to
the front flange of the LPC hub, and the front cone segment. The front cone segment fastened by bolts to the
front of the segment assembly to complete the streamlined structure.
Low Pressure Drive Turbine Shaft
The LPC drive turbine shaft is single-piece construction, fabricated of heat-resistant steel. The shaft is
splined at the forward end to accept the LPC hub, and at the rear to accept the low pressure turbine rear hub.
Axial support and alignment of the shaft is provided by the No. 1 and No. 2 bearings at the forward end, with
support at the rear provided by the turbine rear hub and No. 5 bearing
Fan Case
The fan case is bolted to the front flange of the fan exit case assembly and provides for containment of the
fan blades. The inlet cowl fastens to the front flange of the fan case assembly. The steel case construction
ensures adequate blade containment, and includes abradable Kevlar-epoxy fan blade rubstrip segments and
sound-suppressing acoustic treatment.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 25
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fan Exit Case
The fan exit case provides primary structural support for the fan cases. The fan exit case is secured to the
outer diameter of the intermediate case and is constructed of a steel inner case section with ten compositematerial aerodynamic vanes which straighten fan discharge air and 76 base and vane assemblies. The outer
aluminum case incorporates acoustic treatment to suppress fan-generated sound. The 2.5 bleed valve and
actuating linkage are also incorporated in the fan exit case.
Externally mounted accessory components consist of the electronic engine control (EEC), EEC Wiring
Harnesses, EEC fuel temperature thermocouple probe, fuel/oil heat exchanger and the fuel/oil heat exchanger
bypass valve. All components are mounted on the left side of the case.
No. 1 and No. 2 Bearings
The No. 1 bearing is a ball bearing and absorbs thrust loads transmitted during engine operation.
The No. 2 bearing is a roller bearing and provides axial load support for the LPC shaft.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 26
For Training Purposes Only
ATA 71-80 PW - Page 27
AeroEd LLC – Aviation Education Resource
IF PW-11
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Intermediate Case
The intermediate case is a major structural element of the engine, providing support for No. 1, 2 and 3
bearings, and absorbing thrust forces generated by both the low pressure and high pressure rotors. Thrust
loads are transmitted to the airframe structure by mount brackets located on the intermediate case.
The intermediate case also includes the radial drive assembly, driveshaft, and bevel gear which transmit
power to the gearbox group. The 5th-stage compressor stator is bolted to the front flange of the intermediate
case.
High Pressure Compressor (HPC) Section
The high pressure compressor (HPC) consists of a 12 stage rotor with 12 stages of stators surrounded by a
two-section titanium and steel case.
The function of the high pressure compressor is to further compress and accelerate air delivered by the low
pressure compressor and to direct this highly compressed air to the diffuser and combustor section.
Driving force for the high pressure compressor is provided by the high pressure turbine through the splined
driveshaft.
Externally mounted accessory components consist of the stator vane actuator, turbine vane cooling air valve,
and two turbine cooling air valves.
High Pressure Compressor Rotor
The high pressure compressor rotor consisting of Stages 6 through 17 is a bolted assembly consisting of a
front hub (Stage 6), a nine-stage drum rotor (Stages 7 through 15), two disks (Stages 16 and 17) and a drive
shaft. The high pressure compressor rotor is supported at the front hub by the No. 3 bearing, and is
supported at the rear (drive shaft) by the No. 4 bearing. The drive shaft is linked directly to the high pressure
turbine by a splined joint.
High Pressure Compressor Case and Stators
The high pressure compressor stators consist of five stages (Inlet, and Stage 6 through Stage 9) of variable
stator vanes and seven stages (Stages 10 through 16) of fixed vanes.
The five variable vane stages are contained in the front section of the high pressure compressor case, which
is axially split into two, 180 degree segments with bolted flanges at the 3 and 9 o'clock locations. Variable
vane outer trunnions fit into bushings in the case segments, and are actuated through unison rings
mechanically linked to the stator vane actuator.
Borescope inspection access is provided at seven ports located in the high pressure compressor case.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 28
For Training Purposes Only
ATA 71-80 PW - Page 29
AeroEd LLC – Aviation Education Resource
IF PW-12
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
DIFFUSER AND COMBUSTION SECTION
Overview
The diffuser and combustion section consists primarily of a nickel alloy diffuser case into which are installed
a single-piece combustion chamber, the compressor exit stator assembly, and the No. 4 bearing, bearing
seals, and bearing housing.
Diffuser Section
The function of the diffuser and combustor is to straighten air flowing from the high pressure compressor,
diffuse the flow to the optimum velocity and pressure for combustion, and to create the high temperature
gases necessary for producing thrust. These gases are created by mixing fuel with the air and supporting
ignition of the mixture.
Diffuser Case
The main structural member of this group is the diffuser case, which extends from the last two stages of the
high pressure compressor rearward to the turbine nozzle group. The diffuser case is constructed with a
gradually increasing cross-section from front to rear, which serves to decrease the velocity of airflow,
converting velocity energy to pressure energy.
The inner portion of the diffuser case, in addition to forming an aerodynamic contour for the flow path,
provides structural support for the No. 4 bearing, and supports the compressor exit stator assembly.
The diffuser case incorporates mounting provisions for 24 externally removable fuel injectors, and two igniter
plugs. Borescope ports are incorporated at four locations around the diffuser case.
Rear Compressor Exit Vane Assembly
The rear compressor exit vane assembly consists of 72 nickel alloy exit vanes installed in inner and outer
shroud rings, and retained by bolts to the front inner flange of the diffuser case. A stepped pair of airseal
lands retained at the same bolt circle are aligned with knife-edge seals on the drive shaft of the high pressure
compressor. Seven nozzle assemblies also secured at this location provide cool air to 17th-stage air sealing
ring.
