I. A. E V2500 COURSE NOTES
I.A.E. V2500
COURSE NOTES
CONTENTS
INTRODUCTION TO V2500 PROPULSION UNIT
PART ONE - ENGINE
PART TWO - NACELLE
SECTION 1
INTRODUCTION
SECTION
1
INTRODUCTION
SECTION 2
MECHANICAL ARRANGEMENT
SECTION
2
MECHANICAL ARRANGEMENT
SECTION 3
ELECTRONIC ENGINE CONTROL
SECTION
3
THRUST REVERSER
SECTION 4
POWER MANAGEMENT
SECTION
4
NACELLE VENTILATION AND
SECTION 5
FUEL SYSTEM
FIRE PROTECTION
SECTION 6
OIL SYSTEM
SECTION
SECTION 7
HEAT MANAGEMENT SYSTEM
SECTION 8
COMPRESSOR AIRFLOW CONTROL
PART THREE - GENERAL
SYSTEM
SECTION
1
TROUBLE SHOOTING
SECTION 9
SECONDARY AIR SYSTEMS
SECTION
2
COMPONENT LOCATION GUIDE
SECTION 10
ENGINE ANTI-ICE SYSTEM
SECTION
3
ENGINE G.A. DIAGRAM
SECTION 11
ENGINE INDICATING
SECTION
4
BORESCOPING
SECTION 12
STARTING AND IGNITION SYSTEM
SECTION
5
TRIM BALANCING
5
ENGINE REMOVAL/INSTALLATION
PROPULSION UNIT - INTRODUCTION
I.A.E. V2500
PROPULSION UNIT – INTRODUCTION
The V2500 is an advanced technology aircraft propulsion unit
designed primarily for the 150 seat, short to medium range
aircraft.
V2500 is a new advanced design and incorporates technology
from five major engine manufacturers: Rolls-Royce
Pratt & Whitney
Japanese Aero
Engine Corporation
M.T.U.
Fiat Aviazion
- England
- U.S.A.
- Japan
- Germany
- Italy
The propulsion unit shown below is the V2500 for the Airbus
A320 Aircraft.
Propulsion Unit - Data
T.O. Thrust (S.L. Static)
Flat Rated Temperature
Total Airflow
By-pass Ratio
Overall Pressure Ratio
Fan Diameter
Propulsion Unit overall length
Engine overall length
Propulsion Unit Weight
Bare Engine Weight
: 25000 lbs (111205 KN).
: I.S.A. + 15°C
: 783 lbs (355kgs)/second
: 5.42:1
: 29.4:1
: 63 inches (160cm)
: 198.39 inches (503.91cm)
: 126 inches (320cm)
: 7300 lbs (3311kgs)
: 4942 lbs (2242kgs)
V2500 PROPULSION UNIT
Propulsion Unit Introduction
Gas Path
A simplified view of the propulsion unit is shown below. All
the air entering the engine passes through the inlet cowl to
the fan.
At the fan exit the air stream divides into two flows:the core engine flow
• the by-pass flow
the by-pass ratio is 5.42:1.
•
Core Engine Flow
The core engine flow passes through the fixed inlet guide
vanes to the three stage booster then to the HP
compressor, the combustion section and the HP & LP
turbines and finally exhausts into the CNA.
By-pass Flow
The fan exhaust air (cold stream) entering the by-pass duct
passes through the fan outlet guide vanes and flows along
the by-pass duct to exhaust into the CNA.
Common Nozzle Assembly (C.N.A.)
The core engine (hot) exhaust and the (coo1) by-pass flow
are mixed in the CNA before passing through the single
propelling nozzle to atmosphere.
PROPULSION UNIT OUTLINE
INTRODUCTION
ENGINE MARK NUMBERS
For easy identification of the present and all future variants
of the V2500, International Aero Engines has introduced a
new engine designation system.
•
All engines will retain V2500 as their generic name.
•
The first three characters of the full designation are V25,
identifying each engine as a V2500.
•
•
•
The next two figures indicate the engine's rated
sea-level takeoff thrust.
The following letter shows the aircraft manufacturer.
The last figure represents the mechanical standard
of the engine.
This system will provide a clear designation of a particular
engine as well as a simple way of grouping by name,
engines with similar characteristics.
• The designation V2500-D collectively describes,
irrespective of thrust, all engines for McDonnell
•
Douglas applications and V2500-A all engines for Airbus
Industrie.
Similarly, V2500-5 describes all engines built to the 5
mechanical standard, irrespective of airframe
application.
The only engine exempt from these is the current service
engine, which, having already been certified, will retain the
original and current designation V2500-A1
V2530-A5
V25
: Generic to all V2500 engines
30
: Take off thrust in thousands of pounds
A(D) : Air frame manufacture
5
A : for Airbus Industrie
D : for McDonnell Douglas
: Mechanical standard of engine
MK NO
Takeoff Thrust (lb)
Aircraft
V25OO - A1
25,000
A32O-20O
V253O - A5
3O,000
A321-100
V2525 - A5
2 5,000
A3 2O-2OO
V2528 - D5
2 8,000
MD-9O-4O
V2522 - D5
2 2,000
MD-9O-10
V2500
COURSE NOTES
PART ONE ENGINE
V2500
COURSE NOTES
PART ONE ENGINE
CONTENTS
SECTION 1
SECTION 2
SECTION 3
SECTION 4
SECTION 5
SECTION 6
SECTION 7
SECTION 8
SECTION 9
SECTION 10
SECTION 11
PART ONE - SECTION 1
ENGINE INTRODUCTION
PROPULSION UNIT
INTRODUCTION
INTRODUCTION - ENGINE
The V2500 is a twin spool, axial flow, high by-pass ratio turbofan.
The engine incorporates several advanced technology features
which include:•
•
•
•
•
Full Authority Digital Electronic Control - FADEC
Wide chord fan blades
Single crystal HP. turbine blades
Powdered Metal HP. turbine discs .
A two piece, annular combustion system which
utilises segmental liners.
Engine Mechanical Arrangement
Low Pressure (LP) spool
•
The low pressure (LP) spool comprises:
o a single stage fan
o three stage axial flow (booster) or LP Compressor,
o five stage LP turbine.
•
The booster stage has an annular bleed valve to improve
starting and handling.
LP spool speed is indicated as N1 (%).
The LP Spool is supported in three bearings, one ball and
two roller.
•
•
High Pressure (HP) spool
•
The HP spool comprises:
o a ten stage axial flow compressor
o two stage turbine.
•
The HP compressor has:
o variable inlet guide vanes (VIGVs),
o four stages of variable stator vanes (VSVs) and
o four bleed valves which are all used to improve starting
and handling.
HP spool speed is indicated as N2 (%).
•
•
The HP spool is supported in two bearings, one ball and one
roller.
Lubrication
The lubrication system is:
• self-contained,
• re-cirulatory,
• full flow (unregulated pressure).
Primary oil cooling is by a fuel/oil heat exchanger located in the
LP fuel system, additional cooling, as required, is provided by
an air/oil heat exchanger.
V2500 ENGINE CUTAWAY
Introduction
Active Clearance Control A.C.C. (Turbine)
Full Authority Digital Electronic Control (FADEC)
Active clearance control (ACC) is used on both the LP and
HP turbine casings, this system uses cool air taken from
the fan duct.
The heart of the FADEC is the Electronic Engine Control
(EEC).
Engine Air Bleeds
Engine air bleed is utilised for: •
•
•
•
•
•
Aircraft systems
Compressor stability system
HP and LP active clearance control
10th stage make-up cooling air (turbine cooling)
Air cooled air cooler (buffer air)
Air cooled oil cooler
Customer Services Bleed
HP compressor stage 7 and stage 10 bleeds are
available to the aircraft manufacturer.
The EEC receives rotor speed, pressure and temperature
signals from the engine, it uses these parameters along
with aircraft inputs to command outputs to engine
mounted actuators to provide control of:• Engine fuel flow
• Automatic engine starting
• Compressor airflow control system
• Heat Management system
• 10th stage make-up air system
• thrust reverser
The E.E.C. also provides protection for:• Nl overspeed
• N2 overspeed
• Engine surge
SLTO PRESSURE - TEMPERATURE MAP
ENGINE DIMENSIONS AND PRESSURE
TEMPERATURE MAP
V2500
AIR OFFTAKES
FAN AIR
•
•
•
•
•
•
COOLING FLOW FOR ACAC.
ACTIVE CLEARANCE CONTROL SYSTEM (HP TURBINE &
LP TURBINE).
COOLING FLOW FOR AIR COOLED OIL COOLER
(ACOC).
PRE-COOLER, CUSTOMER SERVICES BLEED COOLING
FLOW.
IGNITION EXCITERS & HIGH TENSION LEADS COOLING
C-DUCT ACTUATORS & OIL SUPPLY PIPE COOLING
STAGE 7
•
•
•
STAGE 8
•
BOOSTER - 2.5 HANDLING BLEED VALVE (B.S.B.V.)
STAGE 6
•
SEALING - FRONT BEARING COMPARTMENT (CARBON
SEALS) (1 SERIES ENGINE ONLY)
INLET COWL ANTI-ICING
HANDLING BLEED VALVES
CUSTOMER SERVICES BLEED:o E.C.S.
o WING ANTI-ICING
o POTABLE WATER TANK
o HYD HEADER TANK
•
COOLING
o H.P. COMPRESSOR
o L.P. TURBINE CAVITY
SEALING
o FRONT BEARING COMPARTMENT HYD SEAL
(INTERSHAFT)
o No. 5 bearing compartment (front seal)
V2500
AIR OFFTAKES
STAGE 10
•
•
CUSTOMER SERVICE BLEED
MAKE UP AIR SYSTEM; ADDITIONAL COOLING FOR
SPACE BETWEEN 1 & 2 HP TURBINE, DISCS & STAGE 2
HP TURBINE BLADES
•
No 4 BEARING SCAVENGE VALVE SUPPLY (CONTROL
PARAMETER & MUSCLE AIR)
HANDLING BLEED VALVES
H.P.T. STAGE 2 NGV's
•
•
STAGE 12
•
•
•
•
•
•
BUFFER AIR - No 4 BEARING CHAMBER COOLING
FLOW
MUSCLE AIR FOR HANDLING BLEED VALVES
STAGE 1 HPT NGV's COOLING
STAGE 1 HPT DISC FRONT FACE COOLING.
STAGE 1 HPT BLADES COOLING.
INNER & OUTER COMBUSTOR LINERS.
PART ONE - SECTION 2
ENGINE MECHANICAL
ARRANGEMENT
Mechanical Arrangement
General
The engine is an axial flow, high by-pass ratio, twin spool,
turbo fan.
The general arrangement is shown below.
LP System
Four stage L.P. compressor comprising:
o 1 Fan stage
o 3 Primary stages, driven by a 5 stage axial flow LP
Turbine
• An annular bleed valve is located at the outlet from the
booster stage
•
HP System
•
•
•
•
•
Ten stage axial flow compressor driven by a 2 stage
axial flow HP Turbine
Variable angle Inlet guide vanes
4 stages of variable stator vanes
Handling bleed valves on stage 7 and stage 10
Customer service bleed at stages 7 and 10
Combustion System
•
Annular, two piece, with 20 fuel spray nozzles
Gearbox
•
•
Radial drive via a tower shaft from HP Compressor shaft
to fan case mounted Angle and Main gearboxes .
Gearbox provides mountings and drive for all engine
driven accessories and the pneumatic starter motor
ENGINE - GENERAL ARRANGEMENT
Mechanical Arrangement
Engine Main Bearings
The main bearing arrangement, and the bearing
numbering system is shown below.
The five bearings are located in three bearing
compartments.
•
The front bearing compartment:
o located at the center of the intermediate case,
o houses No's 1,2 & 3 bearings
•
The center bearing compartment located in the
diffuser/combustor case houses No 4 bearing .
•
The rear bearing compartment located in the turbine
exhaust case houses No 5 bearing
No 1 Bearing
•
•
•
LP shaft axial location bearing
Takes the thrust loads of the LP shaft
Single track ball bearing
No 2 Bearing
•
•
•
Radial support for the front of the L.P. turbine shaft
Single track roller bearing
Squeeze film oil damping
No 3 Bearing
•
•
•
•
•
HP shaft axial location bearing
Radial support for the front of the HP shaft
Takes the thrust loads of the H.P. shaft
Single track ball bearing
Mounted in an Hydraulic damper which is centred
by a series of rod springs (Squirrelcage).
No 4 Bearing
•
•
Radial support for turbine end of HP shaft
Single track roller bearing
No 5 Bearing
•
•
•
Radial support for the turbine end of the LP shaft
Single track roller bearing
Squeeze film oil damping
No 2 BEARING
No 1 BEARING
No 3 BEARING
No 4 BEARING
No 5 BEARING
ENGINE - MAIN BEARINGS
Mechanical Arrangement
Bearing Compartments
Front Compartment
Gearbox Drive
The No's 1, 2 and 3 bearings are located in the front
bearing compartment which is at the centre of the
intermediate module (32).
The HP stubshaft, which is located axially by No 3
bearing, has at its front end a bevel drive gear which,
through the tower shaft, provides the drive for the main
accessory gearbox.
The compartment sealing utilises carbon seals, brush
seals and sealing airflows obtained from the 6 th stage
compressor manifold.
The H.P. stubshaft separates from the HP compressor
module at the curvic coupling and remains as part of the
intermediate case module.
An oil filled (hydraulic) seal is used between the two
shafts, this seal is supported by 8 th stage air.
Adequate pressure drops across the seals to ensure
satisfactory sealing, are achieved by venting the
compartment, by an external tube, to the de-oiler. The
bearing compartment pressure, and therefore the sealing
flows, are controlled by a restrictor in the vent tube.
FRONT BEARING COMPARTMENT
Bearing Compartments
Front Compartment Continued
The drawing below shows details of No2 and No3
bearings.
A phonic wheel (1) is fitted to the LP stub shaft, this
interacts with speed probes to provide LP shaft speed
signals (Nl) to the EEC.
The hydraulic seal (6) prevents oil leakage from the
compartment passing rearwards between the HP and LP
shafts.
No3 bearing is hydraulically damped. The outer race is
supported by a series of spring rods (14) which allow
some slight radial movement of the bearing. The bearing
is centralised by oil pressure fed to an annulus (12)
around the bearing outer race.
The gearbox drive gear (8) is splined onto the HP shaft
and retained by No3 bearing nut (7).
1 Fonic wheel
2 LP Stub shaft
3 No.2 bearing
4 Squeeze film damper
5 No.2 bearing support
6 Hydraulic seal
7 No.3 bearing nut
8 Internal gearbox driving gear
9 No.3 bearing rotor center
10 No.3 bearing
11 No.3 bearing housing
12 Hydraulic damper
13 No.3 bearing seat support
14 Spring rod
15 No.3 bearing rear oil sea)
No 2, No 3 BEARING ARRANGEMENT
Mechanical Arrangement
Bearing Compartments
No 4 (Centre) Bearing
The No 4 bearing compartment is situated in a high
temperature, high pressure environment at the centre of
the combustor section.
The bearing compartment is encased in a thermally
insulated shielding.
A supply of cooled air (buffer air) is admitted to the space
between the chamber double skinned walls.
The buffer air is exhausted from the cooling spaces close
to the upstream side of the carbon seals, creating an area
of cooler air from which the seal leakage is obtained. This
results in an acceptable temperature of the air leaking into
the bearing compartment.
Buffer air flow rates are controlled by restrictors at the
outlet from the cooling passages.
The bearing compartment internal pressure level is
determined by the area of the variable scavenge valve,
(called No 4 bearing scavenge valve and described in the
oil system). This valve acts as a variable restrictor in the
compartment vent line.
No 4 BEARING COMPARTMENT
Mechanical Arrangement
bearing Compartments
No 5 Bearing (Rear)
The rear bearing compartment is located at the centre
of the L.P. turbine module (module 50) and houses
No 5 bearing which supports the L.P. turbine rotor.
The compartment is sealed at the front end by a
carbon seal, a simple labyrinth seal provides
secondary sealing to protect against oil loss in the
event of carbon seal failure.
Separate venting is not necessary for this compartment
because with only one carbon seal the airflow induced by
the scavenge pump gives the required pressure drop
across the seal.
The compartment is covered by an insulating heat
shield.
REAR BEARING COMPARTMENT
STAGE 8
SEALING AIR
LP TURBINE SHAFT
CARBON SEAL
REAR BEARING COMPARTMENT
Mechanical Arrangement:
Engine Stations
Stage Numbering
1
:Intake/Engine inlet interface
Compressor
2
:Fan inlet
Stage 1
2.5
:LP Compressor exit 12.5 :Fan exit
Stage 1.5 :Booster Stage
3
:HP Compressor exit
Stage 2
4
:Combustion section exit
Stage 2.5 :Booster Stage
4.5
:HP Turbine exit
4.9
:LP Turbine exit
Stage 3
to
Stage 12
Pressure and Temperature Signals
The following pressures and temperatures are sensed and
transmitted to the Electronic Engine Control (EEC):
P2
T2
P2.5
T2.5
P3 (Pb) T3
P4.9
P12.5
T4.9
:Fan
:Booster Stage
:HP Compressor Stages
:HP Compressor Stages
Turbine
Stage 1 }
Stage 2 } :HP Turbine Stages
Stage 3 }
to
}
Stage 7 }
:LP Turbine Stages
STATIONS
2
ROTATING STAGES
12,5 2.5
1
1.5 2 2.5
3
4
4
5 6 7 8 9 10 11 12 1
COMPRESSOR STAGES
4.5
4.9
2 3 4 5 6 7
TURBINE STAGES
ENGINE STATION AND STAGE NUMBERING
ENGINE
Mechanical Arrangement
Modular Construction
Modular construction has the following advantages:•
•
•
•
•
•
•
•
•
•
•
lower overall maintenance costs.
maximum life achieved from each module.
reduced turn-around time for engine repair.
reduced spare engine holdings.
reduced transportation costs.
ease of transportation and storage.
rapid module change with minimum ground running
easy hot section inspection.
vertical/horizontal build strip.
split engine transportation.
compressors/turbines independently balanced.
Module Designation
Module No
31
32
40
41
45
50
60
Module
Fan
Intermediate
HP System
HP Compressor
HP Turbine
LP Turbine
External gearbox
(31) FAN
(32) INTERMEDIATE
(50) LOW PRESSURE
(60) EXTERNAL GEARBOX
ENGINE MODULES
Mechanical Arrangement
Module 31 - Description
Module 31 (Fan Module) is the complete Fan assembly
and comprises:
•
•
•
•
2 hollow fan blades
2 annulus fillers
the fan disc
the front and rear blade retaining rings
The blades are retained in the disc radially by the dovetail
root.
Axial retention is provided by the front and rear blade
retaining rings.
Blade removal/replacement is easily achieved by
removing the front blade retaining ring and sliding the
blade along the dovetail slot in the disc.
22 annulus fillers form the fan inner annulus.
The nose cone and fairing smooth the airflow into the fan.
REAR BLADE RETAINING RING
WIDE CHORD FAN BLADES (22)
FRONT BLADE RETAINING
FAIRING
FAN DISK
NOSE CONE
ANNULUS FILLERS (22)
LP COMPRESSOR (FAN)
Mechanical Arrangement
Fan
Nose Cone
The Glassfibre cone smoothes the airflow into the fan. It is
secured to the front blade retaining ring by 18 bolts.