The rear compressor exit vane assembly straightens the circular flow of air from the high pressure
compressor, converting velocity to pressure.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 30
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-13
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 31
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 32
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Combustion Chamber
The combustion chamber, fabricated into a single piece from welded, rolled segments of nickel alloy,
incorporates magnesium zirconate coating on the combustion side of the liner. The combustion chamber is
retained in position by six doweled end bolts, and encloses the nozzles of the fuel injectors.
The combustion chamber provides the optimum environment for efficient combustion and directs the flow of
hot gases to the turbine nozzle.
Fuel Drain Valve Assembly
The fuel drain valve assembly is located at the 6 o'clock position on the diffuser case. Residual fuel is drained
overboard at engine shutdown as a finger valve opens with decrease of burner pressure.
IF PW-14
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 33
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
TURBINE AND EXHAUST SECTION
Overview
The turbine and exhaust section consists of the turbine nozzle section, high pressure turbine (HPT), low
pressure turbine (LPT), and the turbine exhaust section. Reaction to the combustion gases passing through
the turbines causes the turbines to rotate and drive their respective compressors.
Turbine Nozzle Section
The turbine nozzle section consists of 36 air-cooled vanes bolted around the 1st-stage cooling duct assembly
and retained at the OD by a support ring. The ring of vanes forms a series of nozzles which increase the
velocity of gases exiting the combustion chamber and direct those gases at the optimum angle onto the 1ststage turbine blades.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 34
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-15
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 35
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
High Pressure Turbine (HPT)
The high pressure turbine (HPT) section provides driving force for the high pressure compressor, and
consists of a two-stage bolted rotor incorporating single-crystal nickel alloy blades, and a turbine case and
stator assembly incorporating the second stage nickel alloy vanes. The turbine front case incorporates
segmented outer ducts matched to the first and second stage rotors. Optimum blade tip clearances are
maintained by use of active clearance control to cool the case with compressor air thus reducing heat
expansion.
High Pressure Turbine Rotor
The 1st-stage turbine disk, with 48 blades installed, is secured to the 2nd-stage turbine hub by a bolted
integral spacer. The 2nd-stage turbine hub incorporating 64 blades, mates to the drive shaft of the rear
compressor by a splined joint.
Turbine Case and Vane Assembly
The turbine case incorporates internal mounting flanges to support 40 2nd-stage turbine vanes, and
segmented outer ducts for both the 1st- and 2nd-stage turbine rotors.
Vanes are retained in the case by bolts securing outer buttresses to internal case flanges.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 36
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-16
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 37
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Low Pressure Turbine (LPT)
The low pressure turbine (LPT) provides driving force for the low pressure compressor and consists of a fivestage rotor and stator enclosed in a single piece rear turbine case. The rotor assembly includes five winged
disks bolted to a steel hub with interstage knife-edge seals captured between the bolted flanges of the disks.
Blades are installed and are retained in the disks by the interstage seals. The five stages of stators are
assembled as clusters of three vanes. These clusters are installed in internal case flanges and are retained by
the OD shroud segments installed in the case, between stages. Optimum blade tip clearances are maintained
by use of active clearance control to cool the case with fan discharge air thus reducing heat expansion.
Provision for borescope inspection is incorporated at the 9 o'clock location on the case.
Low Pressure Rotor
The five low pressure turbine disks are joined together axially at adjacent flanged wing sections and secured
by bolts. The sixth stage disk is bolted to the OD flange of the hub which is mated to the drive shaft of the
low pressure compressor by a splined joint.
Low pressure turbine blades are installed in axial, dovetail slots and are retained in the disks by contact with
rear edges of the interstage seals (Stages 4 through 7), or by contact with the rear edge of the third stage
airseal support (Stage 3). All LPT blade airfoils are hollow in order to reduce weight. Blade platforms are
extended axially to provide effective gaspath sealing.
Low Pressure Turbine Case and Vanes
The low pressure turbine case is a single-piece, machined weldment forging. Internal slotted flanges provide
hooked engagement with stator vane feet and with outer shroud segments positioned behind each of the five
vane stages.
Low pressure turbine vanes are are assembled as clusters of three vanes bonded at adjoining ID and OD
buttresses. The Stages 3 through 7stator incorporates hollow airfoils, with an aerodynamic transition duct
directing gas flow exiting the high pressure turbine.
Active clearance control tubes surround the low pressure turbine case, and supply fan discharge air to cool
the surface of the case during cruise power operation. Cooling the case surface results in shrinkage of the
case with decreased tip clearances between blade tips and airseal lands resulting in improved turbine
efficiency.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 38
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-17
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 39
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Turbine Exhaust Case
The turbine exhaust case is fabricated as a concentric inner cone segment and outer duct ring joined by
fifteen welded radial struts. Two circumferential rails on the OD of the case provide mounting points for the
rear of the engine.
The inner cone portion of the case provides support for the No. 5 bearing, with load forces transmitted from
the bearing, through the struts to the outer case structure. No. 5 bearing oil pressure and scavenge tubes are
routed within case struts at the 8 and 4 o'clock locations respectively.
Pressure sensing provisions are incorporated integrally within the struts of the case, while sleeves for
installation of exhaust gas temperature thermocouple probes are provided within the leading edge of six of the
case struts.
Number 5 Bearing
The No. 5 roller bearing provides radial positioning for the rear of the low pressure compressor shaft and is
retained in an integral support within the exhaust case.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 40
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-18
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 41
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ACCESSORY DRIVES
Overview
The accessory gearboxes consist of an angle gearbox, a main gearbox and their externally mounted
components. The gearboxes provide mechanical connection between the high pressure rotor (N2) and
accessory components mounted on the main gearbox. The accessory components mounted on the main
gearbox are the starter, integrated drive generator, electronic engine control (EEC) alternator, engine driven
hydraulic pump, and the fuel pump. In addition, the main gearbox incorporates an external de-aerator, a
pressure relief valve, oil filters and various oil pumps. Seals are installed in each of the accessory component
drive pads.
Angle Gearbox
The angle gearbox is located beneath the fan exit case and attaches to the main gearbox through a short
horizontal driveshaft. Together, the angle gearbox and the main gearbox are mounted by means of four
flexible support links attached to lugs on the gearbox casings and to brackets bolted to the fan case.
The angle gearbox transmits power from the engine to the main gearbox and from the starter on the main
gearbox to the engine. The angle gearbox is a cast aluminum housing, with transmission of power to and from
the engine accomplished through a tower-shaft geared to the high pressure compressor (HPC) rotor. The
bevel gear set drives a horizontal input gear shaft running forward to the main gearbox. Two metered jets
supply pressure oil for lubrication of bearings and gears within the angle gearbox.
Main Gearbox
The main gearbox is supported along with the angle gearbox by four support links. These support links are
attached to lugs on the gearbox casings and to brackets bolted to the fan case.
The main gearbox transmits power from the engine to the accessories mounted on the gearbox and from the
starter on the gearbox to the engine. The power is obtained from the high pressure rotor, through a driveshaft
in the intermediate case, an angle gearbox and the main gearbox driveshaft. The main gearbox provides
optimum speeds, torques, and monitoring provisions required for the accessories to perform their various
functions. The main gearbox consists of a cast aluminum housing incorporating a gear train and seven drive
pads. Each of the seven accessory drive geartrain sections including the gearbox oil scavenge pump, is an
individually replaceable plug-in unit which fits into the face of the gearbox housing sealing the gearbox core.
Drive pads on the forward face of the gearbox provide for installation of the EEC alternator, hydraulic pump,
and the starter. The lubrication and scavenge oil pump is installed on the forward-facing mount pad of the
gearbox oil scavenge pump. Provision for remote cranking of the high pressure rotor is also located on the
front face. Drive pads on the rear face of the gearbox provide for installation of the integrated drive generator
(IDG) and the fuel pump. The input gearshaft from the angle gearbox joins the main gearbox through a boss
on the rear face.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 42
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-19
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 43
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FUEL AND CONTROL
Overview
The system for the Electronic Engine Control (EEC) provides fuel control and distribution schedules for engine
operation. The EEC system supports engine shutdown operations, maintains stable and efficient engine
operations at idle, part power, full power, and during thrust transient conditions. The EEC system sets and
maintains the engine thrust setting as a function of flight crew commands and engine ratings. The system
also provides most of the data for engine thrust setting and condition indications which are displayed in the
flight deck. System components include:
Electronic Engine Control (EEC) - The EEC is a dual channel digital computer mounted on the engine fan
case.
Fuel Control Unit (FCU) & Fuel Pump - The FCU is mounted to the fuel pump on the engine gearbox and
accepts EEC commands to govern fuel flow and to operate other engine control devices.
The fuel pump is mounted to the engine gearbox and provides pressurized fuel to the fuel control unit.
Permanent Magnet Alternator (PMA) - The PMA is mounted on the gearbox and is the primary power
source of the EEC. The PMA also senses N2 rotational speed.
Actuators and Valves & Sensors - The following actuators and valves, controlled by the EEC, provide
stabilization and increased engine efficiency. They include stator vane actuator, 2.5 bleed valve, 14th-stage
bleed valve and converter valve, the cooling air valve of the high and low turbine case, air/oil heat exchanger,
the bypass valve of the fuel/oil heat exchanger, directional control valve and interlock actuator of the thrust
reverser, and associated control relay logic.
Various sensors are used to adjust thrust commands. These sensors include the resolver for the thrust lever
angle, total temperature probe for the inlet air, N1 speed transducer, sensors for the engine burner pressure,
static pressure sensor, probe for the exhaust gas pressure, fuel temperature probe, and probes for the
exhaust gas temperature.
Fuel Distribution Valve and Fuel Injectors - The fuel distribution system routes fuel to the combustion
chamber. Components used in distributing fuel are the fuel distribution valve, located on the bottom of the
engine under the thrust reverser, fuel manifolds and the fuel injectors, which surround the turbine section.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 44
For Training Purposes Only
ATA 71-80 PW - Page 45
AeroEd LLC – Aviation Education Resource
IF PW-20
757 General Familiarization (7-2005)
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL DISTRIBUTION SYSTEM
Overview
Fuel distribution is provided by the fuel pump, fuel control, fuel distribution valve, and fuel injectors.
The fuel distribution system supplies fuel to the fuel injectors in the engine combustion chamber. System
components include fuel pump, fuel control, fuel distribution valve, and fuel injectors.
Fuel Pump
The fuel pump is mounted on the engine gearbox, located on the bottom of the fan case.
The fuel pump comprises a centrifugal boost stage, a high pressure stage with a single gear, and a filter. The
filter is bypassed when differential pressure exceeds 9 psid.
A pressure switch is actuated when filter bypass is pending, displaying a fuel filter message on EICAS.
IF PW-21
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 46
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fuel Distribution Valve
The fuel distribution valve is located at the six o'clock position on the diffuser case.
The fuel distribution valve receives metered fuel flow from the fuel control and apportions it to the fuel
manifolds and injectors.
Fuel filtering is provided at the valve input. Trimming plugs are provided in the discharge ports to balance the
manifold fuel flows at high engine power.
IF PW-22
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 47
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Fuel Injectors
Twenty-four fuel injectors are mounted on the diffuser case. Twelve fuel manifolds provide fuel to the fuel
injectors. Each manifold supplies fuel to two fuel injectors.
IF PW-23
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 48
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL CONTROL SYSTEM
Overview
Engine fuel control is provided by the electronic engine control.