The nose cone is balanced during manufacture by
applying weights to its inside surface.
The nose cone is unheated. Ice protection is provided by
a soft rubber cone tip.
The nose cone retaining bolt flange is faired by a titanium
fairing which is secured by 6 bolts.
NOTE
Be careful when removing the Nose Cone retaining bolts.
Balance weights may be fitted to some of the bolts. The
position of these weights must be marked before removal
to ensure they are refitted to the same position.
The arrangement is shown below.
TRIM BALANCE
WEIGHT
LP COMPRESSOR
{Fan} MODULE
TRIM BALANCE
WEIGHT
FAIRING
TRIM BALANCE
WtiGHT
FRONT RETAINING
RING
NOSE CONE
Mechanical Arrangement
Fan - Front Blade Retaining Ring
Note:
The blade retaining ring is secured to the fan disc by a ring
of 36 bolts. A second {outer) ring of bolts passes through
the retaining ring and screws into each of the 22 annulus
fillers.
The fan blades and annulus filler positions are not
identified. For this reason it is important to identify the
blade and annulus filler position, relative to the numbered
slots in the fan disc, before disassembly.
Both rings of bolts roust be removed before attempting to
remove the front retaining ring.
This is done using a temporary marker.
After all the securing bolts (22 + 36) have been removed
the retaining ring can be removed by screwing pusher
bolts into the 6 threaded holes provided for this purpose.
Balance weights, if required are located on the retaining
ring, as shown below.
The front blade retaining ring can only be fitted in one
position which is determined by three off-set locating
dowels on the fan disc.
1
2
3
4
5
6
7
8
9
10
11
12
13
T mark
Front blade retaining ring
Stage 1 fan disk
Annulus filler
Guide pin
Headless pin (3 off)
Stage 1 fan blade
Pusher threaded hole (6 off)
Boit (22 off}
Bolt (36 off)
Balance weight flange
BaSance weight
fim balance weight
(engine pass off)
10
LP COMPRESSOR (FAN) BLADE RETAINING RING
Mechanical Arrangement
Fan - Fan Blades and Annulus Fillers
After removal of the Front Blade retaining ring the annulus
fillers can be removed as follows:•
•
•
•
•
lift the front end of the annulus filler 3 to 4 inches
twist the annulus filler through about 60 deg counter
clockwise
draw the annulus filler forward to clear the blades
Remove the annulus fillers on either side of the blade
to be removed.
The blade to be removed can than be pulled forward
to clear the dovetail slot in the fan disc.
1 Stage 1 fan disk
2 Stage 1 fan blade
3 Annulus filler
FAN BLADE/ANNULUS FILLER
Mechanical Arrangement
Fan Blades Inspection/Repair
Fan blade inspection / repair procedures are briefly
described in these notes. This information is for guidance
only.
Before any repair is carried out reference must be made to
the Maintenance Manual Chapter 72-31-11 Page Block
800.
General
The fan blade surface area is divided into three zones A,B
and C as shown below. The acceptance limits for damage
may vary depending on which zone is damaged.
Inspection
Blades are inspected for signs of nicks, cracks, dents,
scores, scratches on the surface, and bends on the
leading or trailing edges. The blades should also be
inspected for signs of arc burns (lightning strikes).
Arc burns or cracks are grounds for rejection and
a replacement blade must be fitted.
The acceptance limits for nicks, scratches, scores and
dents are detailed below. If damage exceeds these limits
refer to the appropriate repair scheme (see MM 72-31-11
Page Block 800 - Approved Repairs).
LEADING EDGE
TRAILING EDGE
0.75m
(19,05 mm)
2.00m
(50,80 mm)
12.00m
(304,80 mm)
1.50 in
(38,10 mm)
3.00 in
(76,20 mm)
MAXIMUM SERVICEABLE LIMITS FOR SURFACE DAMAGE
DEPTH ON CONVEX AND CONCAVE SURFACES
A 0.008in (0,20mm)
B 0.008in (0,20mm)
STAGE 1 FAN BLADES REPAIR LIMITS
Mechanical Arrangement
Fan Blades Inspection/Repair - General
Acceptance Limits
The leading and trailing edges of the blades should
be examined for bends (deformations).
X maximum = 0.2 in (5,08mm)
if X is more than 0.2 in reject the blade.
Note:
Z must be not less than 8 times dimension X
if Z is less than 8 times X reject the blade.
•
the maximum number of bent blades in any fan rotor
assembly is three.
•
no blade may have more than one bend.
•
if any bend has associated cracks, kinks, creases, tears
or nicks -reject the blade.
•
bends must be outboard of the annulus fillers, if any
bend extends below the annulus filler platform reject the
blade.
Y must be not less than 20 times X
if Y is less than 20 times X reject the blade.
Note:
There must be a smooth transition between the
undamaged aerofoil surface and the bent area.
if there is not a smooth transition reject the blade.
STAGE 1 FAN BLADES REPAIR LIMITS
Mechanical Arrangement
Annulus Fillers - Inspection/Repair
The outer surfaces of the annulus fillers should be inspected
for cracks, nicks, dents, scores.
•
if any cracks are found reject the annulus filler.
•
accept dents, nicks, scores up to 0.010 inches (0.25mm)
•
annulus fillers with damage in excess of 0.010 may
be repaired in accordance with the appropriate
repair scheme (see MM 72-31-11 page block 800)
1 0.010 in. (0,25rnm) maximum blend depth
2 Apply heat resistant ES coating to this area
3 0.400in. (10,16mrn) minimum
No blending permitted in this area
ANNULUS FILLERS - REPAIR LIMITS
Mechanical Arrangement
Fan Blades - Cropping
Before carrying out any blade repairs refer to the
Maintenance Manual 72-31-11 page block 800 - Approved
Repairs.
The following pages illustrate typical cropping and
scalloping limits.
The following points must be noted:•
always do a dye penetrant crack test when the
repair is completed.
•
when t he repair is completed write the repair
scheme number e.g. VRS 1002, in the engine log
book.
•
at the next shop visit, after repair, the repair
scheme number should be etched on the blade
root.
•
blades that are repaired on wing must be glass
bead peened at the next overhaul.
FAN BLADE CROPPING LIMITS
Mechanical Arrangement
Fan Blade Repairs
Shown below are examples of blades of blade cropping and
scalloping limits.
TYPICAL LEADING
AND TRAILING EDGES
One scallop only
is permitted in this zone.
POINT
AP
POINT
AR
Two scallops only are
permitted on leading and/or
trailing edges provided that
all other conditions are met
in this zone.
One scallop only is Pitted
in this zone on lead.ng or
trailing edge.
(21,59mm)
One scallop only is permitted
on leading or trailing edge.
FAN BLADE MINIMUM ACCEPTABLE CONDITIONS BETWEEN
TIP CROPPING AND SCALLOPING.
The minimum dimensions quoted apply to any combination of
cropping and scalloping in the blade tip area.
If tip cropping has not been carried out, these dimensions
apply from the blade tip points AR and AP
FAN BLADE SCALLOPING LIMITS
(21,59mm)
Mechanical Arrangement
Fan Blade Repairs
Further examples of scalloping proportions and limits are shown
below.
10XAB
SCALLOP PROPORTIONS
MIN R = 7xAB
MINR =
7xAB
Permitted only if extents
of blend rads do not overlap
— AB
,1AX = AB + 0.064«p, 11,63mm)
MIN = AB
AB —
TYPICAL
BLENDING
AREA
FAN BLADE SCALLOP PROPORTIONS
0.858in (21,79mm) AT TIP
R
0.Q43in(1.09 mm)
0.026in (0,66mm)
LEADING EDGE
VIEW AB
8.955in (227,46mm)
. BLEND LEADING EDGE
BACK TO THIS LINE
4.220in
(107,18mm)
R 2.532in (64,31mm)
LEADING EDGE
TRAILING EDGE
FAN BLADE FLY BACK REPAIRS
FAN BLADE FLY BACK REPAIRS
Mechanical Arrangement
Module 32 - Intermediate Case
The intermediate module comprises:
• the fan case.
• the fan duct.
• the fan outlet guide vanes.
• the three stage L.P.'Booster' compressor.
• the booster stage bleed valve (B.S.B.V.).
• the front engine mount structure.
• the front bearing compartment which houses:
o No 1, 2 and 3 bearings.
o the drive gear for the power off-take
shaft (gearbox drive).
• the L.P. stub shaft.
• 10 inner support struts.
• 10 outer support struts
• Vee gioove locations for the inner and outer barrels of
the C-ducts
Instrumentation
The following pressures and temperatures are sensed and
transmitted to the E.E.C.:•
P12.5
•
P2.5
•
T2.5
1 Fan case
2 Fan case front panel fairing
3 LP compressor outlet guide vane
4 Intermediate structure front fairing
5 LP compressor inlet guide vane
6 Inner ring
7 LP stub shaft curvic teeth
FORWARD
INTERMEDIATE MODULE - FRONT VIEW
Mechanical Arrangement
Module 32 - Intermediate Case
The rear view of the intermediate case is shown below.
1
2
3
4
5
6
7
Fan case
Fan case rear panel
Fan frame outer strut
Intermediatestructure
Forward mount
Fan frame inner strut
No.3 bearing curvic teeth
INTERMEDIATE CASE - REAR VIEW
Mechanical Arrangement
Nodule 40 - H.P. Compressor
The general arrangement of the H.P. compressor is shown
below.
The H.P. compressor has 10 stages. It utilises variable
inlet guide vanes at the inlet to stage 3 and variable stator
vanes on stages 3, 4, 5 and 6.
The front casing, which houses stages 3 to 6, is made in
two halves which bolt together along horizontal flanges, it
is bolted to the intermediate casing (module 32) at the
front and to the outer casing at the rear.
The rear compressor casing has inner and outer casings
as shown. Flanges on the inner case form annular
manifolds which provide 6, 7 and 10 stage air offtakes.
HP
COMPRESSOR
ROTOR
ASSEMBLY
REAR INNER
CASE
REAR OUTER
CASE
FRONT
COMPRESSOR
CASES
VARIABLE
STATOR
VANE
VSV OPERATING
MECHANISM
HP COMPRESSOR
Mechanical Arrangement
H.P. Compressor
Compressor Drums - (Rotor)
The rotor assembly is in two parts:
• the stage 3 to 8 drum
• the stage 9 to 12 drum
The two rotor drums are bolted together with a vortex
reducer installed between the 8 and 9 stages.
The vortex reducer straightens the stage 8 air flow which
passes to the centre of the engine for internal cooling and
sealing.
Mechanical Arrangement
Combustion Section
The combustion section includes the diffuser section, the
combustion inner and outer liners, and the No 4 bearing
assembly.
Diffuser Casing
The diffuser section is the primary structural part of the
combustion section.
The diffuser section has 20 mounting pads for the
installation of the fuel spray nozzles. It also has two
mounting pads for the two igniter plugs.
Combustion Liner
The combustion liner is formed by the inner and outer
liners.The outer liner is located by five locating pins which pass
through the diffuser casing.
The inner combustion liner is attached to the turbine
nozzle guide vane assembly.
The inner and outer liners are manufactured from sheet
metal with 100 separate liner segments attached to the
inner surface. The segments can be replaced
independently.
COMBUSTION SECTION (1)
Mechanical Arrangement
Combustion Section
The drawing below shows the arrangement of the diffuser
casing and the outer section of the combustion liner.
Also shown is the front section of the No 4 bearing
compartment.
COMBUSTION SECTION (2)
DIFFUSER CASE
ASSEMBLY
HP COMPRESSOR
EXIT STATOR
FUEL NOZZLE
(Cool air flow)
OUTER COMBUSTION
CHAMBER LINER
No. 4 BEARING FRONT
HEATSHIELD
THRUST BALANCE
STATIC SEAL
No, 4 BEARING
SUPPORT ASSEMBLY
No. 4 BEARING FRONT
COOLiNG AIR DUCT '
No. 4 BEARING
LOCK AND NUT
No. 4 BEARING FRONT
COMPARTMENT
No. 4 BEARING
FRONT SEAL
ASSEMBLY
COMBUSTION SECTION (2)
Mechanical Arrangement
Combustion Section
The drawing below shows the arrangement of the inner
combustion liner and the H.P. stage 1 nozzle guide vanes.
Also shown is the cooling air inlet arrangement which
provides the cooling air supplies for the H.P. turbine 1st
stage disc and turbine blades.
The cooling air duct is known as the Tangential On Board
Injection (T.O.B.I.) duct.
STAGE 1 HPT
SUPPORT ASSEMBLY
(Coo! air flow)
COMBUSTION CHAMBER
INNER LINER
STAGE 1 HPT VANE
CLUSTER ASSEMBLY
STAGE 1 HPT DUCT
SEGMENT
(Cool air flow}
STAGE 1 HPT COOLING
DUCT ASSEMBLY
(Cool air flow)
COMBUSTION SYSTEM (3)
Mechanical Arrangement
H.P. Turbine
Shown below is the arrangement of the H.P. Turbine.
Cooling airflows are also shown.
10th STAGE COMPRESSOR AIR
FOR 2nd VANE AND 1-2 SEAL
TOB! FEED
TO 1st BLADE
10th STAGE AND HPC
DISCHARGE"AIR
TO 2nd BLADE
HP TURBINE ASSEMBLY
Mechanical Arrangement
Module 50 - L.P. Turbine
The five stage LP turbine extracts energy from the gas
stream to provide the rotational drive for the L.P.
compressor and fan.
The four principal elements of the LP Turbine Module are:
•
•
•
•
LP Turbine case, vanes and static seals
Five stage LP Turbine rotor
LP Turbine shaft
Turbine exhaust case
Seal clearance and L.P turbine case thermal expansion
are controlled by an external Active Clearance
Control.(A.C.C.) system. The A.C.C. system uses fan
discharge air which is directed externally to the L.P.
turbine case via the eight A.C.C. tubes.
Two boroscope ports are provided on the case, one on
each side. These ports enable inspection of the L.P.
turbine (stage 3) rotor blades and also stage 2 H.P.
turbine rotor blades (rear side). Each port is sealed by a
plug which incorporates features to prevent incorrect
installation.
Axial positioning of the L.P. turbine rotor assembly is
achieved by selection of an appropriate adjusting washer
fitted at the front end between the L.P. turbine shaft and
the L.P. compressor stubshaft.
The L.P. turbine shaft is supported at the front by No 2
bearing and at the rear by No 5 bearing.
The turbine exhaust case serves to straighten the gas
flow, provides structural support for the No 5 bearing and
incorporates the rear engine mount lugs. The struts
incorporate provision to sense exhaust gas temperature
T4.9 and pressure P4.9.
REAR ENGINE MOUNT LUGS
STAGE 3 NGV
LP TURBINE SHAFT
LP TURBINE NUT
ACC COOUNG
AIR INLET
TURBINE EXHAUST CASING
ADJUSTING WASHER
STUBSHAFT SPLINES
ACC TUBES
LOW PRESSURE TURBINE MODULE
Mechanical Arrangement
L.P. Turbine Exhaust Case
The L.P. Turbine exhaust case provides the support for No
5 bearing.
The hollow support struts provide the location for the 4
thermocouples which measure the E.G.T. (T4.9). Three of
the struts also house the P4.9 pressure measuring rakes.
The casing also provides the rear engine mounting
location.
OUTER RING
ENGINE MOUNTS
INNER RING
OIL SCAVENGE TUBE
FORWARD
PRESSURE CONTROL MANIFOLD
TURBINE EXHAUST CASE ASSEMBLY
TURBINE EXHAUST
CASE ASSEMBLY
Mechanical Arrangement
Nodule 60 - External Gearbox
The gearbox assembly transmits power from the engine to
provide drives for the accessories mounted on the
gearbox front and rear faces. During engine starting the
gearbox also transmits power from the pneumatic starter
motor to the engine.
The gearbox also provides a means of hand cranking the
H.P. rotor for maintenance operations.
Location
The gearbox is mounted by 4 flexible links to the bottom of
the fan case.
• main gearbox 3 links
• angle gearbox 1 link
Type
Cast aluminium housing.
Features
• individually replaceable drive units
• magnetic chip detectors
o main gearbox 2 magnetic chip detectors
o angle gearbox 1 magnetic chip detector
Front Face Mount Pads
•
•
•
•
•
De-oiler
Pneumatic starter
Dedicated generator
Hydraulic pump
Oil Pressure pump
Rear Face Mount Pads
•
•
•
Fuel pumps and Fuel Metering Unit (FMU)
Oil scavenge pumps unit
Integrated Drive Generator (I.D.G.)
OIL FILTER
FUEL PUMP DRIVE PAD
O!L SCAVENGE PUMP
MAIN GEARBOX
DEOILER
INTEGRATED DRIVE GENERATOR
SYSTEM (iDGS) DRIVE PAD
STARTER
DRIVE PAD
DEDICATED
ALTERNATOR
(PMA)
FRONT VIEW
OIL TANK
HYDRAULIC PUMP
DRIVE PAD
OIL PRESSURE
PUMP
ANGLE AND MAIN GEARBOX
Mechanical Arrangement
Engine Drain System
Leakage and drainage from the engine accessories and
fuel operated actuators is conducted by tubes to the
engine drains mast.
The drains mast discharges to atmosphere through the
bottom of the fan cowls.
10
1
9
2
8
3
4
1 : OIL TANK SCUPPER DRAIN
2 : HYDRAULIC PUMP SEAL DRAIN
3 : AIR STARTER MOTOR SEAL DRAIN
4 : IDG SEAL DRAIN
5 : AIR COOLED OIL COOLER
ACTUATOR DRAIN
6 : DRAIN MAST
7 : CORE ENGINE DRAINS
8 : FUEL DIVERTER VALVE DRAIN
9 : FUEL PUMP SEAL DRAIN
7
5
GASKET
6
BOLT (TYPICAL
4 PLACES)
10 : FUEL MODULATING DRAIN UNIT
ENGINE DRAIN MAST-INSTALLED
PART ONE - SECTION 3
ELECTRONIC ENGINE
CONTROL (E.E.C.)
Electronic Engine Control
Introduction
The V2500 uses a Full Authority Digital Electronic Engine
Control (FADEC).
The FADEC comprises the sensors and data input, the
electronic engine control unit (E.E.C.) and the output
devices which include solenoids, fuel servo operated
actuators and pneumatic servo operated devices. The
FADEC also includes electrical harnesses.
Engine Electronic Control
The heart of the FADEC is the Engine Electronic Control
(EEC) unit. The EEC is a fan case mounted unit which is
shielded and grounded as protection against E.M.I. mainly
lightning strikes.
Features
•
•
•
•
•
•
•
Vibration isolation mountings
Shielded and grounded (lightning strike protection)
Size
:15.9 X 20.1 X 4.4 inches
Weight :41 lbs
Two independent electronic channels
Two independent power supplies from dedicated
generator
Built in handle facilitates removal and handling.