System components include EEC, fuel control, EEC alternator, N1 speed transducer, and inlet probe for the
total pressure/temperature (PT2/TT2). Other components are the thermocouple probe for the fuel
temperature, the thermocouple probe for the oil temperature, the resolver for the thrust lever angle and the
propulsion discrete interface units (PDIU).
Fuel Control (FC)
The fuel control is mounted on the engine fuel pump, located on the gearbox.
The FC meters fuel flow in response to commands from the EEC. The FC also controls the positions of the
stator vane actuator, valve for the air/oil heat exchanger, and intercompressor bleed valve through fuel
pressure.
An automatic overspeed system is built-in to the FC, driven by the speed of the input shaft of the fuel pump.
The speed of the high pressure compressor is governed at a flight-safe speed whenever the trip speed is
exceeded.
IF PW-24
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 49
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Probe for the Inlet Total Pressure/Temperature (PT2/TT2)
The PT2/TT2 probe is located near the 11 o'clock position in the inlet cowl. The probe measures the total
pressure and total temperature of the air mass entering the engine.
Total temperature is measured by two platinum resistance elements. Each channel monitors one element and
converts the resistance to a temperature signal.
Total pressure is routed to one of the pressure transducers in the main channel of the EEC.
The probe has an internal heater to protect the probe from icing conditions.
IF PW-25
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 50
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Electronic Engine Control (EEC)
The EEC is mounted on the engine fan case with vibration-isolation
ation-isolation hardware. The EEC is a dual channel unit
that automatically switches to the opposite channel when the main channel is faulty.
The EEC is linked to the airplane Air Data Computer (ADC) and Thrust Management Computer (TMC) by
digital data links. The ADC provides the EEC with total air pressure and temperature readings and pressure
altitude data. The TMC receives engine performance and rating data from the EEC.
The EEC also receives rotor speed, pressure, and temperature signals from the engine. These parameters are
combined with signals from the ADC and flight compartment thrust levers to command outputs to engine
actuators.
EEC outputs are sent to the fuel control, TMC, cooling valves of the high and low pressure compressors, and
the solenoid valve for the turbine cooling air.
Power to the EEC is provided by the EEC alternator. The EEC senses N2 rotor speed from the EEC alternator.
A data entry plug is connected to the EEC to provide thrust rating and data of the thrust-versus-engine
pressure ratio to the EEC. The plug remains connected to the engine when removing the EEC.
IF PW-26
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 51
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
EEC Alternator
The EEC alternator is mounted on the gearbox. The alternator is composed of a rotor
tor and a stator. The rotor
is mounted on a stub shaft, and the stator fits over the rotor and mounts to the gearbox.
The stator consists of two 3-phase windings and a single phase winding. Each 3-phase winding provides
power and an N2 rotor speed signal to a channel of the EEC. The single phase winding provides an N2 rotor
speed signal to EICAS.
The alternator operates between 3500-37850 rpm. A shaft seal seals the alternator from the gearbox. The
alternator will still operate when filled with oil, in case of shaft seal failure.
IF PW-27
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 52
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
N1 Speed Transducer
The N1 speed transducer is mounted on a dipstick arrangement that inserts through the intermediate case
wall. The transducer pickup unit is mounted in close proximity to 60 teeth machined on the shaft of the low
pressure compressor. The transducer provides two identical, independent outputs - one for each channel of
the EEC.
IF PW-28
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 53
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Data Entry Plug
The data entry plug for the EEC consists of a 19-contact
ontact connector plug and a connector backshell. A lanyard
provides a means of permanently attaching the data entry plug to the engine. The plug identification is
inscribed on the connector backshell. Pins (insulating) are installed in all unused contacts not occupied by
electrical lead assemblies (jumper wires).
IF PW-29
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 54
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Thermocouple Probe for the Fuel Temperature
A thermocouple probe for the fuel temperature is installed in the fuel-out end of the fuel/oil heat exchanger.
The probe consists of two alumel-chromel thermocouples, which provide independent signals to each channel
of the EEC.
IF PW-30
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 55
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Thermocouple Probe for the Oil Temperature
A thermocouple probe for the oil temperature is installed at the oil inlet of the fuel/oil heat exchanger. Two
alumel-chromel thermocouples provide independent oil temperature signals to each channel of the EEC.
IF PW-31
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 56
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Resolver for the Thrust Lever Angle
Resolvers for the thrust lever angle are dual precisionn resolvers which supply electrically isolated thrust
command signals to both channels of the electronic engine control (EEC).
The resolvers for the thrust lever angle are located on the autothrottle assembly.
IF PW-32
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 57
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Propulsion Discrete Interface Unit (PDIU)
There are two PDIUs, one for each engine. The PDIU is a digital processor which collects, buffers, and
converts analog discrete signal data from the airframe to an ARINC digital bus format for use by the EEC, and
which processes ARINC digital data from the EEC for engine starter cutout and for N2 control mode indication.
The PDIUs are located in the forward E/E equipment bay on the E5 shelf.
IF PW-33
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 58
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Thermocouple Probe for the T 3.5 Temperature
(AIRPLANES WITH BG-11 ENGINES AND SUBSEQUENT)
The T 3.5 thermocouple probe is located at an AP-9
-9 borescope port, between Flanges K and M, at
approximately the two o'clock position, when you look at it from the front of the engine. The probe measures
the high compressor discharge total temperature.
The probe consists of two isolated alumel-chromel thermocouples which provide independent signals to each
channel of the EEC.
The temperature value is used by the EEC to cause a lower engine fuel flow rate during the startup of a hot
engine on the ground. The lower fuel flow rate makes the HPC compressor more stable and reduces the
chance of an engine stall or a hung start.