HARNESS
ANTI-ICE
DUCT
P2/T2 PROBE
SENSOR LINE
COOLING AIR OUTLET (2)
EEC COOLING
EJECTOR
EEC-CHANNEL (A)
ELECTRICAL CONNECTORS
EEC COOLING
OUTLET
ANTI-VIB
MOUNTINGS (4)
EEC - CHANNEL (B)
ELECTRICAL CONNECTORS
COOLING AIR INLETS (2)
ELECTRONIC ENGINE CONTROL
The EEC Description
The E.E.C. has two identical electronic circuits which are
identified as Channel A and Channel B. Each channel is
supplied with identical data from the aircraft and the
engine- This data includes throttle position, aircraft digital
data, air pressures, air temperatures, exhaust gas
temperatures and rotor speeds.
This data is used by the E.E.C, to set the correct engine
rating for the flight conditions. The E.E.C. also transmits
engine performance data to the aircraft. This data is used
in cockpit display, thrust management and condition
monitoring systems.
Each of the EEC channels can exercise full control of all
engine functions.
Control alternates between Channel A and Channel B for
consecutive flights, the selection of the controlling channel
being made automatically by the E.E.C. itself.
The dual channels which are contained in the two piece
housing are separated from each other through a unique
circuit mounting board system. The two channels
exchange data through the data crosslink.
ENGINE
ELECTRONIC
CONTROL
UNIT
EEC DESCRIPTION
Electronic Engine Control
Harness (electrical) and Pressure Connections
Electrical Connections
Two identical, but separate electrical harnesses provide
the input/output circuits between the EEC and the relevant
sensor/control actuator, and the aircraft interface.
The harness connectors are 'keyed' to prevent
misconnection.
Front Face
Note: Single pressure signals are directed to pressure
transducers - located within the EEC - the pressure
transducers then supply digital electronic signals to
channels A and B.
The following pressures are sensed:
•
Pamb
:ambient air pressure - fan case sensor
•
Pb
:burner pressure P3/T3 probe
•
P2
:fan inlet pressure - P2/T2 probe
•
P2.5
:booster stage outlet pressure
•
P5 (P4.9) :LP Turbine exhaust pressure - P5 (P4.9)
rake
•
P12.5
- fan outlet pressure - fan rake
Harness Connector Plug Identification
•
•
•
•
•
J1
J2
J3
J4
J11
E.B.U 4000KSA D202P
Engine D202P
Engine D203P
Engine D204P
Engine D211P
Rear Face
•
•
•
•
•
•
J5
J6
J7
J8
J9
J10
Engine D205P
Data Entry Plug
E.B.U. 4000 KSB
Engine D208P
Engine D209P
Engine D210P
COOLING AIR
INLET
J5
EEC
HARNESS
CONNECTORS
THIS SIDE
TOWARD FNGINE
EEC HARNESS
CONNECTORS
Pamb
VIEW A
Pb
P2.5
VIBRATION - ISOLATED
MOUNTS (4)
P2
P5
EEC - HARNESS / PRESSURE
CONNECTIONS
COOLING AIR
OUTLET
VIEW B
Electronic Engine Control
Cooling System
Internal cooling of the E.E.C. is provided by an induced
airflow.
The cooling air is drawn in through an inlet in the R-H. fan
cowl by an ejector which is located in the cooling air
exhaust duct.
The ejector uses 7th stage air tapped from the nose cowl
anti-icing ducting.
The induced airflow is supplemented by ram airflow during
forward movement of the aircraft.
HARNESS
EEC COOLING
AIR EXHAUST
ANTI-ICE DUCT
P2/T2 PROBE SENSOR LINE
EEC COOLING
EJECTOR
SEAL
EEC
UNiT
STARTER DUCT
EEC COOLING
AIR INLET
BRACKET
EEC EXHAUST
COOLING AIR
DUCT
EEC COOLING INLET
(ON FAN COWL DOOR)
ELECTRONIC ENGINE CONTROL-INSTALLATION
Engine Electronic Control (E.E.C.)
Overview
Fault Monitoring
The E.E.C. provides the following engine control
functions:-
The E.E.C- has extensive self test fault isolation logic built
in. This logic operates continuously to detect isolate
defects in the E-E.C.
•
•
•
•
•
•
•
•
•
•
•
•
•
•
Power Setting (E.P.R.)
Acceleration and deceleration times
Idle speed governing
Overspeed limits (N1 and N2)
Fuel flow
Variable stator vane system (V.S.V.)
Compressor handling bleed valves
Booster stage bleed valve (B.S.B.V.)
Turbine cooling (10 stage make-up air
system)
Active clearance control (A.C.C.)
Thrust revsrser
Automatic engine starting
Oil and fuel temperature management
Note: The fuel cut off (engine shut down) command
comes from the flight crew and is not controlled by the
EEC
Electronic Engine Control
Failures and Redundancy
Improved reliability is achieved by utilising dual sensors,
dual control channels, dual selectors and dual feedback.
•
Dual sensors are used to supply all E.E.C. inputs
except pressures, (single pressure transducers
within the E.E.C. provide signals to each channel A
and B).
•
The E.E.C. uses identical software in each of the
two channels. Each channel has its own power
supply, processor, programme memory and
input/output functions. The mode of operation and
the selection of the channel in control is decided by
the availability of input signal and output controls.
Each channel normally uses its own input signals
but each channel can also use input signals from
the other channel required i.e. if it recognises faulty,
or suspect, inputs.
•
An output fault in one channel will cause switchover
to control from the other channel.
•
In the event of faults in both channels a predetermined hierarchy decides which channel is
more capable of control
•
In the event of loss of both channels, or loss of
electrical power, the systems are designed to go to
the fail safe positions.
EEC-REDUNDANCY
Electronic Engine Control
Failures and Redundancy
In the event of loss of both input signals, loss of both
channels, or loss of electrical power, the system is
designed to go to the fail safe positions shown in the table
below.
•
Metering Valve Torque Motor
Minimum Fuel Position
•
Fuel Shut-off Valve
Last Commanded Position
•
Overspeed Valve Solenoid
Normal Fuel Metering
Seventh Stage Bleed Valves
Valve Open
Tenth Stage Bleed Valve
Valve Open
Combined Active Clearance Control unit
•
High A.C.C.
Valve Closed
•
Low A.C.C.
Valve Partially (45%) Open
Low Compressor (2.5) Bleed Actuator
Valve Open
Stator Vane Actuator
Vanes Open
Fuel Diverter and Back to Tank Valve
•
Fuel Diverter Valve
Solenoid De-energised (Mode 4 or 5)
•
Fuel Back to Tank Valve
Valve closed - No Return to Tank (Mode 3 or 5)
Air/Oil Cooler Control Valve Actuator
Valve Open
Tenth Stage "Make-up" Cooling Air Valve
Valve Open
* Thrust Reverser Control Unit
Reverser Stowed
PT2/TT2 Heater Relay Box
•
Ignition Relays
Ignition ON
•
Probe Heater Relays
Heater OFF
•
Starter Air Valve
Valve closed
Note: If there is a failure of the thrust reverser control unit
arming valve while the reverser is deployed, will remain deployed
Operation and Control
E.E.C. Power Supplies
The electrical supplies for the E.E.C. are normally
provided by a dedicated alternator which is mounted to
and driven by the external gearbox - shown below.
Dedicated Alternator
The unit is a permanent magnet alternator which has two
independent sets of stator windings and supplies two
independent, 3 phase, frequency wild A.C. outputs to the
EEC. These unregulated A.C. supplies are rectified to 28
volts DC within the EEC.
The dedicated alternator also supplies the N2 signal for
the EEC.
The EEC also utilises aircraft power to operate some
engine systems:•
115 volts AC 400 Hz power is required for the ignition
system and inlet probe anti-icing heater
•
28V DC is required for some specific functions which
include the thrust reverser, fuel on/off and ground test
power for EEC maintenance.
In the event of a dedicated alternator total failure the EEC
is supplied from the aircraft 28V DC bus bars, 28V DC
from the same source is also used by the EEC during
engine starts until the dedicated alternator comes on line
at approximately 10% N2.
The dedicated alternator comes on line and supplies the
EEC power requirement when the N2 roaches approx
10%.
Switching between the aircraft 28V supply and dedicated
alternator power supplies is dour automatically by the
EEC.
SHAFT SUPPORT
STATOR
DEDICATED ALTERNATOR
EEC Power Supplies
Electrical Harness
The E.E.C. power supplies and speed signal harness
connections are as shown below.
EEC ELECTRICAL POWER SUPPLIES
AND SPEED SIGNALS HARNESS
Electronic Engine Control
E.E.C. Electrical Connections
The harness connections, pin identification and data
identification for the E.E.C. junctions 1-11 are listed in the
next five pages,
These tables may be used in conjunction with the
electrical harness diagrams which appear in the
relevant engine system description.
A. 1
Sine
t II
»'. t
■IfANHEl
8TOO3W.-O
BT004H.M)
1'
u i
Cosine
Return
8T006B20
8T00SW20
T
J
High
Return
8TOO1W2O
8T002B20
h
Line A
Line B
8T021B20
8T929W29
g
t
Line A 8TJ12
Line B 8TJ13
Line A Line B
8TO18W2O
8T019B20
8T027W20
8T028B2O
p
N
e
M
Line A
Line B
8T017B20
8T016W20
2
d
Line A
Line B
8T014H20
8T015B20
Discrete
Return
L/Eng Discrete
R/Eng Discrete
Eng D'crete Rt
T
TRA Exci t .it I- n
i
i
s
h
«T»f><)W.'(i
8IO/0H.M(
!<■'( t i l n
8T072B20
8T071W?0
Cos i ne
Neturn
8T067W20
8T0e8B20
iiigh
Return
a
8T048B20
8T047W20
Line A
ARINC Input
ARINC Input From EIU
p
N
e
8T027W20
8T028B20
Line A
Line B
E
n
Line A
Line B
ARINC II Out
d
8T059B20
BTO58W2O
L
K
ARINC 112 Out
L
K
8T056W20
8TO57B20
Line A
Line B
8TU6B20
8T115W20
A
X
External Reset
A
X
8TU7W20
8T118B20
Discrete
Return
8T024B2O
8T022W20
8T023O20
r
3
r
8T066B20
8T064W20
L/Eng D;-5c-...ia
R/En§ Discs-ira
Eng D 'tiocs Kr
at
t
8T065O20
z
T/R Arming Valve Pressure Switch
s.
c
z
T/R Stow and Lock ncnsoi
s
8T060B20
8T061W20
8T062W20
8T0fi3B20
E
F
a
8T051B20
8TC49W2O
8T050O20
R
8T053W20
8T052B20
8T02SW20
8T026B20
Discrete Discrete
Rtn Discrete Rtn
6TJ11 Discrete
8T013W20
8T012B20
8TO11B2O
8TO1OW2O
On Off
Discrete Rtn
8T009B20
8T007W20
8T008O20
E
F
a
8T02SW20
8T026B20
R
n
Line A
Line B
u
C
Instinctive Disconnect (Autotthrust)
ARINC Input From ADI R.I 7
n
H.E.C. HARNESS CONECTIONS PIN NO S
Line B
Di sere ■'■.■■:
'■■•-8TJO3
.* K
DL*t.i.^:~.
Discra--
on
Off
Discrete R
8TJ02 Line A
Lirse B Line A
8TJ01 Line B
PART ONE - SECTION 4
POWER MANAGEMENT
Power Management
Throttle Control Lever Mechanism
The throttle control lever (Thrust Lever) is based on the "fixed
throttle" concept, there is no motorised movement of the throttle
levers.
Each throttle control lever drives dual throttle resolvers, each
resolver output is dedicated to one E.E.C. channel.
The throttle lever angle (TLA) is the input to the resolver.
The resolver output, which is fed to the E.E.C, is known as the
Throttle Resolver Angle (TRA)
The relationship between the throttle lever angle and the throttle
resolver angle is linear and:1 deg TLA = 1.9 deg TRA.
CONTROL LEVER
REVERSER LATCHING
LEVER
AUTOTHRUST INSTINCTIVE
DISCONNECT PUSHBUTTON
ENGINE POWER SETTING
Power Management
Throttle Control Lever Mechanism
The throttle control mechanism for one engine is shown
below.
The control system consists of:•
•
•
the throttle control lever
the mechanical box
the throttle control unit
The throttle control lever movement is transmitted through
a rod to the mechanical box. The mechanical box
incorporates 'soft' detents which provides selected engine
ratings, it also provides "artificial feel" for the throttle
control system.
The output from the mechanical box is transmitted by a
second rod to the throttle control unit. The throttle control
unit incorporates two resolvers and six potentiometers.
Each resolver is dedicated to one E.E.C. channel, the
output from the potentiometers provides T.L.A. signals to
the aircraft flight management computers.
A rig pin position is provided on the throttle control unit for
rigging the resolvers and potentiometers.
Note
The E.E.C. incorporates resolver fault accommodation
logic which allows engine operation after a failure/loss of
T.R.A. signal.
Bump Rating Push Button
In some cases (optional) the throttle control levers are
provided with "Bump" rating push buttons, one per engine.
This enables the E.E.C. to be re-rated to provide
additional thrust capability for use during specific aircraft
operations.
REVERSE LATCHING AUTOTHRUST INSTINCTIVE
LEVER
DISCONNECT PUSHBUTTON
CONTROL LEVER
REVERSER LATCHING
LEVER
AUTOTHRUST
INSTINCTIVE
DISCONNECT
PUSHBUTTON
THROTTLE CONTROL
LEVER
MECHANICAL BOX
BUMP
PUSHBUTTONS
ROD
THROTTLE CONTROL
UNIT
ENGINE CONTROL-THROTTLE MECHANISM
Power Management
Throttle Control Lever Mechanism
The throttle control lever moves over a range of 65
degrees, from minus 20 degrees to plus 45 degrees.
An intermediate retractable mechanical stop is provided at
0 degrees.
Forward Thrust Range
The forward thrust range is from (0 to plus 45) degrees.
0 deg. = forward idle power
45 deg. = rated take off power
Two detents are provided in this range:at 25 degrees: max climb (MCLB)
at 35 degrees: max continuous (MCT)/flexible
(deated) take off power (FLTO)
Reverse Thrust Range
Lifting the reverse latching lever allows the throttle to
operate in the range 0 degrees to minus 20 degrees.
Two detents are provided in this range:at minus 6 degrees: reverse idle power, corresponds
to thrust reverse deploy
commanded
at minus 20 degrees: max reverse power.
Auto Thrust System (A.T.S.)
The Auto Thrust System can only be engaged between
0 degrees and plus 35 degrees.
ENGINE CONTROL - THROTTLE SETTING
Power Management
EEC-Fuel System Interface
Movement of the pilots throttle control lever is sensed by
the dual resolvers which signal the TRA to the EEC.
The EEC computes the fuel flow which will produce the
required thrust.
The computed fuel flow request is converted to an
electrical current (I) which drives the torque motor in the
Fuel Metering Unit (FMU) which modulates fuel servo
pressure to move the Fuel Metering Valve (FMV) and sets
the fuel flow.
Movement of the FMV is sensed by a dual resolver which
is located in the fuel metering unit next to the FMV.
The dual resolver translates the fuel metering valve
movement into an electrical feedback signal which is fed
back to the EEC.
The basic control loop is shown below.
POWER SETTING - BASIC CONTROL LOOP
Power Management
Basic Control Loop
NOTE
The EEC uses closed loop control based on Engine
Pressure Ratio (EPR) or N1 if EPR is unobtainable.
EEC controls the rate of change of fuel flow, and thus
acceleration/deceleration times, as a function of the rate of
change of N2.
EPR Closed Loop Control
The E.E.C. computes a Target EPR as a function of:T.R.A.
Tamb
T2
Alt
Mn
: (Throttle Resolver Angle)
: (Ambient temperature)
: (Engine air inlet temperature)
: (Altitude)
: (Mach Number)
The EPR target is compared to the actual EPR to
determine the EPR error.
The EPR error is converted to a rate controlled fuel flow
command (WF) which is summed with the measured fuel
flow (WF actual) to produce the WF error.
The W.F. error is converted to a current I which is sent to
the FMU to drive the torque motor, this moves the FMV to
change the fuel flow. The change in fuel flow causes the
engine to accelerate/decelerate and bring about a change
in actual EPR.
This process continues until there is no EPR error.
Nl Reversion
Loss of the P2 or P4.9 signal will cause an automatic
reversion to Rated N1 closed loop control.
If the T2 signal also fails this causes reversion to Unrated
N1 closed loop control.
NOTE:
•
•
N1 Rated control is dispatchable
N1 Unrated control is non dispatchable
POWER SETTING -EPR/N1 CLOSED LOOP CONTROL
Power Management
Thrust Modes
The engine operates in one of three thrust modes, AUTO,
MEMO and MANUAL. Entering, exiting these three modes
is controlled by inputs to the Engine Interface Unit (EIU)
In the Memo mode the thrust is frozen, to the last actual
EPR value, and will remain frozen until the throttle lever is
moved manually, or, auto thrust is reset.
Auto Thrust Mode
Manual Thrust Mode
The Auto Thrust mode is only available between idle and
MCT when the aircraft is in flight.
This mode is entered anytime the conditions for AUTO or
MEMO are not present. In this mode thrust is a function of
throttle lever position.
After take off the throttle is pulled back to the max climb
position, the auto-thrust system will be active and the
Automatic Flight system will provide an E.P.R. target to
provide either:•
•
•
•
max clistib thrust
an optimum thrust
a minimum thrust
an aircraft speed (mach number)
in association with the auto pilot.
Memo Mode
The memo mode is entered automatically, from Auto
mode if:• the EPR target is invalid
• one of the instinctive disconnect buttons on the
throttle is activated
• Auto thrust is disconnected by the E.I.U
Alpha Floor Protection
If an aircraft stall is imminent the Auto Thrust System sets
the engine power to T/O regardless of throttle position.
POWER MANAGEMENT -THRUST MODE SETTING
PART ONE - SECTION 5
FUEL SYSTEM
Fuel System
Introduction
The primary purpose of the fuel system is to provide a
completely controlled continuous fuel supply in a form
suitable for combustion, to the combustion system.
Control of the fuel supply is by the EEC via the FMU. High
pressure fuel is also used to provide servo pressure
(actuator muscle) for some actuators.
The major components of the fuel system
include:
•
•
•
•
•
•
•
high and low pressure fuel pumps (dual unit)
fuel-oil heat exchanger
LP fuel filter
fuel metering unit (FMU)
fuel distribution valve
fuel injectors (20)
fuel diverter and back to tank valve (FDRV)
The fuel system is shown schematically below.
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
LP fuel pump.
LP fuel filter by-pass valve.
Differential pressure switch.
LP fuel temperature probe.
HP fuel pressure relief valve.
HP fuel pump.
Fuel metering valve (FMV).
Overspeed valve (OS)
Pressure raising and shut-off valve (PRSOV).
Pressure drop governor and spill valve
ACC actuator.
ACOC air control valve actuator.
VSV system actuator.
BSBV - Master actuator.
BSBV - Slave actuator.
V2500 FUEL SYSTEM - SIMPLIFIED OVERVIEW
Fuel System Components Description
Fuel Pumps
H.P.STAGE
The fuel pump unit, shown below, consists of low pressure
and high pressure stages which are driven from a
common gearbox output shaft.
Purpose.
LP Stage
Purpose
To provide the necessary pressure increase to:
•
•
•
Type
account for pressure losses through the FCOC and
the LP fuel filter.
suppress cavitation.
maintain adequate pressure at the inlet to the HP
stage
Shrouded, radial flow, centrifugal impeller, with an axial
inducer.
To inciease the fuel pressure to that which will ensure
adequate fuel flow and good atomisation at all engine
operating conditions.
Type
Two gear (spur gear) pump.
Features
• provides mounting for fuel metering unit (F.M.U.)
• integral relief valve
UNDERSIDE VIEW
FORWARD
1 LP pump discharge port
2 LP pump inlet port
3 LP return from FMU
4 HP pump discharge port
5 HP pump inlet port
6 Adapterhousing
7 Accessory drive clamp
LP/HP FUEL PUMPS UNIT
Fuel System Component
Fuel Cooled Oil Cooler
Purpose
To transfer heat from the oil system to the fuel system to:•
•
reduce the temperature of the engine lubricating oil
prevent fuel ing
Type
Single fuel pass/multi pass oil flow.
Features
•
•
a single casing houses the FCOC and the LP fuel
filter.
provides location for the fuel diverter and back to
tank valve (unit not shown).
•
fuel differential pressure switch.
•
fuel temperature thermocouple.
1
2
3
4
5
6
7
8
9
10
11
:Fuel inlet connection
:Oil outlet connection
:Oil let connection in
:Fuel/oil inlet heat exchanger assembly
:Fuel inlet connection from FD and RV
:Fuel fitter inlet connection from FD and RV
:Fuel filter by-pass valve
:Fuel filter pressure plate
:Fuel drain plug
:Fuel outlet connection
:Fuel filter element
FUEL OIL HEAT EXCHANGER
Fuel System Component
Low Pressure Fuel Filter
Purpose
To remove solid contaminants from the fuel.
TYPE
Woven, glass fibre, disposable, 40 micron (nominal)
Features
•
a differential pressure switch, which generates a flight
deck message "Fuel Filter Clog" if the differential
pressure across the filter reaches 5 psid.
•
a by-pass valve which opens and allows fuel to bypass the filter if the differential pressure reaches 15
psid.
•
a fuel drain plug, used to drain filter case or to obtain
fuel samples.
FUEL COOLED OIL
COOLER
END SEAL
1
2
3
4
5
6
7
FUEL PLUG
LP FUEL FILTER
Fuel Drain plug
O-ring
Bolt(5 off)
Washer (5 off)
O-ring
Fuel filter element
cap assembly
Fuel System Component
Fuel Metering Unit (F.M.U.)
The FMU is the interface between the EEC and the fuel
system.
It is located on the dual fuel pumps unit, on the rear of the
main gearbox, and is retained by four bolts as shown
below.
All the fuel delivered by the HP fuel pumps which is much
more than the engine requires passes to the FMU.
The FMU, under the control of the EEC, meters the fuel
supply to the fuel spray nozzles, it also supplies HP fuel
for the operation (muscle) of a number of actuators. Any
fuel supplied by the HP pumps which is not needed for
these two uses is returned, from the FMU to the LP side of
the fuel system.
In addition to the fuel metering function the FMU also
houses the Overspeed Valve and the Pressure Raising
and Shut Off Valve. The Overspeed valve under the
control of the EEC, provides overspeed protection for the
LP (N1) and HP (N2) rotors.
The Pressure Raising and Shut Off Valve provides a
means of isolating the fuel supplies to start and stop the
engine.
Note:
There are no mechanical inputs to, or outputs from, the
FMU.
7
1
1
2
3
4
5
6
7
8
9
IP/HP fuef pumps unit
Fuel metering unit-FMU
LP pump in let
LP pump outlet
HP pump inlet
HP pump outlet-to FMU inlet
FMU spill-to LP
FMU outlet to ffowmeter
EEC-Channel (A) Electrical connectors
10 EEC-Channel (B) Electrical connectors
11 PRSOV Electrical connectors
FUEL METERING UNIT DISCONNECT POINTS
Fuel System Component
Fuel Metering Unit
A simplified schematic representation of the Fuel Metering
Unit is shown below.
The three main functions of the FMU are:
1. metering the fuel supplies to the fuel spray nozzles.
2. overspeed protection for both the LP (N1) and HP
(N2) rotors.
3. isolation of fuel supplies for starting / stopping the
engine.
These three functions are carried out by three valves,
arranged in series, as shown:
•
•
•
the Fuel Metering Valve (FMV).
the Overspeed Valve (OS).
the Pressure Raising and Shut Off Valve (PRSOV).
The position of each valve is monitored and positional
information is transmitted back to the EEC. This ensures
that the EEC always knows that the valves are in the
commanded position.
TO FLIGHT DECK
EEC
TM
TM
FEED TO EXT
SERVICES
HP
FUEL
PUMP
SERVO
PRESSURE
REGULATOR
POSITION
RESOLVER
FUEL
METERING
VALVE
SPEED
VALVE
TM
M.SW
M.SW
PRESSURE
RAISING
AND
SHUT-OFF
VALVE
PRESSURE DROP
GOVERNOR AND
SPILL VALVE
TM = TORQUE MOTOR
M.SW = MICROSWiTCH
FUEL METERING UNIT SCHEMATIC
TO FLOWMETER
Fuel System Component
Fuel Metering Unit - operation
The three major functions:
• fuel metering
• overspeed protection
• pressure raising and shut off
are explained in more detail in the following pages.
CHANNEL B
CHANNEL A
TO FUEL
DISTRIBUTION
VALVE
HP FUEL
FROM PUMP
HP FEED TO EXT
SERVOES
FMU SPILL
AND SERVO
RETURN
TO LP PUMP
INLET
FUEL METERING UNIT
1 :SPILL PRESSURE RAISING VALVE
2 :PRESSURE DROPREGULATOR AND
SPILL VALVE
3 :FLOW WASH FILTER
4 :SERVO PRESSURE REGULATOR
5 :CONNECTORS
6 :MV TORQUE MOTOR DENSITY ADJUST
7 :POSITION RESOLVER
8 :MICRO SWITCH
9 :DRY DRAINS
10 :LATCHING SHUT-OFF MOTOR
11 :HEAT SHIELD AND EMI COVER
12 :MICRO SWITCH
13 :DRY DRAINS
14 :PRESSURE RAISING AND
SHUT OFF VALVE
15 :OVERSPEED VALVE
16 :METERING VALVE
17 :MV BY-PASS ADJUSTER
18 :MIN FLOW ADJUSTER
19 :SERVO SWITCHING VALVE
Fuel Metering Operation
Fuel metering is achieved by the Fuel Metering Valve and
the Pressure Drop Regulator and Spill Valve, which act
together in the following sequence.
Signals from the EEC, cause the torque motor to change
position which directs fuel servo pressure to re-position
the Fuel Metering Valve. This changes the size of the
metering orifice through which the fuel passes which in
turn changes the pressure drop across the metering valve.
The change in the pressure drop is sensed by the
Pressure Drop Regulator which will re-position the spill
valve and so increase/decrease the fuel flow through the
fuel metering valve until the pressure drop is restored to its
datum value.
The increase/decrease in fuel flow causes the engine to
accelerate/decelerate until the actual EPR is that
demanded by the EEC signal.
Movement of the Fuel Metering Valve is transmitted
through a rack and pinion mechanism to drive a dual
output position resolver, the resolver output is fed back to
the EEC.
The fuel metering valve incorporates a manual adjuster to
allow setting for changes in fuel density. Automatic
compensation for changes in fuel temperature is provided
by bi-metallic washers located in the pressure drop
governor and spill valve assembly.
CHANNEL CHANNEL
B
FUEL METERING
VALVE-POSITION
FEED-BACK SIGNAL
TO EEC
TORQUE MOTOR
FUEL DENSITY ADJUSTER
SERVO PRESSURE
REGULATOR
POSITION RESOLVER
HP FUEL
FROM
PUMP
FLOW WASH
FILTER
LP SPILL
RETURN
MIN FLOW STOP
PRESSURE DROP
REGULATOR AND SPILL
VALVE
FMU FUEL METERING VALVE
PRESSURE DROP
REGULATOR
AND SPILL VALVE
Fuel Metering Unit
Overspeed Protection
The Overspeed Valve is positioned down stream, in
series, with the Fuel Metering Valve.
Note:
It should be understood that this device is not incorporated
to provide the usual N1/N2 red line limiting of max T.o
speed of 100% The EEC will act through the Fuel
metering Valve to trim the fuel flow if Nl or N2 reach 100%.
Operation
The overspeed valve is spring loaded to the closed
position, it is opened by increasing fuel pressure during
engine start and during normal engine operation is always
fully open.
In the event of an overspeed the EEC sends a signal to
the overspeed valve torque motor which changes position
and directs H.P. fuel to the top of the overspeed valve this
fully closing the valve.
A small by-pass flow is arranged around the overspeed
valve to prevent engine flame out
The overspeed valve is hydraulically latched in the closed
position, thus preventing the engine from being reaccelerated - the recommended procedure is for the flight
crew to close the engine down, and not re-start.
Closing down the engine is the only way to release the
hydraulic latching.
Note:
Because the overspeed valve is spring loaded to the
closed position, and opened by fuel pressure, the
overspeed valve will close on every engine shut down.
CHANNELA
CHANNEL B
DE-ENERGIZED
TORQUE
MOTOR
ENERGIZED
TO PRESSURE
RAISING VALVE
OVERSPEED
VALVE
METERING
VALVE
MIN FLOW
ORIFICE
PRESSURE DROP REGULATOR
AND SPILL VALVE
FMU OVERSPEED PROTECTION
Fuel Metering Unit
Pressure Raising and Shut off Valve
The third valve in the F..M.U. is the Pressure Raising and
Shut off Valve (PRSOV). Its primary function is to Isolate
the fuel supplies to the fuel spray nozzles for starting and
stopping the engine, it also acts as a pressure raising
valve to ensure that, during engine starts, fuel is not
passed to the fuel spray nozzles until fuel pressures in the
F.H.U. are high enough to ensure the control devices will
function correctly.
The two position torque motor which controls H.P. fuel
pressure to operate the PRSOV also controls a spill valve
servo line. When the torque motor is selected to close the
PRSOV, to shut down the engine, the spill valve servo line
is opened, this will fully open the spill valve and direct all
the H.P. fuel pump delivery back to the L.P. fuel system.
The PRSOV torque motor is commanded open by the
EEC during AUTO starts. It is commanded open by a
cockpit mounted switch during MANUAL starts.
The PRSOV can be commanded closed by the EEC
during AUTO start sequences if the sequence has to be
stopped for any reason. The EEC's ability to close the shut
off valve is inhibited above 50% N2. Above 50% N2, and
in flight, the PRSOV can only be closed by the crew
operated switch in the cockpit.
LATCHING SHUT-OFF MOTOR
{Shown in open position)
CH A CH B
MICROSWITCH
COCKPIT
SHUT-OFF
SHUT
PX
FIXED ORIFICE
SPILL VALVE
SERVO
LP
PRV AND SHUT-OFF PISTON
(Shown in shut-off position)
FMU PRESSURE RAISING AND SHUT-OFF VALVE
Component Description
Fuel Distributor Valve
Purpose
To accept the fuel supplied from the FMU and apportion
it to the ten fuel manifold tubes.
The ten fuel manifolds each supply two fuel spray nozzles.
Location
At the 4 o'clock position on the front flange of the diffuser
casing.
Features
•
•
•
integral fuel filter with by-pass valve
single fuel metering (check) valve
o spring closed
o fuel pressure opened
ten fuel outlet ports
REAR OUTER CASING
FRONT FLANGE
BRACKET
REAR MOUNT
BRACKET
FUEL SUPPLY
MANIFOLD
SHUT OFF SEAL
FUEL INLET
FUEL DISTRIBUTION
VALVE
DIFFUSER CASE
FRONT
FLANGE
FILTER
(With bypass)
TRANSFER TUBE
FUEL SUPPLY TUBES
FUEL DISTRIBUTION VALVE
DRAIN LINE
(8 places)
FUEL OUT TO MANIFOLD (10 places)
Component Description
Fuel Distributor Valve (Cont)
The diagram below shows an 'exploded' view of the fuel
distributor valve. It also shows how access is gained to the
filter for the periodic inspection which is called up in the
Aircraft Maintenance Manual.
1
1
2
3
4
5
2
3
: FUEL DISTRIBUTION VALVE
: PACKING
: FILTER
: PACKING
: FILTER COVER
4
5
6
7
8
6 : PACKING
7 : WASHER
8 : BOLT
9 : BOLT
10 : FUEL INLET LINE BOLT
FUEL FILTER - FUEL DISTRIBUTION VALVE
9
10
Component Description
Fuel Manifold
To conduct fuel from the fuel distribution valve to the fuel
spray nozzles.
Location
Mounted around the diffuser casing.
Type
Single pipe with twin branch.
Features
•
•
10 manifolds
each manifold supplies two fuel spray nozzles
10 NOZZLE CLUSTERS
(2 nozzles each)
20 FUEL NOZZLES
IGNITER PLUGS
FUEL IN FROM FMU
FUEL DISTRIBUTION VALVE
VIEWED FROM REAR
FUEL DISTRIBUTION TUBES
Component Description
Fuel Spray Nozzles
Purpose
To inject the fuel into the combustion chamber in a form
suitable for combustion by:• atomizing the fuel
• mixing it with air
• controlling the spray pattern
Location
Mounted around the diffuser casing.
Type
Single orifice - air blast.
Features
• 20 fuel spray nozzles
• inlet fitting houses fuel filter
• internal and external heat shields to reduce coking.
FUEL FLOW
FUEL INLET
CONNECTOR
CONNECTOR
MOUNTING
FLANGE
FUEL STRAINER ELEMENT
INSERT
METERING
PLUG
FAIRING
(Both sides)
SUPPORT
FAIRING
FUEL-
FUEL-
AIR FLOW
FAIRING
(Both sides)
NOZZLE
HEAT SHIELD
FAIRING
BAFFLE
FUEL SPRAY NOZZLES (20)
NOZZLE
Fuel Spray Nozzles
Removal/Installation
The installation details for one of the fuel spray nozzles is
shown below, the other 19 are similar but there are slight
differences. Reference must be made to the Engine
Maintenance Manual CH 73-13-41.
The following points must be observed:• all the gaskets and packing used must be discarded
on removal and replaced with new items on re• lubricate the bolt threads with engine oil
• observe the torque loading limitations quoted in the
Engine Maintenance Manual.
• apply anti-galling compound (V10-032) to the
shoulders of the end fittings of the fuel supply tubes
on re-assembly.
• apply white petrolatum (V10-041) to the rubber
packings on re-assembly.
Note
Reference numbers - e.g. V10-032 refer to consumable
materials - a full list of these can be found in the Engine
Maintenance Manual CH 70-30-00.
DIFFUSER CASE
BOSS
FLANGE K
FORWARD
GASKE
FUEL NOZZLE
PACKINGS
BRACKET
FUEL TRANSFER
TUBE
BOLTS
GASKET
NUT
FUEL SUPPLY
TUBE
BOLT
CLAMP
FUEL SPRAY NOZZLE REMOVAL/INSTALLATION
Electrical Harness
Fuel Control (F.M.U.)
The fuel control electrical harness connections are shown
schematically below.
401VC
409VC
IDG POWER
CABLE
454VC
450VC
452VC
407VC
403VC
457VC
406VC
405VC
408VC
448VC
402VC
404VC
447VC
FUEL METERING
UNIT
FUEL METERING UNIT
1 Fuel metering valve - T/M
2 Fuel metering valve - resolver
3 Overspeed valve - T/M
4 Overspeed valve - M/Sw
5 PRSOV-T/M
6 PRSOV- M/Sw
FUEL SYSTEM CONTROL HARNESS
Fuel System
Electrical Harness,
The fuel system electrical harness connections are
as shown below.
407VC
403VC
401VC
409VC
IDG POWER
CABLE
454VC
450VC
452VC
D600
1
1
2
3
4
457VC
406VC
405VC
408VC
3
4
2
Fuel Flowmeter
Fuel Filter ∆P S/W
Fuel Temperature/ TC
LP Fuel Filter
FUEL SYSTEM HARNESS
448VC
402VC
404VC
447VC
PART ONE - SECTION 6
OIL SYSTEM
OIL SYSTEM-INTRODUCTION
The oil system is a self contained, full flow recirculating
type design to ensure reliable lubrication and cooling
under all circumstances.
Oil cooling is controlled by a dedicated Heat Management
System which ensure that engine oil, I.D.G. oil and fuel
temperatures are maintained at acceptable levels while
ensuring the optimum cooling configuration for the best
engine performance.
A simplified system diagram is shown below.
SYSTEM MONITORING
The operation of the engine oil system may be monitored
by the following flight deck indications:
• engine oil pressure
• engine oil temperature
• oil tank contents
In addition warnings may be given for the following nonnormal conditions:
• low oil pressure
• scavenge filter clogged or partly clogged (high
differential pressure)
• No. 4 compartment scavenge valve inoperative
V2500 SIMPLIFIED OIL SYSTEM
Oil System
No 1, 2 and 3 Bearing Lubrication
The oil feeds, and scavenge return, for No's 1, 2 and 3
bearings, including oil damping supplies for No 3 bearing
are shown in detail below.
RESTRICTOR
OIL
DlSTRIBUTORS
STRAINERS
STRAINER
No 1 BEARING (LP)
No 3 BEARING
- DAMPER
JET TO No
2 BEARING
SEAL
JET TO No 1
BEARING
JET TO
HYDRAULIC SEAL
JET TO No 3 BEARING
CARBON SEAL
JET TO BEVEL
GEAR MESH
SEAL DRAINS
ET TO BEVEL GEAR
BEARINGS AND STEADY
BEARING
STEADY
BEARING
No 1 2 3 BEARING LUBRICATION
Oil system
No 4 Bearing Lubrication
The oil feed, and scavenge, arrangements for No 4
bearing are shown in detail below.
STRAINER
THERMALLY
INSULATED
SHIELDING
HP COMPRESSOR DELIVERY
AIR VIA AIR COOLED AIR COOLER
(ACAC) TO No 4 BEARING
FAN AIR USED AS COOLANT
JET TO No 4 BEARING
AND REAR CARBON
SEAL
JET TO FORWARD
CARBON SEAL
No 4 BEARING BUFFER AIR FEED
(Cooling air)
SCAVENGE OIL
SCOOP AND STRAINER
No 4 BEARING LUBRICATION
Oil System
No 5 Bearing Lubrication
The oil feed and scavenge arrangements for No 5 bearing
are shown in detail below.
'
LAST CHANCE
STRAINER
5 BEARING FEED
RESTRICTOR
SQUEEZE FILM JET
JFT TO No 5 BEARING
CARBON SEAL
STRAINER
OIL SCALLOPS
INSULATION BLANKET
EXHAUST CASING
No 5 BEARING LUBRICATION
Oil System
Gearbox Lubrication
The gearboxes, main and angle, are lubricated and
scavenged as shown below.
ANGLE
GEARBOX
MAGNETIC
CHIP DETECTOR
STRAINER
SCAVENGE PUMPS
STARTER
RIGHT HAND
SCAVENGE
PICK-UP
LEFT HAND
GEARBOX SCAVENGE
PICK-UP
DEDICATED
GENERATOR
HYDRAULIC
PUMP
PRESSURE PUMP
AND FILTER
FUEL PUMPS
GEARBOX LUBRICATION
Component Description
Oil Tank
Purpose
To store the oil supply.
Location
Located to the top L.H. side of the external gearbox.