IF PW-34
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 59
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL FLOW INDICATING SYSTEM
Overview
The rate of fuel flow is measured by a fuel flow transmitter and displayed on the lower EICAS display.
The amount of fuel used is calculated by the flight management computer (FMC).
The main system component is the fuel flow transmitter.
Fuel Flow Transmitter
The fuel flow transmitter is mounted in the fuel line between the fuel control and the fuel distribution valve.
The fuel flow transmitter consists of two rotors and a swirl generator. One rotor is free-wheeling and one is
restrained by spring force. The free-wheeling rotor contains two magnets - one to provide a start pulse and
one to provide a stop pulse. The restrained rotor contains the magnetic pickup for the stop pulses.
Fuel entering the transmitter is directed to the swirl generator, which imparts angular velocity to the fuel. The
angular velocity of the fuel causes the rotors to turn.
The restrained rotor turns until spring force equals the force of the fuel turning the rotor. Thus, the position
of the stop pickup changes with fuel flow.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 60
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-35
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 61
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
FUEL FILTER BYPASS WARNING SYSTEM
Overview
The bypass warning system for the fuel filter indicates an impending bypass of the fuel pump filter. The main
system component is the differential pressure switch for the fuel pump filter.
Differential Pressure Switch for the Fuel Pump Filter
The differential pressure switch for the fuel pump filter is mounted on a boss on the left side of the fuel
pump. The switch is actuated by a differential pressure of 5.5 ±.5 psid, and deactuates when
pressure falls below 3.5 ±.5 psid.
IF PW-36
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 62
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IGNITION SYSTEM
Overview
The ignition system supplies electrical energy at high voltage to start or sustain combustion of the fuel-air
mixture in the engine. Each engine ignition system uses two capacitive-discharge exciter units, each rated at 4
joules. The two circuits are electrically and physically independent. You can use either ignition system to start
the engine or for continuous ignition when selected for flight in bad weather conditions.
The ignition system consists of the Ignition Power Supply, High Tension Distribution, and Engine Ignition
Control.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 63
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IGNITION POWER SUPPLY
Overview
The ignition power supply includes two independent ignition exciters installed on the compressor case at the
three o'clock position. Each ignition exciter is of the capacitive-discharge type, converting
erting 115-volt ac to 24kilovolts dc (nominal). The airplane main ac buses supplies primary power to the ignition exciters; the left
bus powers left engine ignition exciters, the right bus supplies power to the right engine ignition exciters.
Alternate ac power is automatically supplied by the 115-volt ac standby bus when primary power is
unavailable.
Ignition Exciter
The ignition exciters are mounted on a bracket located at the three o'clock position on the compressor case.
The ignition exciter consists of a power transformer, a voltage doubler section, a storage capacitor, a
discharger tube, and a high tension transformer. A filter network on the ignition exciter input keeps high
frequency noise from getting onto the main AC buses. A bleeder resistor is provided to decrease the energy in
the ignition exciter if the igniter plug fails to fire. Each ignition exciter is capable of 4 joules continuous-duty
output.
IF PW-37
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 64
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
HIGH TENSION DISTRIBUTION SYSTEM
Overview
The high tension distribution system conducts the high voltage dc from the ignition exciters to the igniter
plugs. The system consists of exciter-to-igniter plug cables and igniter plugs.
Exciter-to-Igniter Plug Cable
The exciter-to-igniter plug cable includes an insulated high tension lead in an air-cooled flexible steel conduit.
The steel conduit is fed with fan discharge air and provides shielding for the high tension lead. The exciterto-igniter plug cables are installed along the right side of the engine, and connect the ignition exciters to the
igniter plugs.
IF PW-38
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 65
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Igniter Plug
The igniter plugs are installed at the 4 and 5 o'clock positions of the diffuser case, aft of the fuel injectors.
Each igniter plug is fed by one ignition exciter.
IF PW-39
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 66
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 67
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE IGNITION CONTROL
Overview
Ignition control for normal operation is provided by switches on the engine start/ram air turbine control
module and the FUEL CONTROL switches. Positioning the switches properly will supply electrical power to the
ignition exciters. The ignition exciters convert the input power into a high energy pulsating spark across the
igniter plug gap to initiate or sustain combustion in the engine. Power for ignition control system one is 115
volts ac supplied from the overhead circuit breaker panel P11, left ac bus for the left engine ignition system,
and right ac bus for the right engine ignition system. Power for ignition system two is 115 volts ac supplied
from overhead circuit breaker panel P11, stby ac bus for both left and right engine ignition systems.
If primary ac power is not available, standby power is automatically used. Standby power is 115 volts ac
supplied from the overhead circuit breaker panel P11. L IGN STBY BUS and R IGN STBY BUS messages show
on the EICAS maintenance mode when ignition control is powered from the standby bus.
Engine Start/Ram Air Turbine Control Module
The engine start/ram air turbine control module is comprised of two ENG START switches, an ignition
selection switch and two starter control VALVE disagreement lights. There is one ENG START switch for each
engine. Each ENG START switch has five positions: OFF, AUTO for automatic ignition if certain conditions
exist, GND for ground start, CONT for continuous ignition, and FLT for flight start. The ignition selection
switch has three positions: 1, 2, and BOTH for ignition system 1, ignition system 2, and both ignition systems,
respectively.
Fuel Control Module
The FUEL CONTROL switch has two positions: CUTOFF and RUN. To energize either ignition system, the FUEL
CONTROL switch must be in the RUN position.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 68
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-40
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 69
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 70
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE AIR SYSTEM
Overview
Many engine systems are made to control airflow through the engine.
Air is bled from the low pressure compressor to decrease the temperature of some engine external
accessories and engine case and turbine vanes.
The angle of the stator vanes in the high pressure compressor is controlled to provide maximum compressor
stability.
Air is bled from the compressors to provide adequate stall margin at low power settings, during engine
transients and in reverse thrust.