Type
Pressurised, hot tank.
Features
•
•
•
•
•
•
oil system servicing:
o gravity fill port
o prismalite oil level indicator
o tank capacity, 28.3 US quarts usable oil 24 US
quarts.
Internal cyclone type de-aerator
Tank pressurization valve (6 psi) ensures adequate
pressure at inlet to oil pressure pump.
Strainer in tank outlet to pressure pump.
Provides mounting for scavenge filter on rear face.
Provides mounting for scavenge filter and master
magnetic chip detector (MMCD).
OIL TANK PRESSURIZATION VALVE
OIL TANK BREATHER TUBE
PRISMALITE OIL
LEVEL INDICATOR
OIL TANK
CONTENTSTRANSMITTER
BREATHER TUBE
RETAINING FLANGE
MASTER MAGNETIC
CHIP DETECTOR
MANUAL OIL
FILLER
SCUPPER
DRAIN
OIL TANK
OIL SCAVENGE FILTER ASSY
OIL TANK
Component Description
Pressure Pump
Pressure Filter
Purpose
Purpose
To supply oil under pressure to the engine bearings,
gearbox drive, and accessory drives.
To trap solid contaminants.
Type
Mesh type filter - cleanable - nominal 125 micron rating.
Standard gear type (speed is 21.4% N2).
Type
Location
Location
On the pressure pump housing.
Attached to front face of external gearbox on L.H. side just
below oil tank. Oil from the tank is taken to the pressure
pump by a passage cast within the gearbox.
Features
Features
• provides the housing for the pressure
• filter
• cold start pressure limiting valve
• flow trimming valve
•
•
pressure priming connection
anti-drain valve
FLOW TRIMMING VALVE
(for adjusting oil fow to a standard
level during pass-off testing)
ANTI-DRAIN VALVE
(prevents oil loss when
fitter is removed)
COLD START PRESSURE
LIMITING VALVE
(opens at 450 psi)
PRESSURE PUMP '
(Driven from gearbox)
FILTER
ELEMENT
(125 microns (ju)
filtration)
OIL PRESSURE PUMP
STRAINER
OIL PRESSURE PUMP FILTER UNIT
OIL TANK
Air/Oil Heat Exchanger (ACOC)
Purpose
To reject excess heat from the oil system to the fully
modulated cooling (fan) air flow, as required by the E.E.C.
Type
Corrugated fin and tube - double pass.
Location
Attached to the fan casing on the lower R.H. side.
Features
•
•
oil by-pass valve
modulated air flow as commanded by E.E.C. (heat
management system) Air flow regulated by air
control valve.
AIR CONTROL
VALE
OIL IN
OIL OUT
AIR INLET
DUCT
Air/Oil Heat Exchanger
Oil System
Air/Oil Beat Exchanger
The cooling air and the oil flows through the air/oil heat
exchanger are shown below.
MODULATING VALVE
(fuel powered and
controlled by signal
from engine EEC
electronic control unit)
NACELLE OUTLET
PANEL
AIR FLOW
OIL FLOW
Full oil flow at
all times, heat
rejection controlled
by varying
air flow
BYPASS VALVE
(used in event of cooler
blockage - opens at
50 psi)
AIR AND OIL FLOWS
REAR OF FAN CASE
Component Description
Air/Oil Heat Exchanger
Air Modulating Valve
Purpose
To govern the flow of cooling (fan) air through the air/oil
heat exchanger, as commanded by the Heat
Management Control System. (EEC)
Type
Plate type supported at either end by stubshafts.
Operated by an Electro-Hydraulic Servo Valve
mechanism.
Location
Bolted to the outlet face of the air/oil heat exchanger.
Features
•
•
•
fails safe - valve fully open - maximum cooling
position
fire seal forms an air tight seal between the
unit outlet and the cowling orifices.
control by either channel A or B of E.E.C.
•
•
•
valve position feed back signal via L.V.D.T.
to each channel of E.E.C.
valve positioned by fuel servo pressure acting on a
control piston
fuel servo pressure directed by the ElectroHydraulic
Servo
Valve
assembly
which
incorporates a Torque motor.
TORSION SPRING
FIRE SEAL
VALVE HOUSING
FUEL (SERVO)
RETURN
AIR VALVE
STOP
FUEL (SERVO)
SUPPLY
STOP PLATE
VIEW ON A
LINEAR VARIABLE
DIFFERENTIAL
TRANSFORMER
ASSEMBLY
(LVDT)
ELECTRO-HYDRAULIC
SERVO VALVE
ASSEMBLY
(EHSV)
CHANNEL A
ELECTRICAL
RECEPTACLES
CHANNEL 8
PUSH ROD
ACTUATING LEVER
AIR MODULATING VALVE
Component Description
Air Control Valve
Electro Hydraulic Servo Valve (E-H.S.V.)
Purpose
It provides the muscle to move the air control valve to the
EEC commanded position.
Type
two stage directional flow valve
• stage 1: electrically activated torque motor and Jet
pipe
•
stage 2: spool valve
Location
Bolted to the air control valve casing.
Features
•
•
•
•
•
two independent torque motor windings -one
connected to each channel of EEC.
operation, from either channel of EEC.
jet pipe protected by 90 micron filter.
E.H.S.V. biased to ensure air control valve fully
open at engine start condition.
single fuel servo supply from FMU.
CHANNEL A
FUEL SUPPLY
FUEL RETURN
DRAIN
1 Torque motor
2 Torque motor jet pipe
3 Spool vaive
4 Actuating piston
5 Air valve
6 Spring
7 Push rod and lever
8 LVDT
ACOC ELECTRO-HYDRAULIC SERVO VALVE
Component Description
Fuel/Oil Heat Exchanger
Purpose
•
•
cool the engine oil
heat the fuel
Type
Single pass fuel flow multi pass oil flow. Forms an integral
unit with the LP fuel filter.
Location
Bolted to the fan casing - LH side - on the engine centre
line.
Features
Differential pressure relief valve permits oil by-pass if oil is
congealed or cooler blocked
Anti-syphon drilling
Mttrix assy
tubes & paffle p|ate
Fuel filter
Fuel in from
LP pump
Fuel out to
FMU & HP pump
Pressure relief valve
opens at 50 psi
Bypass setction of
cooler when mattrix
is blocked or when
oil is congeealed.
FUEL-OIL HEAT EXCHANGER
Component Description
Scavenge Pumps Unit
Purpose
Returns scavenge oil to tank.
Type
Standard gear type pumps (6). All pumps rotate at the
same speed (22% N2), pump capacity is determined by
the width of the gears.
Location
All 6 scavenge pumps are housed together as a single unit
on rear L.H, side of the gearbox.
Features
combines the flows from all the pump outlets to return to
oil tank
DEVELOPED COMPOSITE
SECTION THROUGH PUMPS
SHOWING FUNCTION
OIL FILTER
FUEL PUMP DRIVE
PAD
GEAR PUMPS
OIL SCAVENGE PUMP
ANGLE GEARBOX
1
STRAINER
6
4
2
3
5
LEFT HAND
GEARBOX
PICKUP
MAGNETIC
CHIP
DETECTOR
STRAINER
SCAVENGE PUMPS UNIT
RH HAND
GEARBOX
PICK UP
Component Description
De-oiler
Purpose
•
•
To separate the breather air/oil mixture, it return the
oil to the oil scavenge system via its own scavenge
pump, and
vent the air overboard through the R.H. fan cowl.
Type
Centrifugal separator.
Location
Bolted to the front face (R.H. side) of the external gearbox.
Features
•
provides mounting for the No 4 bearing chamber
scavenge valve.
•
provides location for No 4 bearing Magnetic Chip
Detector housing.
DEOILER
OIL TANK
OIL PRESSURE PUMP
DEOILER
Component Description
Scavenge Filter
Purpose
To trap solid contaminants.
Type
Disposable.
Location
Mounted to the rear of the oil tank -shown below.
Features
•
•
•
•
disposable.
by-pass valve.
∆P switch connections.
provides housing for master magnetic chip detector.
OIL FILTER
PRESURE RELIEF VALVE
(Opens at 20 psi
pressure drop)
FUEL PUMP
DRIVE PAD
OIL SCAVENGE PUMP
ANGLE GEARBOX
MASTER MAGNETIC
CHIP DETECTOR
(Vatric type)
SCAVENGE FILTER
PRESSURE DROP
WARNING SWITCH
(∆p 12 psi
pressure setting)
FILTER ELEMENT
(30 microns
filtration fine)
OIL TEMPERATURE SENSOR
(Weston instruments)
VIEW ON A
OIL SCAVENGE FILTER
Component Description
No 4 Bearing Scavenge Valve
Purpose
Maintains No 4 bearing compartment seal deferential
air/oil mixture to the de-oiler.
Type
Pneumatically operated, two position valve;
• fully open at low speeds.
• closed to minimum area at high engine speeds.
Location
Mounts on the front face of the de-oiler casing.
Features
•
•
•
position feed back signal to EEC.
uses stage 10 air as servo air.
uses value of pressure of stage 10 air as operating
parameter:
o stage 10 air < 150 psi: valve max open position
o stage 10 air > 200 psi: valve min open position
10TH
STAGE
MAGNETS
SCAVENGE
FLOW
REED
SWITCH
4
5
Bolt (3 off)
Sealing ring
DE-OILER CASE
No. 4 BEARING SCAVENGE VALVE
ELECTRICAL
CONNECTOR
Component Description
Magnetic Chip Detectors (M.C.D.)
A total of 7 MCD's are used in the oil scavenge system.
Each bearing compartment and gearbox has its own
dedicated M.C.D. (two in the case of the main gearbox)
although that for the No 4 bearing is located in the De-oiler
scavenge outlet.
Access to the dedicated MCD's is by opening the L and R
hand fan cowls.
The Master M.C.D
It is located in the combined scavenge return line, on the
scavenge filter housing.
The Master MCD is accessible through an access panel in
the LH fan cowl.
If the master MCD indicates a problem then each of the
other MCD's is inspected to indicate the source of the
problem.
FUEL METERING
UNIT
HP/LP FUEL
PUMPS
OIL SCAVENGE
PUMPS
Nos 1, 2 AND 3 BEARING
SCAVENGE CHIP
ANGLED GEARBOX
SCAVENGE CHIP
DETECTOR
No 5 BEARING
SCAVENGE CHIP
DETECTOR
OIL SCAVENGE
FILTER
MASTER CHIP
DETECTOR
INTEGRATED DRIVE
GENERATOR
DE-OILER
OIL
TANK
DE-OILER CHIP
DETECTOR
FORWARD
OIL PUMP AND
FILTER
GEARBOX
CHIP DETECTORS
STARTER
DETECTORS
HYDRAULIC
CHIP DETECTORS - LOCATION
2 POSITION OIL
SCAVENGE VALVE
Component Description
Master Magnetic Chip Detectors
The master magnetic chip detector is located in the
scavenge filter case.
The master MCD samples the combined scavenge return
oil flow.
Access to the master MCD is through a dedicated access
panel in the LH fan cowl.
1 Magnetic probe
2 Seal ring (2 off)
3 Detector housing
MASTER CHIP
DETECTOR
HOUSING
PROBE
(With 'O' ring missing)
SPRING SAFETY PIN
View showing pin against
groove when 'O' ring missing)
MASTER MAGNETIC CHIP DETECTOR
Component Description
Magnetic Chip Detectors
Location
The MCDs for:
•
•
•
No's 1, 2 and 3 bearings
main gearbox - L.H. scavenge pick up
angle gearbox
are located to the rear of the main gearbox on the LH side
as shown below.
Nos 1, 2 & 3 BEARltfG
SCAVENGE CHIP
DETECTORS
ANGLED GEARBOX
SCAVENGE CHIP
DETECTOR
GEARBOX LH
CHIP DETECTOR
LOCATION MAGNETIC CHIP
DETECTORS LEFT-HAND
Magnetic Chip Detector
Location
The MCDs for:
•
•
•
No 5 bearing
De-oiler (No 4 bearing)
Main gearbox - R.H. scavenge pick up
are located as shown below.
DE OILER CHIP
DETECTOR
(NO 4 BEARING)
No 5 BEARING
SCAVENGE CHIP
DETECTOR
GEARBOX RH
CHIP DETECTOR
VIEW A
VIEW B
LOCATION MAGNETIC CHIP
DETECTORS RIGHT-HAND
OIL SYSTEM INDICATIONS
The oil system parameters are displayed on the Engine
page on the Lower ECAM screen.
1. OIL TEMPERATURE (deg C)
• Normal - green indication
• 170 deg C or above:
o flashing green indication
• 190 deg C or above:
o steady amber indication
o master caution light
o single chime
o message (upper ECAM) OIL HI TEMP
• Oil low temperature warning:
- Throttle above idle and
- Engine running
o message OIL LO TEMP
o single chime
2. Oil Quantity
•
•
Normal - green
Less than 4 quarts - flashes green
3. Oil Pressure
•
•
o
o
o
o
Normal: green indication
60 psid or below:
flashing red indication
master warning light f
audio warning
message (upper ECAM)
ENG 1(2) OIL LOW PRESS
THROTTLE 1(2) IDLE
4. Scavenge Filter Clog
• If filter ∆p > 12 psi:
o OIL FILTER CLOG message appears on Engine
page.
OIL SYSTEM INDICATIONS
Component Description
Differential Oil Pressure Transmitter &
Low Oil Pressure Warning Switch
Both units are located on the upper LH side of the fan
case.
OIL PRESSURE
TRANSMITTER
LOW OIL PRESSURE
WARNING SWITCH
OIL PRESSURE TRANSMITTER &
LOW OIL PRESSURE WARNING SWITCH
Component Description
Oil System Electrical Harness
The oil system sensors, controls, actuators are connected
by the main Electrical harness to the EEC and ECAM
computer, as shown below.
1 Low oif pressure switch
2 Lube oil - pressure tx
3 Lube oil temp sensor
4 Lube oif qty tx
5 Scav fitter AP sw
6 IDG oil temp (HMS)
7 ACOC air modulating valve
8 No 4 bearing scav valve
9 No 4 bearing compartment pressure tx
10 Lube oil temp sensor - HMS
OIL SYSTEM HARNESS
Engine Oil System
Pressure Filter Removal/Installation
Refer to the AMM Ch 79
Task 79-21-44-000-010
The procedure is summarised below.
Note: the pressure filter can be cleaned ultrasonically.
OIL PRESSURE
PUMP ASSEMBLY
1 Filter casing
2 Lock wire
3 Seal ring
4 Drain plug
5 Bolt
6 Washer
7 Filter cover
8 Seal ring
9 Filter element
OIL PRESSURE FILTER REMOVAL/INSTALLATION
Oil system
Scavenge Filter - R & I
Refer to the AMM Ch 79
Task 79-22-44-000-010
Note
The scavenge filter is not re-usable.
OIL OUTLET
(Return to oil tank)
MASTER MAGNETIC
CHIP DETECTOR
1
OIL INLET
(From
scavenge
pumps)
1 Filter casing
2 Guide pin
3 Lock wire
4 Seal ring
5 Drain plug
6 Bolt
7 Washer
8 Filter cover
9 Seat ring
10 Filter element
OIL SCAVENGE FILTER - REMOVAL/INSTALLATION
PART ONE - SECTION 7
HEAT MANAGEMENT SYSTEM
Beat Management System
Purpose
Mode 1
The system is designed to provide adeguate cooling, to
maintain the critical oil and fuel temperatures within
specified limits, whilst minimising the requirement for fan
air offtake.
This is the normal mode and is shown below. In this mode
all the heat from the engine oil system and the I.D.G. oil
system is absorbed by the L.P. fuel flows. Some of the fuel
is returned to the aircraft tanks where the heat is absorbed
or dissipated within the tank.
Three sources of cooling are available:•
•
•
the LP fuel passing to the engine fuel system
the LP fuel which is returned to the aircraft fuel tanks
fan air
There are four basic configurations between which the
flow paths of fuel in the engine L.P. fuel system are varied.
Within each configuration the cooling capacity may be
varied by control valves which form the Fuel Diverter and
Back to Tank Valve.
The transfer between modes of operation is determined by
software logic contained in the E.E.C. The logic is
generated around the limiting temperatures of the fuel and
oil within the system together with the signal from the
aircraft which permits/inhibits fuel spill to aircraft tanks.
This mode is maintained if the following conditions are
satisfied:•
•
•
engine not at high power setting (T/O and early part
of climb).
cooling spill fuel temperature less than 100 deg C.
fuel temperature at pump inlet less than 54 deg C.
Heat Management System
Mode 3
The second mode shown below is the mode adopted
when the requirements for fuel spill back to tank can no
longer be satisfied i.e.
engine at high power setting.
spill fuel temperature above limits (100 deg C
tank fuel temperature above limits (54 deg C)
In this condition all the heat from the engine and I.D.G. oil
systems is absorbed by the burned fuel. If however, the
fuel flow is too low to provide adequate cooling the engine
oil will be pre-cooled in the air/oil heat exchanger, by a
modulated air flow, before passing to the fuel/oil heat
exchanger.
Heat Management System
Mode 4
Mode 4 as shown below is the mode adopted when the
burned fuel flow is low i.e. at low engine speeds with a
high H.P. fuel pump inlet temperature.
In this mode the fuel/oil heat exchanger is operating as a
fuel 'cooler' and the heat passed to the engine oil is
extracted by the air/oil heat exchanger.
Heat Management System
Mode 5
Mode 5, shown below, is the mode which is used when
system conditions demand operation as in Mode 3 but this
is not permitted because:•
•
the IDG oil system temperature is excessive,
or
fuel spill to aircraft tank is not permissible because
of high spill fuel temperatures.
Heat Management System
Fuel Diverter and Back to Tank Valve
The fuel diverter valve and the back to tank valve together
form a single unit. The unit is bolted to the rear of the
fuel/oil heat exchanger as shown below. The valves are
positioned on commands from the E.E.C.
Fuel Diverter Valve
This valve is a two position valve and is operated by a
dual coil solenoid. The control signals to energise / deenergise the solenoid come from the E.B.C.
•
•
solenoid energised :Mode 4 or 5.
Solenoid de-energised : Mode 1 or 3 (fail safe
position)
Back to Tank Valve
This valve is a modulating valve and will divert a proportion
of the L.P. fuel back to the aircraft tanks as directed by the
E.E.C.
The interface between the E.E.C. and the valve is a
modulating torque motor, the torque motor (or will direct
H.P. servo fuel to position the valve.
The fail safe position is with the valve fully closed - no fuel
return to tank
10
1 Fuel cooled oil cooler (FCOC) (79-21-43)
2 Fuel filter element (73-12-42)
3 HP servo supply connection
4 LP servo return connection
5 To/from IDG FCOC port connection
6 To/from FCOC connection
7 Fuel filter in ret connection
8 Fuel metering unit spill flow connection
9 FCOC inlet connection
10 Aircraft fuef tank inlet connection
11 Drainconnection
12 ECU LVDT connection (19 pins)
10 ECUT/M • solenoid valve - microswitch
connection (26 pinsf
FUEL DIVERTER AND RETURN TO TANK VALVE
PART ONE - SECTION 8
COMPRESSOR AIRFLOW
CONTROL SYSTEM
Compressor Airflow Control System
Introduction
The engine incorporates two air bleed systems and a
variable stator vane (VSV) system which together are
used to:•
ensure stable airflow through the compressor at "off
design" conditions
•
ensure smooth, surge free, accelerations and
decelerations (transient conditions)
•
improve engine starting characteristics
•
assist in re-stabilising the engine if surge occurs
(surge recovery)
The complete system comprises three subsystems, which
are:
•
an LP compressor air bleed located at engine
station 2.5 and known as the Booster Stage Bleed
Valve (BSBV).
•
HP compressor air bleeds on stages 7 and 10
•
the VSV system which comprises variable inlet
guide vanes, at the inlet to the H.P. compressor,
and 4 stages of variable stator vanes.
The three systems are controlled by the E.E.C.
AIR-FLOW CONTROL SYSTEM - SCHEMATIC
Booster Stage Bleed Valve (B.S.B.V.)
Purpose
B.S.B.V. Mechanical Arrangement
The B.S.B.V. bleeds air from the rear of the L.P.
compressor at engine station 2.5, the bleed air is vented
into the fan air duct.
The annular bleed valve comprises 27 flaps which are
attached by 25 link arms and 2 power arms to a
synchronise ring. Two actuating rods connect the two
power arms to two actuators. The two actuators utilise
H.P. fuel as an hydraulic medium, and are hydraulically
"linked" to ensure simultaneous movement.
The bleed valve provides improved surge margin during
starting, low power and transient operations.
The bleed valve is controlled by the E.E.C. and is fully
modulating, between the fully open and fully closed
positions, as a function of:•
•
•
Nl corrected speed
Altitude
Aircraft forward speed (Mn)
For starting the bleed valve is fully open and will
progressively close during engine acceleration, during
cruise and take off the valve is fully closed.
For
decelerations and operation in reverse thrust the valve is
opened. In the event of an engine surge the valve is
opened to enhance recovery.
The mechanical arrangement is shown below.
1
2
3
4
5
6
Support ring
Power arm (2off)
Valve flap
Link arm
Synchronize ring
Bleed duct
BOOSTER STAGE BLEED VALVE
Compressor Airflow Control
BSBV Actuators Description
The two BSBV actuators utilise HP fuel as a hydraulic
operating medium.
The actuators are located on the rear of the intermediate
casing on either side of the HP compressor.
Only one of the actuators, the one on the left hand side,
interfaces with the EEC. This actuator is called the Master
actuator, the RH actuator is called the Slave actuator.
The two actuators are hydraulically linked by two tubes
which pass across the top of the H.P. compressor case.
The master actuator incorporates an LVDT which
transmits actuator positional information back to the EEC.
The slave actuator incorporates two overload relief valves
which prevent overpressurisation of the actuators in the
case of faults, such as a mechanically seized actuator.
1
2
3
4
5
Bleed valve actuating rod
Piston jack fork end
LPC bleed-master actuator
LPC Weed-slave actuator
Intermediate structure
BSBV ACTUATORS
BSBV Master Actuator
Removal / Installation
Removal/installation of the master actuator is quite
straightforward.
The following points should be noted:all sealing rings must be discarded on removal and
new sealing rings fitted on installation
all threads should be lubricated with clean engine
oil on installation
observe the torque
maintenance manual
loadings
quoted
in
the
the bolt which secures the actuator fork end to the
actuating rod is locked by a double key washer a
new washer must be used on installation
on completion of the actuator change carry out Test
No. 1 or 3 (leak test), followed by Test No. 11 (High
Power Assurance test).
The full procedure to remove/install the B.S.B.V.
master actuator can be found in the AMM CH 7531-42
EXTEND LINE TO
SLAVE ACTUATOR
RETRACT LINE TO
SLAVE ACTUATOR
GUIDE PIN
(2 off)
HP FUEL
BLEED VALVE
ACTUATING
ROD
INTERMEDIATE
STRUCTURE
DRAIN TUBE
HARNESS
(Feed back signals)
HARNESS
BSBV MASTER ACTUATOR
REMOVAL / INSTALLATION
Compressor Airflow Control
BSBV Slave Actuator
Removal/Installation
Removal/installation
straightforward.
of
the
slave actuator is quite
The disconnect points are shown below.
Points to note are the same as the notes for the Master
actuator shown on the previous page.
For full removal/installation procedures refer to the aircraft
maintenance manual CH 75-31-43.
INTERMEDIATE
STRUCTURE
RETRACT LINE
EXTEND LINE
PISTON JACK
FORK END
BLEED VALVE
ACTUATING
ROD
FUEL DRAIN
GUIDE PIN
(2 off)
MOUNT BRACKET
BSBV SLAVE ACTUATOR
REMOVAL/INSTALLATION
Compressor Airflow Control
BSBV Electrical Harness
The BSBV electrical harness connections are as shown
below.
BSBV (2.5) BLEED ACTUATOR HARNESS
Compressor Airflow Control
Variable Stator Vane System (V.S.V.)
Introduction
The entry of air into the H.P. compressor is controlled by
Variable Incidence Stator Vanes. The variable vanes
control the angle at which the air enters the first five
stages of the H.P. compressor.
The five unison rings are connected by short rods to a
crankshaft. The crankshaft is connected by a short rod to
an actuator which utilizes HP fuel as a hydraulic operating
medium.
The angle varies with the HP compressor speed (N2), this
reduces the risk of blade stall and compressor surge.
Signals from the EEC direct HP fuel to extend/retract the
actuator. Actuator movement causes the crankshaft to
rotate, and, through the unison rings, reposition the
variable stator vanes.
The five stages of variable incidence stators comprise inlet
guide vanes to stage 3 and stages 3, 4, 5 and 6 stator
vanes.
Mechanical Arrangement
Each vane has pivots at its inner and outer ends which
allow the vane to rotate about its longitudinal axis.
The outer end of each vane is formed into a shaft which
passes through the compressor case and is attached by a
short lever to a Unison ring, (one unison ring for each
stage).
The actuator incorporates an LVDT which signals actuator
positional information back to the E.E.C.
UNISON RINGS
FUEL POWERED RAM
VARIABLE STATOR VANE
ACTUATION CRANKSHAFT
AND MOUNTINGS
FRONT BRG
HOUSING
IGV CRANK SHAFT
LEVER ASSEMBLY
STAGE 3 LEVER
STAGE 5 LEVER
BRIDGE
PIECE
REAR
BRG
ENGINE
SPLIT
CASING
FUEL
POWERED
RAM
STAGE 4
LEVER
INPUT AND STAGE 4
UNDERSIDEVIEW
OF ENGINE
HP COMPRESSOR VSV ACTUATION SYSTEM
Compressor Airflow Control
Variable Stator Vane System (S.V.S.)
Actuator Removal/Installation
The following notes summarize the removal / installation
procedures.
For a full description of the procedure refer to the aircraft
maintenance manual CH 75-32-41.
Access to the actuator, which is mounted on the HP
compressor case, L.H. lower side, is by opening the L.H.
'C duct, see Part Two, Section Two of these notes and
refer to the aircraft maintenance manual CH 78-32-00.
Note
•
The VSV actuator fuel pressure and return lines
must be drained before work begins.
•
The fuel lines are drained at the union locations
shown below.
•
The fuel is drained at this point because it is the
lowest point in the system, and also, because fuel
drained from here is least likely to contaminate the
engine electrical harness.
FUEL METERING
UNIT (FMU)
1 UNION
2 UNION
VSV ACTUATOR-FUEL DRAIN UNIONS
Compressor Airflow Control
Variable Stator Vane System (VSV)
Actuator Removal/Installation (Cont)
Before the actuator is removed it is important that the VSV
crankshaft assembly is locked in order to prevent damage
to the stator vanes.
Rig pins are provided to lock the crankshaft and the
actuator, as shown below.
After the fuel supply and return tubes have been
disconnected the crankshaft should be rotated to align the
rig pin holes in the input lever and the front bearing
housing. Spanner (Wrench) flats are provided on the
crankshaft for this purpose. Installing the rig pin locks the
crankshaft assembly with the actuator and vanes in the
high speed position (actuator fully retracted).
VSV ACTUATOR
REMOVAL/INSTALLATION (3)
Compressor Airflow Control
Variable Stator Vane System (V.S.V.)
Actuator Removal/Installation (Cont)
The actuator disconnect points are shown below.
Note: Discard the sealing rings, from the fuel lines, on
removal
•
•
fuel tubes are lock-wired
a 15/16 inch bi-hex crowsfoot spanner is required to
disconnect the fuel supply tube
ELECTRICAL
CONNECTOR
VARIABLE STATOR
VANE ACTUATOR
ELECTRICAL
CONNECTOR
LP RETURN TUBE
SEALING RING
HP SUPPLY TUBE
ELECTRICAL
CONNECTOR
FUEL DRAIN TUBE
ELECTRICAL
CONNECTOR
VSV ACTUATOR - REMOVAL/INSTALLATION (1)
Variable Stator Vane System (V.S.V.)
Actuator Removal/Installation (Cont)
The following points should be noted:•
During removal do not allow the upper support
bracket to take the full weight of the actuator since
this can damage the bracket.
•
The surfaces marked * should be cleaned with V01003 (cleaning fluid) and coated with V04-0O4
(jointing compound) on assembly.
•
Ensure the torque loading instructions contained in
the aircraft maintenance manual are carried out.
•
Fit the rigging pin to the replacement actuator, with
the actuator in the fully retracted (high speed)
position before installation.
•
If it is necessary to adjust the length of the control
rod end ensure the control rod ends are in "safety"
on completion.
•
On completion carry out Test No 1 or 3 (leak
checks) followed by Test No 11 (high powor
assurance test).
VSV ACTUATOR - REMOVAL/INSTALLATION (2)
Compressor Airflow Control
V.S.V. System - Electrical Harness
The VSV Electrical harness connect ions are as shown
below.
VSV ACTUATOR HARNESS
Compressor Airflow Control
Handling Bleed Valves
Introduction
Handling bleed valves are fitted to the H.P. compressor to
improve engine starting, and prevent engine surge when
the compressor is operating at off-design conditions.
A total of four bleed valves are used, three on stage 7 and
one on stage 10.
The handling bleed valves are two position only - fully
open or fully closed, and are operated pneumatically by
their respective solenoid control valve.
The solenoid control valves are scheduled by the E.E.C.
as a function of N2 and T2.6 (N2 corrected).
When the bleed valves are open air bleeds into the fan
duct through ports in the inner barrel of the 'C ducts.
The servo air used to operate the bleed valves is H.P.
compressor delivery air known as PB or Pb.
The bleed valves are arranged radially around the HP
compressor case as shown below.
Silencers ate used on some bleed valves.
All the bleed valves are spring loaded to the open position
and so will always be in the correct position (open) for
starting.
COMPRESSOR HANDLING BLEED VALVES
Compressor Airflow Control
Handling Bleed Valves - Location R.H.
The diagram below shows the location of the three bleed
valves mounted on the R.H. side of the engine H.P.
compressor case.
STAGE 7 UPPER RH BLEED VALVE
STAGE 10 RH BLEED VALVE
STAGE 7 LOWER RH BV
BLEED VALVES - RH
Compressor Airflow Control
Handling Bleed Valves - Location L.H.
The diagram below shows the location of the bleed valve
mounted on the L.H. side of the engine.
BLEED VALVES - LH
Compressor Airflow Control
Handling Bleed Valves
Operating Schedule
The handling bleed valves have tluec operating regimes:•
•
•
steady state
transient
surge/reverse
Operation of the bleed valve is scheduled against N2
corrected for changes of H.P. compressor (T2.6) inlet
temperature -known as N2C26.
Steady State
The valves are commanded open whenever N2C26 is
below the steady state closing speed.
Transient
The valves are commanded open at the beginning of
accelerations/decelerations and will close when either the
speed limits are exceeded or timers expire.
Surge/Reverse
The valves will be commanded open in the event of a
surge within the speed range shown.
In reverse thrust laws similar to the transient laws apply
but the reverse speeds, shown below, are used.
BLEED VALVE
REGIME
OPEN
CLOSE
Corrected N2
9000
9400
below 15000 ft
10000
10400
above 15000 ft
Transient
9000
9400
Surge/Reverse
12562
12772
7B
Steady state
7650
8000
7C
Steady state
6800
7000
Transient
11800
12250
below 15000 ft
12000
12450
above 15000 ft
Surge/Reverse
12352
12562
Steady state
7650
8000
Surge/Reverse
10667
10667
7A
10
Steady state
BLEED VALVE OPERATING SCHEDULE
Compressor Airflow Control
Handling Bleed Valves
operating Schedule
The schedule for one bleed valve (7C) is shown, in detail,
below.
Note the transient regime is slightly modified for operation
above 15000 ft but operates in the same way.
Steady State
Surge/Reverse
It can be seen that the valve will be commanded closed at
stabilized min. idle, 8600 N2C26, and will not be opened
again in Steady state.
If the engine is operating in reverse thrust operation is the
same as Transient but different speeds apply.
Transient
The valve will be commanded open during engine
acceleration whenever N2C26 is below the transient
closing speed. Thus during an acceleration from min idle
to max speed the valve will be opened and
will remain open until the speed passes the transient
closing speed. If the acceleration is to a speed below the
transient closing speed the valve will remain open until the
acceleration timer expires (30 seconds).
During decelerations the valve will be commanded open
whenever N2C26 is below the transient opening speed.
The valve remains open until the deceleration ceases and
a deceleration time, 2 seconds, expires.
In the event of an engine surge the valve will be
commanded open, if the speed is below the open speed,
and will remain open until the engine restabilises.
BLEED VALVE OPERATING SCHEDULE (7C)
Compressor Airflow Control
Handling Bleed Valves - Operation
The bleed valves and the solenoid control valves all
operate in the same manner. The operation of one bleed
valve only is described.
Bleed Valves
The bleed valve is a two position valve and is either fully
open or fully closed. The bleed valve is spring loaded to
the open position and so all the bleed valves will be in the
correct position - open -for engine start.
When the engine is started the bleed air will try to close
the valve. The valve is kept in the open position by servo
air (P3) supplied from the solenoid control valve, (solenoid
de-energised) as shown below.
P3 SUPPLY
VENT
VENT
SOLENOID
DE-ENERGISED
OPENING CHAMBER
BLEED VALVE
OPEN
SOLENOID CONTROL
VALVE
BLEED VALVE - OPERATION
Compressor Airflow Control
Handling Bleed Valves - Operations
The bleed valves will be closed at the correct time during
an engine acceleration by the EEC.
The bleed valve is closed by the EEC which energises the
solenoid control valve, as shown below.
Energising the solenoid control valve vents the P3 servo
air from the opening chamber of the bleed valve, and the
bleed valve will move to the closed position.
During an engine deceleration the reverse operation
occurs and the bleed valve opens.
P3 SUPPLY
VENT
VENT
SOLENOID
ENERGISED
BLEED VALVE
CLOSED
SOLENOID CONTROL
VALVE
BLEED VALVE - OPERATION
PART ONE - SECTION 9
SECONDARY AIR SYSTEMS
ACTIVE CLEARANCE CONTROL
SYSTEM
10TH STAGE HAKE-UP AIR SYSTEM
AIRCRAFT SERVICES BLEED SYSTEM
Secondary Air Systems
Active Clearance Control (A.C.C.) System
Introduction
The system improves engine performance by ensuring
that the HP and LP turbines operate with the optimum
turbine blade tip clearances.
This is achieved by directing a controlled flow of cooling
air to reduce the thermal growth of the turbine casings.
This minimises the increase in turbine blade tip clearances
which otherwise occurs during the climb and cruise
phases.
Operation
An air scoop directs fan air to a dual control valve which
modulates the flow to two cooling manifolds, on the H.P.
and L.P. turbine casings.
The modulating air control valves are positioned by a fuel
pressure operated actuator. The actuator input to the air
control valves is through a cam mechanism which
provides different cooling flow rates to the two separate
manifolds.
The actuator is positioned by signals from the E.E.C.
which thus controls the cooling flows as a function of:•
•
corrected N2
aircraft altitude
An actuator mounted L.V.D.T. transmits cooling valve
position feedback signals to the E.E.C.
Loss of control (EEC) or loss of fuel pressure drives the
actuator to the fail safe position to provide maximum
turbine blade tip clearances.
ACC SYSTEM - SCHEMATIC
Secondary Air Systems
Active Clearance Control
Component Location
The components in the system comprise:•
•
•
•
the HP Turbine cooling manifold
the LP Turbine cooling manifold
the dual air control valve
the actuator
ACTIVE CLEARANCE
CONTROL (ACC) ACTUATOR
ACTIVE CLEARANCE
CONTROL (ACC) VALVE
AIR FROM THRUST REVERSER DUCT
TO HP/LP CASE COOLING
HP TURBINE
CASE COOLING
LP TURBINE
CASE COOLING
FROM RH T/R DOOR
ACTIVE CLEARANCE CONTROL-INSTALLATION
Secondary Air Systems
Active Clearance Control
Operation
The A.C.C. system is shown diagrammatically below.
Operation of the system is summarized as follows:•
•
•
•
•
signals from the E.E.C. channel A or B position the
jet pipe servo valve
the jet pipe servo valve moves to direct HP fuel to
one end of the spool valve
the spool valve is positioned to port HP fuel to one
side of the actuator piston, the other side of the
piston is ported to LP fuel return
the actuator extends/retracts to position the L.P.
and HP cooling air control valves to the
commanded position
actuator movement is sensed by the LVDT which
signals actuator position to the E.E.C, channels A
and B.
JET PIPE
SERVO VALVE
90 µ FILTER
EEC
PRESSURE
PORT
DUAL LVT
FEEDBACK TRANSDUCER
RETURN
PORT
DRAIN
HP Turbine
LP Turbine
ACC SYSTEM
Secondary Air Systems
Active Clearance Control
H.P. Turbine Manifold
The assembly consists of LH and RH tube assemblies
which are a simple push fit into the manifold. The tube
assemblies are sealed off at their upper ends.
Air from the air control valve enters the manifold and is
directed to the left and right tubes.
Air outlet holes on the inner face of the tubes direct the air
onto the H.P. turbine casings.
HP TURBINE - ACC MANIFOLD
Secondary Air Systems
Active Clearance Control
L.P. Turbine Manifold
The assembly consists of upper and lower tube
assemblies with integral manifolds, both ends of the
cooling tubes are sealed.
Air from the air control valve enters a supply tube which
then splits to feed air into two tubes which supply the
upper and lower manifolds. The manifolds direct the air
into the cooling air tubes.
Air outlet holes on the inner surfaces direct the air onto the
L.P. turbine cases.
LP TURBINE - ACC MANIFOLD
Secondary Clearance Control
Operating Schedule
The graph shown below shows control valve position, and
actuator position related to operation points A to E.
Engine Stopped
With the engine stopped, the position of the actuator
piston is point A. At this point:•
The control valve for the H.P. turbine ACC is closed.
•
The control valve for the LP turbine ACC is not less
than 44 per cent opened.
Engine Operation
During engine operation, the E.E.C. controls the position
of the actuator piston between point B and point E.
Take-off
During take-off, the position of the actuator piston is at
point C. At this point:•
The control valve for the HP turbine ACC is closed.
•
The control valve for the LP turbine ACC is not less
than 70 per cent opened.
Note
The actuator position between point C and point E
depends on Altitude.