Various airplane components require pneumatic power. Air can be bled from the 10th and 14th stages of the
high pressure compressor to provide pneumatic power.
The air/oil heat exchanger utilizes fan discharge air which is controlled by a modulating butterfly valve
commanded by the EEC.
Stator vane anti-ice is not required, so there is no system to provide stator vane anti-ice.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 71
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COMPRESSOR AND TURBINE COOLING (CTC) SYSTEM
Overview
The compressor and turbine cooling (CTC) system controls blade tip clearances in the two turbines and the
high pressure compressor for higher efficiency.
The CTC system has two important functions - engine case cooling and turbine vane cooling.
The CTC system includes the LP turbine case cooling air valve, HP compressor and HP turbine case cooling
air valve, case cooling manifolds, and turbine vane cooling air tubes.
The Turbine Cooling Air (TCA) system provides necessary cooling air for internal turbine parts.
Low Pressure Turbine (LPT) Case Cooling Air Valve and HP Compressor (HPC) and HP Turbine (HPT)
Case Cooling Air Valve.
These valves operate the same, but the valves supply to different sections of the engine case.
The two valves are attached to case cooling air ducts to the rear of the intermediate case at approximately the
ten o'clock position.
The valves are fully modulating butterfly valves which are hydraulically operated with engine fuel.
The electronic engine control (EEC) supplies input to a torque motor in the valve which lets fuel pressure to
position the valve.
Each valve has a dual variable differential transducer (LVDT) which provides position feedback to the EEC.
If there is no power to the torque motor in the valve, the HPC/HPT valve will automatically close, while the
LPT valve will automatically open.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 72
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-41
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 73
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Turbine Vane Cooling Air Tubes
Four ducts supply 14th-stage air to the turbine vanes for continuous cooling.
The two turbines have cooling manifolds which go around the full circle of the turbine cases.
IF PW-42
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 74
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 75
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Turbine Cooling Air (TCA) System
The turbine cooling air system consists of three manifolds which port 14th-stage bleed air to the high
pressure turbine second vane outer cavity.
Two air valves are controlled by a single actuator featuring
uring hydraulic (fuel) actuation, operating on command
from the EEC.
This actuator is normally in the extended (valves open) position.
When the actuator retracts, on command from the EEC, the air valves close reducing the airflow through the
TCA system.
Turbine temperature information is provided to the EEC by means of a temperature probe in the high pressure
turbine inter-stage inner air seal cavity.
The inter-stage inner seal cavity probe consists of two redundant thermocouples which are used to establish
the cooling level necessary to maintain optimum compartment temperature.
The EEC transmits an electrical signal commanding the position of the actuator for the turbine cooling air
system valves based on the inter-stage seal cavity probe temperature.
IF PW-43
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 76
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-43A
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 77
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 78
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COMPRESSOR AIR CONTROL SYSTEM
Overview
The compressor control system has two systems:
(1) Stator Vane Control System
(2) Bleed Control System
The compressor stator vane control system allows scheduling of the HPC stator vanes. This increases the stall
margin and engine performance.
The compressor bleed control system bleeds air from the compressor section to decrease the pressure
difference across the compressor. This will make the engine start easier, and will increase the stall margin at
low power settings and in reverse thrust.
The stator vane control and engine bleed control systems increase the compressor performance, decrease fuel
consumption and keep the stall margin.
Compressor Stator Vane Control System
This system includes the electronic engine control, the fuel control, the stator vane actuator and the bellcrank,
linkage and unison rings.
Compressor Bleed Control System
A modulating bleed valve and a two-position bleed valve supplies compressor bleed control.
The 14th-stage bleed valve bleeds air through a bleed duct and exhausts it out the left core cowl. This air is
vented during low power settings to decrease the possibility of engine surge. The 14th-stage bleed valve
operation is controlled by the EEC. The 14th-stage bleed valve is operated by engine bleed air from the
converter valve for the 14th-stage bleed. The EEC gives a signal to the converter valve which in sequence
operates a torque motor to generate a pressurized fuel signal. This signal produces a pressurized air
command for input to the 14th-stage bleed valve.
A modulating bleed valve vents intercompressor air during moderate power setting. The modulating bleed
system includes the intercompressor (2.5) bleed valve actuator, the 2.5 bleed valve ring and the fuel control.
The 2.5 bleed operation is controlled by the EEC.
All bleed valves are closed during high power settings.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 79
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COMPRESSOR STATOR VANE CONTROL SYSTEM
Overview
The angle of stator vanes improve the stall margin and engine performance.
Compressor variable stator vane scheduling is done by the EEC. The EEC adjusts the stator vane angle as a
function of N2 rotor speed and current vane angle and supplies a signal to set the stator vane angle when
compressor stall (surge) occurs.
An electrical signal from the EEC is sent to a torque motor in the fuel control. The torque motor changes the
hydraulic pressures sent to the stator vane actuator.
As the stator vane actuator gets differential hydraulic pressures the actuator piston extends or retracts.
The EICAS EPCS page shows the rod end travel of the stator vane actuator.
The end of the stator vane actuator attached to the bellcrank moves and turns the bellcrank.
Unison rings, attached to the bellcrank by the unison ring links, change the variable stator vane angle.
Stator Vane Actuator
The stator vane actuator is installed on the right side of the high pressure compressor case.
The stator vane actuator is a linear actuator operated hydraulically with fuel pressure from the fuel control as
the hydraulic medium.
The stator vane actuator has a dual linear variable differential transformer (LVDT) in it to supply the EEC with
position feedback of the actuator piston.
Variable Stator Linkage
The stator vane actuator is attached to a bellcrank which has many attach points for linkages which connect
the bellcrank to the unison rings.