Fail Safe
When there is no torque motor current or no fuel servo
pressure, the actuator piston moves to point A. The
actuator piston remains at this point at all defective
conditions.
PISTON TRAVEL (%)
ACC OPERATING SCHEDULE
Secondary Air Systems
Active Clearance Control
The ACC Electrical harness connections are as shown below.
HPT/LPT ACC HARNESS
Secondary Air Systems
10th Stage Make-up Air System
Introduction
The purpose of this system is to provide additional cooling
airflows to the H.P. turbine 2nd stage disc and blades.
The cooling air used is taken from the 10th stage manifold,
and is controlled by a two position pneumatically operated
valve.
The valve position is controlled by the EEC as a function
of corrected N2 and altitude.
Operation
Signals from the EEC will energise/de-energise the
solenoid control valve. This directs pneumatic servo
supplies to position the 10th stage air valve to the
open/close position.
In the open position the valve allows 10th stage air to flow
through two outlet tubes down the left and right hand side
of the diffuser case and then pass into the engine across
the diffuser area. The air then discharges into the area
around No4 bearing housing.
The make up air supplements the normal airflows in this
area and increases the cooling flow passing to the H.P.
turbine, stage 2.
The EEC will keep the air valve open at all engine
operating phases except cruise. The valve incorporates a
micro switch when transmitting valve position feedback
signals to the EEC.
The failsafe position is valve open.
MAKE-UP AIR VALVE - OPERATION
MAKE-UP AIR SYSTEM -SCHEMATIC
Secondary Air Systems
10th Stage Make-up Air System
Component Location
The components in this system comprise:•
•
•
the two-position stage 10 on-off valve bolted to the
10th stage manifold at the top of the engine
compressor case.
the solenoid control valve located on the lower RH
fan case.
two air supply tubes.
TURBINE COOLING CONTROLLED
AIR TUBES
ON/OF
VALVE
SOLENOID
CONTROL
VALVE
VIEW ON A
MAKE-UP AIR SYSTEM - COMPONENT LOCATION
Secondary Air Systems
10th Stage 'Make-up' Air System
10th Stage Air Valve Removal/Installation
These notes are for guidance only, reference must be
made to the Maintenance Manual -75-23-51.
Access to the valve is by opening the 'C ducts.
The disconnection and location points are shown below.
The removal and installation procedure is straightforward
but the following points should be noted:•
•
a new 'C seal (3) must be fitted.
the threads of the retaining bolts (4) must be
cleaned and coated with anti-seize compound,
1
2
3
4
5
6
7
Electrica (connector)
Outlet air tube
C-seal
Bolt
Air Offtake to No. 4 Bearing Scavenge Valve
Lock wire
Servo air tube
10 STAGE AIR VALVE REMOVAL/INSTALLATION
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
The system provides pressure/flows for:•
•
•
•
•
cabin pressurisation and conditioning
wing anti-icing
engine crossfeed starting
hydraulic system pressurisation
water system pressurisation
The required Sir is bled from the HP compressor of each
engine.
APU BLEED VALVE
TO WING ANTI-ICING
CROSS8LEED VALVE
TO AIR CONDITIONING PACKS
TO WING ANTI-ICING
HP GROUND CONNECTOR
PRECOOLER
FAN AIR VALVE
OVER PRESSURE VALVE
IP CHECK VALVE
PRESSURE REGULATING VALVE
HP VALVE
AIRCRAFT PNEUMATIC SYSTEM MANIFOLD
Secondary Air System
Aircraft Pneumatic System
The aircraft pneumatic system is shown, schematically,
below.
No 1 Engine installation only is shown, No 2 engine is
identical.
AIRCRAFT PNEUMATIC SYSTEM
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
The system is identical on each engine.
One engine system is shown, schematically below.
Two air off-takes are provided, one each from:. stage 7
. stage 10
Air is taken from stage 10 or stage 7.
Stage 10 supplies the requirements at low power
setting.
Stage 7 supplies the requirements at higher
power settings.
Automatic change over from stage 10 to stage 7
occurs during engine acceleration.
TO AIRCRAFT
PNEUMATIC
SYSTEM
TO NOSE COWL
TAI
OVERBOARD
STARTER
AIR
VALVE
STARTER
STAGE 7 STAGE 10
HPC
AIR OFF-TAKES - SCHEMATIC
1
2
3
4
5
HP Regulating valve
PRV
0/PV
FAV
Check valve
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
Component Location
The engine mounted components of the aircraft pneumatic
system comprise:•
•
•
•
high stage control valve
pressure regulating valve
check valve
associated ducting up to the engine / aircraft
interface
All these components are located on the L.H. side of the
engine core as shown below.
The remainder of the components of the system are
supply, fit and responsibility of the aircraft manufacturer.
STARTER
DUCT
OVER/PRESSURE
VALVE
PRECOOLER
INLET COWL
ANTI-ICE DUCT
FAN AIR
7TH STAGE
CHECK VALVE
ANTI-ICE
VALVE
STARTER
DUCT
REGULATING
VALVE
HIGH STAGE
CONTROL VALVE
CORE COMPARTMENT
TEMPERATURE SENSOR
NACELLE - PNEUMATIC SYSTEM - COMPONENTS
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
The system operates under the control of the Bleed Air
Monitoring Computer (BMC) and will automatically:•
•
•
select the compressor stage from which air is bled.
regulate bleed air pressure.
control bleed air temperature.
TCT
CTS
A/C
SYSTEMS
TLT
PRECOOLER
PNEUMATIC SYSTEM SCHEMATIC
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
Sensing Lines
The arrangement of the pneumatic sensing lines is shown
below.
PRESSURE
REGULATING
VALVE
CHECK
VALVE
HIGH PRESSURE
CONTROL VALVE
PNEUMATIC SYSTEM COMPONENTS-SENSE LINES
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
The air bleed electrical harness connections are as
shown below.
402VC
404VC
447VC
DUAL OUTPUT
454VC
450VC
PRESSURE
REGULATING
VALVE V
HIGH STAGE
VALVE
D510P(A)
D550P(B)
HYDRAULIC LP
WNG SWITCH
MISC SYSTEMS HARNESS
ANTI-ICING
CONTROL VALVE
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
High Stage Control Valve Removal / Installation
Refer to M.M.
Removal
Refer to AMM, Task 36-11-51-000-010
Installation
Refer to AMM, Task 36-11-51-400-010
The procedure is illustrated below.
Note
Discard the seals on removal and always fit new seals on
installation
ELECTRICAL
CONNECTOR
MANUAL
LOCKOUT PIN
PRESSURE FROM
PRV
CLOSED POSITION
OPEN POSITION
POSITION INDICATOR
HIGH PRESSURE VALVE - REMOVAL/INSTALLATION
Secondary Air Systems
Engine Air Bleeds - Aircraft Services
Pressure Regulating Valve R & I
Removal
Refer to AMM, Task 36-11-52-000-010
Installation
Refer to AMM, Task 36-11-52-040-010
Note
Discard the seals on removal and always fit new seals on
installation
PRESSURE FROM
TEMPERATURE
LIMITATION
THERMOSTAT (TLT)
ELECTRICAL
CONNECTOR
PRESSURE TO
HIGH PRESSURE
VALVE
MANUAL
LOCKOUT PIN
(STOWED)
MANUAL
LOCKOUT
PIN
OPEN
ROTATED VIEW 180°
POSITION INDICATOR
PRESSURE REGULATOR VALVE - REMOVAL/INSTALLATION
PART ONE - SECTION 10
ENGINE ANTI-ICE SYSTEMS
Engine Anti-Icing System
General
P2/T2 Probe Heating
Ice may form in the inlet cowl when the engine is operating
in conditions of low temperature and high humidity.
The P2/T2 probe is continuously heated, during engine
operation, by an integral 115V heating coil.
Ice build up in, and on, the inlet cowl could affect engine
performance and could cause compressor damage from
ice ingestion. To prevent ice formation anti-icing protection
is provided in the following areas:-
Rotating Fairing
•
•
•
the inlet cowl (thermal).
the P2/T2 probe (thermal)
the rotating fairing - spinner (dynamic)
Inlet Cowl Anti-Icing
The inlet cowl anti-icing system utilises air taken from the
7th compressor stage which is ducted down the R.H. side
of the engine to an anti-icing control valve located on the
rear diaphragm of the nose cowl.
From the anti-icing control valve the hot air passes to a
distribution manifold located inside the inlet cowl lip. The
used air is vented overboard through an exhaust grille on
the lower R.H. side of the inlet cowl.
The rotating fairing (spinner) is protected against ice build
up by a solid rubber nose tip which vibrates naturally to
break up and dislodge the ice immediately it starts to form.
ANTIICE
ENG 1
7TH STAGE
DUCT
INLET COWL
ANTI-ICE VALVE
ANTI ICE DISCHARGE
GRILLE
ANTI-ICE SYSTEM - INSTALLATION
ENG2
Engine Anti-Icing System
Inlet Cowl Ducting
Hot air from the anti-icing control valve is ducted between
the inner and outer skins of the inlet cowl to a spray ring
located inside the inlet cowl lip skin.
The spray ring has multiple outlets which direct the hot air
to heat the inner surfaces of inlet cowl lip skin. The air is
then exhausted overboard through an outlet grille on the
lower R.H. side of the inlet cowl.
The arrangement is shown below.
FORWARD BULKHEAD
AFT BULKHEAD
FORWARD
BULKHEAD
AIR SUPPLY
EXHAUST DUCT DISCHARGE
VIEW ON A
INLET COWL ANTI-ICE DUCT
Engine Anti-Icing System
Anti-Icing Control Valve
The anti-icing control valve is located in the air supply
ducting, as shown below, which is attached to the rear
of the inlet cowl.
Manual Override and Lock
The valve can be manually overridden and locked in
either the open or closed position. Without inlet
pressure applied to the unit, the valve can be
manually moved to the desired position by applying a
standard wrench to the hexagonal nut attached to the
butterfly shaft. The valve is locked in the selected
position by removing the locking pin from its stow
position and inserting it through the locking hole on
the valve exterior and the mating hole in the valve
piston. The pin is retained by a ball detent mechanism
in the end of the locking pin.
LOCKING PIN
OP
(OPEN)
RING
MID
CL
CLOSE
LOCK
RING
INLET COWL ANTI-ICE VALVE - REMOVAL/INSTALLATION
Engine Anti-Icing System
Anti-Icing Control Valve Description and operation
The anti-icing control valve is a butterfly type valve located
in the anti-icing supply duct.
The valve is positioned through a mechanical linkage by a
piston which is operated by pneumatic pressure. The
piston is spring loaded to the valve closed position.
The two valve positions are shown below.
ANTI-ICING OFF
Solenoid Energised
Inlet supply pressurised
When the solenoid is energised, the solenoid plunger is
retracted against the spring force. Inlet pressure shuttles
the pilot ball against the pilot vent seat, thus admitting inlet
pressure to chamber "A". Inlet pressure acting on the
chamber "A" side of the piston, overcomes the opposing
force of inlet pressure acting on the chamber "B" side of
the piston which has less effective area. This pneumatic
closing force, aided by the power spring force, moves the
valve to the closed position.
ANTI-ICING ON
Solenoid De-energised
Inlet supply pressurised
When the solenoid is de-energised, the solenoid plunger is
extended by the spring. The pilot ball is held against the
inlet pressure seat on the pilot valve, thus venting
chamber "A" to ambient. Inlet pressure in chamber "B",
acting on the reduced piston area, overcomes the power
spring force and moves the valve to the open position.
DUCT PRESSURE
ANTI-ICE OFF
SOLENOID ENERGISED
VALVE CLOSED
ANTI ICE ON
SOLENOID DE ENERGISED
VALVE OPEN (Fail safe position)
INLET COWL ANTI-ICE VALVE
Engine Anti-Icing System
Inlet Cowl Ducting
The ducting which connects the anti-icing control
valve to the inlet cowl ducting is shown below
INLET COWL ANTI-ICE DUCTING - RH NLET COWL
PART ONE - SECTION 11
ENGINE INDICATING
Engine Indicating
Flight Deck Indications
The primary engine parameters listed below are displayed
on the upper ECAM CRT.
Primary Engine Display
•
•
•
•
Engine Pressure Patio (EPR)
Exhaust Gas Temperature (EGT)
N1
N2
Secondary Engine Display
The secondary engine parameters listed below are
available for display on command or, during engine start,
automatically.
•
•
•
•
•
•
•
•
•
Total fuel used
Oil quantity
Oil pressure
Oil temperature
Nacelle temperature
Vibration – N1 and N2
Oil filter clog
Fuel filter clog
No 4 bearing Scavenge Valve position
FLIGHT DECK INDICATIONS
Engine Indicating
Speed Indicating System
Purpose
To provide signals of N1 & N2 speeds to be used for:•
•
flight deck indications
the EEC control circuits
Note
In addition to the speed signals a dedicated signal from
the LP rotor (N1) is provided for trim balancing operations.
Type
N1: speed probes used in conjunction with a phonic wheel
N2: uses the frequency of the dedicated alternator output
Trim Balance: speed probe (used with the phonic
wheel as N1).
Operation
A schematic arrangement of the speed indicating system is
shown below.
FRONT BEARING
COMPARTMENT
BREAK CONNECTION
(Bifurcation panel)
SPEED
PROBE 'A'
SPEED
PROBE 'B'
ARINC
429 RMS
EEC
SPEED
PROBE
(Spare)
N1
SIGNALS
N1 and
N2
signals)
TRIM
BALANCE
PROBE
TO ELECTRONIC
CENTRALIZED
AIRCRAFT
MONITORING
SYSTEM
(ECAM)
AIRCRAFT/ ENGINE
INTERFACE
JUNCTION
BOX
MAIN
ACCESSORY
GEARBOX
CHANNEL A' POWER SUPPLY
CHANNEL B' POWER SUPPLY
DEDICATED
GENERATOR
N2 SIGNALS
ENGINE SPEED MEASUREMENT -SCHEMATIC
TO ENGINE
VIBRATION
MONITORING
UNIT (EVMU)
Engine Indicating
Speed Indicating System
Speed Probes (N1)
The probes comprise two pole pieces, a permanent magnet,
and a coil wound on to one of the pole pieces. The pole pieces
span two teeth of the phonic wheel.
The phonic wheel is an integral part of the fan stubshaft and
has 60 teeth.
As the shaft rotates and the teeth of the phonic wheel pass the
pole pieces, a voltage 'pulse' is produced in the winding. The
number of pulses produced is directly proportional to the speed
of the shaft. This signal is passed to the E.E.C., processed and
is used to display N1 speed on the flight deck and also for the
engine control circuits as required.
Trim Balance Probe
The signal from this probe is only used during trim balance
operations and provides the phase relationship between any
out of balance forces present and a datum position.
The trim balance probe senses the passage of one specially
modified tooth on the phonic wheel and produces one pulse per
revolution.
POLE PIECES
LP STUB
SHAFT
SPEED PROBE
PERMANENT
MAGNET
FAN SPEED PROBE
PHONIC WHEEL
PERMANENT
MAGNET
POLE PIECES
TRIM BALANCE
PROBE
1 PULSE/REVOLUTION
SLOT
ENGINE SPEED MEASUREMENT
SPEED PROBES
Engine Indicating
Speed Indicating System
Speed Probes - Location
The N speed probes and the trim balance probe are
located, as shown below, in the front bearing
compartment.
No 3 STRUT
No 2 BEARING
SUPPORT
INNER
STRUT
FAN SPEED
PROBE
RWARD
LP STUB
SHAFT
TRIM BALANCE
PROBE
LP SHAFT
PHONIC WHEEL
ENGINE SPEED MEASUREMENT
SPEED PROBES LOCATION
1 Terminal block
2 NF tube
3 Electrical lead
4 Fan speed
probe
5 Trim balance
probe
Engine Indicating
Exhaust Gas Temperature (E.G.T.)
Indicating System
The EGT is measured by 4 thermocouples which are
located in the turbine exhaust case support struts (engine
station 4.9).
The 4 thermocouples are connected to the junction box by
a thermocouple harness.
The materials used for the thermocouples and harness are
Chrome (CR) and Alumel (AL).
An extension harness connects the EGT junction box to
channels A and B of the EEC.
Indication
The EGT indication appears on the upper ECAM display
unit. The ECAM provides the EGT indication :
•
•
in analog form with a pointer which deflects in
front of a dial
in digital form, in the lower section of the dial
The indication is normally green.
When EGT is greater than (TBD) deg C:
•
•
•
•
the indication becomes amber
the MASTER CAUT light comes on
accompanied by the single chime
The following message appears on the ECAM (TBD)
When EGT is greater than (TBD) deg C :
•
•
•
•
the indication becomes red
the MASTER WARN light comes on
accompanied by the repetitive chime
the following message appears on the ECAM (TBD)
If the EGT has exceeded TBD deg C :
• the maximum value reached is memorised
• a small red line remains positioned on the analog
scale, at that value (max point).
T/C 3
A CIRCUIT
B CJRCUJT
T/C 4
JUNCTION BOX
T/C 2
CHROMEL B CIRCUIT
CHROMEL A CIRCUIT
ALUMEJ 8 CIRCUIT
ALUMEI A CIRCUIT
EGT MANAGEMENT - SCHEMATIC
T/C 1
Engine Indicating
EGT Indicating System
Thermocouple locations
The 4 thermocouples are located in four of the hollow LP
turbine exhaust struts.
PISTON RING
SEALS
ALUMEL STUDS
VANE No IT
(Similar - 3, 6 & 8)
CHROMEL
VIEW ON B
TYPICAL AT 4
POSITIONS
EGT MEASUREMENT - THERMOCOUPLE
Engine Indicating
EGT Indicating System
EGT Junction Box
The location of the EGT junction box is shown below.
ENGINE FLANGE
ACAC EXHAUST
DUCT
FIRE DETECTION
SUPPORT BRACKET
EGT JUNCTION BOX
Indicating
The P3/T3 Sensor
The P3/T3 sensor is a dual purpose aerodynamically
shaped probe.
It measures the pressure and temperature of the air
stream at the inlet of the diffuser case.
The data which results is transmitted to the EEC for
control purposes.
BOLT
P3/T3
RIGHT LOWER FRONT-COWLS OPEN
BOSS
DIFFUSER CASE ASSEMBLY
P3/T3 PROBE
Engine Indicating
E.G.T. and P3/T3 Measurement
Electrical Harness
The EGT and T3 measurement harness electrical
connections are shown below.
TEMPERATURE MEASUREMENT HARNESS
ENGINE Indicating
Engine Pressure Ratio (EPR) Indicating System
The Engine Pressure Ratio (EPR) is used to set and
control the engine thrust
EPR = P4.9 : P2
P2 is measured by the P2/T2 Probe located in the inlet
cowl.
The associated indications are:•
EPR max : thick amber line
•
EPR limit : max EPR value corresponding to thrust
limit mode, which can be:
o
o
o
o
P4.9 is measured by a pressure rake located at the
LP Turbine exhaust.
The pressures from these sensors are routed to the EEC.
The EEC processes the pressure signals to totm actual
EPR and transmits the EPR value to the ECAM for display
on the upper screen. Each of the two EEC channels
perform this operation independently.
EPR: normal.
EPR: if exceeds limit.
EPR: if exceeds Red line limit (TBD) with master caution
and single chime.