The linkages from the bellcrank are attached to unison rings which turn to adjust the stator vane angle.
Variable Stator Vanes
The first five stages of the high pressure compressor have variable stator vane angles.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 80
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-44
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 81
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
COMPRESSOR BLEED CONTROL SYSTEM
Overview
14th Stage Bleed System
Air is bled from the high pressure compressor during an engine start to increase the compressor stability as
the engine N2 rpm increases.
The converter valve for the 14th-stage bleed is a servo-piloted, hydraulically (engine fuel) operated, threeway, two-position pneumatic valve. The converter valve is installed on the rear of the intermediate case at
approximately the 4 o'clock position. The converter valve is used to pressurize and bleed the servo port of the
14th-stage bleed valve with high compressor air (17th-stage).
The EEC controls the position of the converter valve for the 14th-stage bleed.
When the energy in the converter valve is removed, compressor discharge air is bled to ambient air and the
14th-stage bleed valve stays open.
When the 14th-stage bleed valve is open, 14th-stage air is bled through the 14th-stage bleed valve, along a
duct and discarded overboard through an opening in the left core cowl panel.
When the EEC energizes the converter, compressor discharge air is routed to the 14-stage bleed valve.
The pressure at the inlet to the 14th-stage bleed valve closes the valve.
EICAS/EPCS STATUS messages report bleed system malfunctions.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 82
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Modulating Bleed Air System
The electronic engine control (EEC) schedules the modulating bleed valve position at some power settings and
during reverse thrust.
The scheduling of the 2.5 bleed valve is accomplished by the electronic engine control (EEC).
The signal from the EEC goes to the torque motor in the 2.5 bleed valve actuator. The torque motor adjusts
the position of a flapper nozzle valve.
The flapper nozzle valve adjusts the hydraulic pressure received from the fuel control to position the 2.5 bleed
valve ring.
The 2.5 bleed valve actuator is made to move the 2.5 bleed valve ring fully open if there is no power to the
torque motor.
If the 2.5 bleed valve actuator fails open, the Engine Indicating and Crew Alerting System (EICAS) makes a
record a status message L (R) ENG BLEED CONT in the nonvolatile memory.
EICAS/EPCS STATUS messages report bleed system malfunctions.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 83
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
14th-Stage Bleed Valve
The 14th-stage bleed valve is installed on the left side of the high pressure compressor case.
The 14th-stage bleed is closed pneumatically when compressor discharge air is supplied to the 14th-stage
bleed valve.
A vent at the end of the 14th-stage bleed valve lets 14th-stage air flow from the compressor, through a duct
and flow overboard through the left core cowl.
The 14th-stage bleed valve is spring-loaded open.
IF PW-45
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 84
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
14th-Stage Bleed Converter Valve
The converter valve for the 14-stage bleed
eed is installed on the rear of the intermediate case at approximately
the 6 o'clock position.
The solenoid is a three port, electromagnetically controlled valve.
An integral torque motor and pilot valve use inlet pressure from the fuel control to position the pneumatic
section of the converter valve. The torque motor contains dual coils which are independently controlled by
each channel of the EEC. If the power or fuel supply pressure decreases, the converter valve will open the
pneumatic outlet port and cause the 14th-stage bleed valve to open.
IF PW-46
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 85
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Intercompressor (2.5) Bleed Actuator
The 2.5 bleed actuator is installed at approximately the 8 o'clock position on the intermediate case bulkhead.
The 2.5 bleed actuator is hydraulically operated by engine fuel pressure from the fuel control and is used to
position the intercompressor bleed ring in a position from open to closed.
Control of the 2.5 bleed actuator is done through a pilot valve which is controlled by an electrical signal from
the EEC to a torque motor. The torque motor and pilot valve are integral with the 2.5 bleed actuator. Actuator
position feedback is by a dual linear variable differential transformer (LVDT) installed on the side of the 2.5
bleed actuator.
IF PW-47
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 86
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 87
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE CONTROL SYSTEM
Overview
Thrust commands from the flight crew or the thrust management computer are transmitted to the engine by
the engine control system. The engine control system consists of the thrust lever assembly, autothrottle
assembly, and thrust lever interlock.
Thrust Levers
Selection of a level of forward thrust, from idle thrust to maximum rated thrust of the engine, is made by
moving the forward thrust lever forward until the EPR COMMAND indication on EICAS is at the desired level.
The engine will accelerate to this thrust level, as indicated by the ACTUAL EPR indication.
Selection of reverse thrust is made by placing the forward thrust lever to idle and then rotating the reverse
thrust lever up and aft. This causes the thrust reverser to deploy and the engine to deliver up to full rated
reverse thrust based on the reverse lever position.
The forward thrust lever is locked at the idle position when the reverse thrust lever is in use. The reverse
thrust lever is locked when the forward thrust lever is moved forward out of idle.
Autothrottle Servomechanism
The autothrottle servomechanism is a single unit to position both thrust levers as commanded by the thrust
management computer. Motion of the thrust levers is transmitted by a rod to the thrust control unit which
incorporates a clutch, which provides both a friction feel force for the thrust lever and the means by which the
autothrottle servomechanism is coupled to the thrust levers.
Thrust Lever Interlock
The thrust lever interlock operates based on commands from EEC to prevent movement of the thrust levers,
in the forward or the reverse direction, until the thrust reverser has translated to the correct commanded
position.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 88
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-48
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 89
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Thrust Levers
Thrust levers are found together between the captain and first officer so that they are easily operated by
either the captain or first officer.
The flight crew selects increased or decreased and forward or reverse thrust with the thrust levers.
IF PW-49
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 90
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Autothrottle Servomechanism
The autothrottle servomechanism is a single unit to position
tion both thrust levers as commanded by the thrust
management computer. A clutch provides a friction feel force for the thrust lever and coupling to the
autothrottle servomechanism.
IF PW-50
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 91
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Thrust Lever Interlock
The thrust lever interlock prevents thrust lever movement during thrust reverser translation. The thrust lever
interlock is located on the autothrottle assembly.
IF PW-51
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 92
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 93
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE FIRE EMERGENCY SHUTDOWN
Overview
This section describes the procedure for an engine fire emergency shutdown.
An engine fire is indicated on the Engine Indication and Crew Alerting System (EICAS) display, illumination of
the master warning lights, illumination of the FIRE light, a continuous ringing of the fire warning bell,
illumination of the fire switch, and illumination of the fuel control switch.
Engine fire emergency shutdown is accomplished by moving the thrust lever to idle position, the fuel control
switch to CUTOFF position, and pulling the fire switch. Moving the thrust lever to idle position decelerates the
engine. Moving the fuel control switch to CUTOFF, closes the engine fuel valve, stops engine ignition, and
deactivates the thrust reverser control system. Pulling the fire switch arms the engine fire extinguisher,
silences the fire bell, depressurize the hydraulic system, and isolates the engine. Engine isolation deenergizes the generator control relay, closes the hydraulic fluid and pneumatic valves, and ensures the fuel
valve is closed. After engine isolation is accomplished, the fire switch is turned to the left or right to
discharge the left or right extinguisher bottles.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 94
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-52
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 95
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
THIS PAGE INTENTIONALLY LEFT BLANK
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 96
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE INDICATING SYSTEM
Overview
The engine indicating system is comprised of: Engine Pressure Ratio (EPR) Indicating System, Engine
Tachometer System, Exhaust Gas Temperature (EGT) Indicating System, Airborne Vibration Monitoring (AVM)
System, Electronic Engine Control (EEC) Monitoring System, and Standby Engine Indication System.
EPR Indicating System
The EEC senses PT4.9 input and uses the total pressure input (P2) from the Air Data Computer (ADC) to
calculate EPR which equals the ratio of PT4.9/P2. EPR parameters are transmitted from the EEC to EICAS
over a 429 data bus, and are displayed on the upper EICAS display.
Engine Tachometer System
The engine tachometer system provides both N1 and N2 speeds to EICAS. N1 speed is displayed on the upper
EICAS display, and N2 speed is displayed on the lower EICAS display.
EGT Indicating System
EGT is sensed by thermocouple probes located in the struts of the turbine exhaust case. The EEC transforms
the EGT probe signals into signals suitable for transmission on a 429 data bus. EGT parameters are displayed
on the upper EICAS display. The EGT probe signals are also sent (in analog form) to the SEI.
AVM System
Engine vibration is sensed by accelerometers mounted on the engine case. The AVM monitor unit converts the
accelerometer signals into signals suitable for display on EICAS. Vibration levels are displayed on the lower
EICAS display.
EEC Monitoring System
The EEC monitoring system allows for testing and fault monitoring of the electronic engine controls. EEC fault
data are stored in non-volatile memory inside the EEC monitor unit.
System for the Standby Engine Indication
A liquid-crystal display provides a readout of EPR, N1, N2, and EGT parameters in case of EICAS failure.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 97
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
ENGINE PRESSURE RATIO (EPR) INDICATING SYSTEM
Overview
The indicating system for the engine pressure ratio (EPR) shows engine power output. The system is used for
setting engine thrust and for monitoring engine performance.
Engine inlet and exhaust pressures are sensed by probes and sent to the electronic engine control (EEC). The
EEC converts pressure signals to an EPR signal and sends the EPR signal to the flight compartment.
The system consists of a probe for the inlet total pressure/temperature, exhaust probes, EEC, EICAS, and the
standby engine indicator (SEI).
EPR Indication
The upper EICAS display shows a continuous analog and digital readout of EPR. The analog readout of EPR
consists of two white round dials with scale markings of 1.0, 1.4, and 1.8. Each scale graduation represent a
change of 0.1 EPR.
A reference limit readout that is green colored is displayed above each respective round dial to provide thrust
reference information. When the THRUST REF SET knob on the display select panel for the EICAS is pulled
out, manual EPR settings from 1.0 to 1.8 EPR can be displayed selectively or simultaneously. When the
THRUST REF SET knob is reset (pushed in), the thrust reference limits are provided by the TMC.
Actual thrust readout is enclosed in a white box directly below each reference limit readout. The actual thrust
pointer is a straight white line extending from the center of the dial to the edge of the scale.
The command sector is located adjacent to each scale and indicates the momentary difference between
commanded EPR and actual EPR. As engine thrust changes, the actual thrust pointer will move the
commanded thrust level, erasing the sector.
Maximum thrust limit is indicated by two yellow colored bars on the analog dial. The value is normally
acquired from the EEC, but is obtained from the TMC if the EEC fails.
Setting the SEI mode switch to AUTO causes the SEI display to blank unless an EICAS display failure occurs.
Turning the mode switch to ON causes the SEI display to be on continuously.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 98
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
IF PW-53
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 99
For Training Purposes Only
AeroEd LLC – Aviation Education Resource
Probe for the Inlet Total Pressure/Temperature (PT2/TT2)
The PT2/TT2 probe supplies inlet pressure to each channel of the EEC as a backup value for the EPR
calculation.
Exhaust Pressure (PT4.9) Probes
Exhaust pressure data is furnished by probes integral to the struts of the turbine exhaust case. A manifold
averages each pressure signal and routes the average pressure to the EEC.
Electronic Engine Control
The electronic engine control takes the inlet and exhaust pressure signals and calculates PT4.9/P2 (EPR).
Each channel of the EEC determines EPR and sends EPR data to the flight compartment via separate 429 data
buses.
757 General Familiarization (7-2005)
ATA 71-80 PW - Page 100
For Training Purposes O