G.A Go Around mode
FLX Flexible Take Off mode
MCT Maximum Continuous Thrust mode
CL
Climb mode
•
Flex T/O Temperature
Assumed temperature entered by crew through the
FMS - MCDU
•
EPR Reference
Predicted EPR value according to TRA.
EPR INDICATIONS
Engine Indicating
P2/T2 Sensor
The P2/T2 sensor is a dual purpose probe which
measures the total air temperature and pressure in the
inlet air stream. The temperature and pressure signals are
fed to the E.E.C
The sensor is installed at the 12 o'clock position in the air
inlet cowl.
The temperature is measured by two platinum resistance
elements, each channel of the E.E.C. monitors one of the
elements.
The pressure signal is fed to a pressure transducer in the
E.E.C.
The sensor is electrically heated to provide anti-ice
protection. The E.E.C. software corrects any temperature
signal errors caused by heating.
PROBE
ELECTRICAL
CONNECTIONS
REAR
BULKHEAD
P2/T2
MOUNTiNG
BOLTS
MOUNTING BRACKET
SEALING BLOCK
PROBE
PRESSURE
CONNECTION
P2/T2 PROBE
ACOUSTIC LINER
P2/TZ PROBE - INSTALLATION
Engine Indicating
P2/T2 Sensor - Pneumatic Line
The single pressure signal from the P2/T2 probe is routed
to a pressure transducer located within the E.E.C. The
pressure transducer has two outputs, one to channel A,
one to channel B.
Routing of the pressure signal - probe to E.E.C. - is shown
below.
PY LON
INLET COWL
REAR BULKHEAD
ELECTRONIC ENGINE
CONTROL UNiT
RELAY BOX
P2/T2 PROBE PNEUMATIC LINE - ROUTING
Engine Indicating
P2/T2 Probe - Electrical Harness
The electrical harness connections for the P2/T2 probe
are as shown below.
Note the probe anti-icing heater utilises 115V AC from the
aircraft: electrical system.
P2/T2 PROBE HARNESS
Engine Indicating
Engine Vibration Indicating System
The system monitors engine vibration for engine 1 and
engine 2. Monitoring is done by a vibration transducer on
each engine fan case. This produces an electrical signal in
proportion to the vibration detected and sends it to the
Engine Vibration Monitoring Unit (EVMU). Two channels
come from each engine. The EVMU provides signals of
Vibration, N1, N2 for display on the Engine page of the
ECAM.
The vibration transducer is installed on the fan case at the
top left side of the engine. It is attached with bolts and is
installed on a mounting plate.
Indications
The engine vibration indications are displayed in green on
the lower ECAM display unit on the engine and cruise
pages.
The ECAM display unit receives the information through
the ARINC 429 data bus via the SDAC 1 and SDAC 2.
If the advisory level is reached, the indication flashes (0.6
sec bright, 0.3 sec normal).
If the indication is not available, the corresponding
indication is replaced by 2 amber crosses.
VIBRATION ACCELEROMETER - REMOVAL/INSTALLATION
PART ONE - SECTION 12
STARTING AND IGNITION SYSEM
Engine Starting and Ignition System
General
The system comprises:•
•
•
•
•
a pneumatic starter motor
a starter air control valve
a dual ignition system
pneumatic ducting
start control panels on the flight deck
Starting and cranking cycles are initiated through
switches/controls on the engine start panel and provide,
on selection:•
•
•
•
•
automatic starting
manual starting
dry cranking
wet cranking
continuous ignition
Pneumatic supplies for the starter motor can be provided
by:•
•
•
the A.P.U.
ground supply
the other engine (if already started)
During auto start attempts the critical parameters are
monitored by the FADEC and in the event of a faulty start
i.e.
•
•
•
•
hot start
hung start
no light-up
HP fuel valve failure
the FADEC performs an automatic shut down (start abort)
and provides a motoring cycle to clear fuel vapours and
cool the engine.
SIMPLIFIED STARTING SYSTEM
Engine Starting and Ignition System
Starter Air Duct
Air supplies for the pneumatic starter motor may be
supplied from:•
•
•
the aircraft APU.
the other engine - if already running
a ground starter trolley
Note: minimum duct pressure for starting should be
approx 25 psi.
All ducting in the system is designed for high pressure and
high temperature operation and gimbal joints are
incorporated to permit working movement.
Air leakage is prevented by E-type seals interposed
between all mating flanges, mating flanges are secured by
Vee-band coupling clamps.
STARTER
STARTER
SHUT OFF
VALVE
STARTRT DUCT- INSTALLATION
Engine Starting and Ignition System
Starter Motor
The pneumatic starter motor is mounted on the forward
face of the external gearbox and provides the drive to
rotate the HP compressor to a speed at which light up can
occur.
The starter motor is connected by ducting to the aircraft
pneumatic system.
The starter motor ge.as and bearings arc lubricated by an
integral lubrication system.
Servicing features include:•
an oil filler-level plug
•
drain plug with a built in magnetic chip detector
The starter motor is attached to the external gearbox by a
quick attach/disconnect adaptor (QAP).
A quick detach Vee clamp connects the tarter
motor to the adaptor.
(FGEARBOX)
STARTER
INSTALLATION
COUP I ING
STARTER ADAPTER
ENGINE
STARTER
ENGINE STARTER-INSTALLATION
Engine Starting and Ignition System
Starter Motor - Operation
The starter is a pneumatically driven turbine unit that
accelerates the H.P. rotor to the required speed for engine
starting. The unit is mounted on the front face of the
external gearbox.
The starter comprises a single stage turbine, a reduction
gear train, a clutch and an output drive shaft - all housed
within a case incorporating an air inlet and exhaust.
Compressed air enters the starter, impinges on the turbine
blades to rotate the turbine, and leaves through the air
exhaust. The reduction gear train converts the high speed,
low torque rotation of the turbine to low speed, high torque
rotation of the gear train hub. The ratchet teeth of the gear
hub engage the pawls of the output drive shaft to transmit
drive to the external gearbox, which in turn accelerates the
engine H.P. compressor rotor assembly.
When the air supply to the starter is cut off, the pawls
overrun the gear train hub ratchet teeth allowing the
turbine to coast to a stop while the engine HP turbine
compressor assembly and, therefore, the external gearbox
and starter output drive shaft continue to rotate.
When the starter output drive shaft rotational speed
increases above a predetermined rpm centrifugal force
overcomes the tension of the clutch leaf springs, allowing
the pawls to be pulled clear of the gear hub ratchet teeth
to disengage the output drive shaft from the turbine.
TURBINE ROTOR
REDUCTION GEARS
DRIVE SHAFT
AIR INLET
DISENGAGING MECHANISM
TURBINE
ROTOR
GEAR TRAIN
TURBINE
NOZZLES
CLUTCH
AIR INLET
TWO PIECE
OUTPUT SHAF
AIR EXHAUST
(Elongated for clarity)
STARTER - SCHEMATIC
ROTATING ANNULUS
Engine Starting and Ignition System
Starter Air Control Valve
The starter air control valve is a pneumatically operated,
electrically controlled shut-off valve positioned on the
lower right hand side of the L.P. compressor (fan) case.
The start valve controls the air flow from the starter air
duct to the starter motor. The start valve basically
comprises a butterfly type vaive housed in a cylindrical
valve body with in-line flanged end connectors, an
actuator, a solenoid valve and a pressure controller.
Manual Operation
The starter air valve can be opened/closed manually using
a 0.375 in square drive. Access is through a panel in the
R.H. fan cowl. A valve position indicator is provided on the
valve body.
A micro switch provides valve position feed back
information to the FADEC.
STARTER SHUT-OFF VALVE - INSTALLATION
Starting and Ignition System Starter
Air Valve - Operation
Valve Opening
Air from immediately upstream ot the butterfly valve is
filtered and routed through an orifice in the solenoid valve.
Air upstream of the orifice is also admitted to the smaller
piston of the double acting actuator. When the solenoid is
energised the ball valve opens to admit air to the larger
piston whilst simultaneously closing the vent port. The air
acting on the larger piston oveicomes the combined force
of upstream air pressure acting on the smaller piston and
the actuator spring. Movement of the actuator is translated
through the linkage to rotate the butterfly valve towards
the open position.
Valve Closing
When the solenoid is de-energised, at approximately 6000
rpm (43%) N2, the ball valve closes and air acting on the
larger piston is vented to atmosphere through the vent- Air
pressure and actuator spring pressure acting on the
smaller piston then closes the butterfly valve. Any loss of
air pressure will cause the butterfly valve to close under
the action of the actuator spring.
STARTER AIR VALVE -SCHEMATIC
Starting and Ignition System
Starter Air Valve - Electrical Harness
The starter air control valve electrical harness connections
are as shown below.
Starting and Ignition System
Ignition
Two independent ignition systems are provided.
Continuous ignition may also be selected manually.
The system comprises:-
The ignition exciters provide approx 22.26 KV and the
igniter discharge rate is 1.5/2.5 per second.
•
•
•
two ignition exciter units' - located as shown below
two igniter plugs - located in the combustion system
adjacent to No's 7 & 8 fuel spray nozzles
two air cooled High Tension ignition connector
leads
Dual ignition is automatically selected for: •
•
•
all in flight starts
manual start attempts
continuous ignition
Single alternate ignition is automatically selected for
ground auto starts.
Continuous ignition is automatically selected when the
engine anti-ice system is 'ON' or when the aircraft flaps
are extended for:•
•
•
take-off
approach
landing
Test
Operation of the ignition system can be checked on the
ground, wj th the engine shut down, through the
maintenance menu mode of the CFDS.
CLAMP
INPUT LEAD
IGNITER PLUG
COOLING AIR
INLET
COOLING AIR
EXHAUST
HIGH
TENSION LEAD
(inside cooling jacket)
IGNITION EXCITER
IGNITION SYSTEM
Starting and Ignit ion System
Ignition - Relay Box
The ignition system utilises 115VAC supplied from the AC
115V Normal and standby bus bars to the relay box.
The 115V relays - which are used to connect / isolate the
supplies - are located in the relay box and are controlled
by signals from the E.E.C.
Note the same relay box also houses the relays which
control the 115V AC supplies for P2/T2 probe heating.
RELAY BOX
Starting and Ignition System
Electrical Harness
The ignition system electrical harness connections are
shown below.
Starting and Ignition System
Engine Starting
Controls and Indication
Two Master switches and a 3-position rotary control switch
are mounted on the pedestal for auto mode engine start
and pushbutton switches are mounted on the overhead
panel for alternate mode (manual) start.
The rotary switch is used in conjunction with either Master
switch during auto mode starts or with MAN START push
button switches in the alternate mode.
IGN position
•
•
used to perform starting in the auto and alternate
modes.
to call ENG START page on the lower ECAM display
unit.
NORM position
•
•
rest position after engine starting.
clears ENG START page
CRANK position
•
used to pet form dry or wet motoring.
Indication is displayed by lower and upper ECAM display
units and by warning lights.
The ENG MAN START button incorporates a blue ON
legend and is normally in the released position with the
ON legend off. Pressing the switch opens the pneumatic
starter valve and illuminates the ON legend. An amber
"fault" warning light on the pedestal mounted control panel
illuminates when a disagreement occurs between the
pneumatic starter valve position and that commanded by
the starting sequence in the "auto" mode. ECAM lowet
and upper screens display various warning and caution
messages as well as the ENGINE start page.
STARTING - CONTROLS AND INDICATIONS
Starting and Ignition System
Electrical Control
Engine Interface Unit (EIU)
Receives discrete electrical signals from the cockpit.
Digitizes these signals and transmits them to the EEC.
Also sends discrete signals to close air conditioning pack
flow valves and to speed up the Auxiliary Power Unit
(APU) if required.
Electronic Engine Control (E.E.C.)
The EEC generates pneumatic starter valve opening /
closing signal in respect of control switch selection (rotary
selector, master lever, MAN START push button switch)
and N2 speed signal.
Generates warning and caution messages for display on
the ECAM.
Starting and Ignition System
Start Procedures
(A) Pre-start
2. Manual Start Procedure
• thrust lever
: IDLE
• master switch
: OFF
• mode selector
: AUTO IGN
• manual start button
: OFF
• aircraft booster pumps
: ON
(B) Start
The engine may be started using the:
(1) Auto Start or
(2) Manual Start procedure
1. Auto Start Procedure
• mode selector
• master switch
• on completion of start,
: IGN/START
: ON
: return mode selector to NORM
• mode selector
• manual start button
• when N2 reaches max
motoring speed (min 15%)
• on completion of start
: IGN
: ON
: master switch ON
: mode selector to AUTO IGN
manual start push button OFF
Starting and Ignition System
Start Procedures
Auto Start - Automatic actions
•
•
•
at approx 15% N2: ignition is activated
at approx 19% N2: fuel PRSOV opens
at approx 42% N2: starter valve closes
ignition off
Manual Start
•
•
•
•
Starter air valve opens when manual start push
button is pressed ON
Fuel PRSOV opens and ignition is switched on
when master switch is moved to ON position
Starter air valve closes and ignition is cancelled
automatically at approx 42% N2
Manual start push button must be selected OFF on
completion of start
General
Auto starts or manual starts can be interrupted at any time
by moving the Master switch to the OFF position.
Starting and Ignition System
Wet and Dry Cracking (Motoring)
A dry motoring cycle will be required to:
•
•
ventilate (blow through) the engine
carry out leak checks.
Method
•
•
During this operation the starter motor is engaged but the
fuel PRSOV remains closed and both ignition systems are
OFF.
Method
•
•
Place ignition mode selector to CKANK
on overhead panel, push MAN START push button
for appropriate engine
o
o
o
ON legend illuminates
starter air valve opens
engine accelerates to max motoring speed
Wet Motoring
During wet motoring cycles the starter motor is engaged,
fuel is ON, both ignition systems are OFF.
Place ignition mode selector to CRANK
On overhead panel, push MAN START push button
for appropriate engine
ON legend illuminates
starter air valve opens
o
engine accelerates to max motoring speed
Place master control switch to ON
o
o
•
o
fuel PRSOV opens
V2500 COURSE NOTES
PART TWO - NACELLE
V2500 COURSE NOTES
PART TWO - NACELLE
CONTENTS
PART TWO - SECTION 1
INTRODUCTION
PART TWO - SECTION 2
NACELLE – MECHANICAL ARRANGEMENT
PART TWO - SECTION 3
THRUST REVERSEH
PART TWO - SECTION 4
NACELLE VENTILATION AND FIRE PROTECTION
PART TWO - SECTION 5
ENGINE REMOVAL / INSTALLATION
PART TWO - SECTION 1
INTRODUCTION
Introduction - Nacelle
The propulsion unit comprises the Engine and the Nacelle.
The major components which comprise the Nacelle are:•
the air inlet cowl
•
the fan cowls (left and right hand)
•
the C - ducts (left and right hand) which incorporate
the hydraulically operated thrust reverser unit
•
the combined nozzle assembly
Components Weights
Nose Cowl
: 238 lbs (107.98 kg)
Fan Cowl R.H.
: 86 lbs (39.01 kg)
Fan Cowl L.H.
: 79 lbs (35.84 kg)
C Ducts
: 578 lhs (each) (262.25 kg)
FAN AiR VALVE
DISCHARGE
P2/T2
PROBE
C DUCT
TRANSLATING COWL
CARBON FIBRE
ANT-ICING
D CHAMBER
ALUMINIUM
FAN COWL DOORS
CARBON-FIBRE
INLET
COWL
CARBON FIBRE
V2500 NACELLE
COMMON NOZZLE
ASSEMBLY (CNA)
TITANIUM
Introduction - Nacelle
Access to Engine Mounted Units
Access to units mounted on the LP compressor (fan) case
and external gearbox is gained by opening the hinged fan
cowls.
Access to the core engine, and the units mounted on it, is
gained by opening the hinged 'C ducts.
RIGHT HAND C-DUCT
THRUST REVERSER
PYLON
RIGHT HAND
FAN COWL DOOR
COMMON NOZZLE
ASSEMBLY (CNA)
HOLD OPEN
ROD
LEFT HAND C-DUCT
THRUST REVERSER"
INLET
COWL
HOLD
OPEN
ROD
NACELLE SYSTEM-DETAILS
LEFT HAND FAN
COWL DOOR
Introduction - Nacelle
Access Panels and Doors
Access panels and servicing doors are provided as shown
below.
ACTUATOR ACCESS
(TYPICAL 4 PLACES)
P2/T2 PHOBE
ACCESS
EEC GOOLING
AIR EXHAUST
EEC GOOLING
AIR INLET
INIfcHPHONI.
JACK
ANI-ICE
DISCHARGE
GRILLE
AIR COOLED
OIL COOLER
GEARBOX
OVERBOARO
DISCHARGE
MASTER CHIP
DETECTOR
STOW LOCKOUT
(TYPICAL 2 PLACES)
STARTER SHUTOFF
VALVE/PRESSURE
RELIEF DOOR
OIL TANK
SERVICE DOOR
VENTILATION
EXIT GRILLE
DRAIN MAST
NACELLE ACCESS PANELS AND DOORS
STOW LOCKOUT PIN
STOWAGE (TYPICAL
PART TWO - SECTION 2
NACELLB MECHANICAL ARRANGEMENT
NACELLE
Mechanical Arrangement
The propulsion unit comprises the Engine and the Nacelle.
The major assemblies which together form the propulsion
unit are shown below.
1. Inlet Cowl
Bolted to the LP Compressor (Fan) case.
2,3. Fan Cowl Doors
Secured to the strut at 4 hinge points on each side.
4,5. C-Ducts/Thrust Reverser
Secured to the strut at 4 hinged points on each side.
The C - ducts incorporate the hydraulically operated, cold
stream, thrust reverser.
6. Common Nozzle Assembly (C.N-A.)
Bolted to the rear flange of the turbine exhaust case
7,8. Engine Mounts
Front mount
•
•
attaches to the engine intermediate case with two
brackets and a monoball mount.
It transfers vertical, lateral and thrust loads.
Rear mount
•
•
attaches to the LP turbine exhaust case.
It transfers vertical, lateral and torque loads.
Propulsion Unit General Arrangement
NACELLE
Mechanical Arrangement
Inlet Cowl
The inlet cowl is bolted to the front of the L.P. Compressor
(Fan) case.
Purpose
To supply all the air required by the engine, with minimum
pressure losses and with an even pressure face to the fan.
To minimise nacelle drag.
Construction
Hollow, inner and outer skins supported by front (titanium)
and rear (Graphite / Epoxy composite) bulkheads
Inner and outer skins manufactured from composites.
Leading edge - Aluminium.
Ice Protection
Integral thermal anti-icing utilizing stage 7 air off take.
P2 / T2 Probe
• Probe located at TDC, attached to inner skin
• Senses total inlet pressure (P2), and total inlet
temperature (T2).
• Access panel provided for maintenance.
Handling
• Two hoisting points provided.
• Inlet cowl weighs 238 lbs.
Ventilation Intake
Ram air inlet to provide ventilation of Zone 1.
INLET COWL
NACELLE
Mechanical Arrangement
Inlet Cowl Details
•
Door locators:
Automatically align the fan cowl doors to ensure good
sealing.
•
Strut brackets:
Provide location for the L and R hand fan cowl door
support struts (front struts only).
•
Alignment pins:
Ensure correct location of the inlet cowl to the fan case
when the inlet cowl is installed.
OUTER BARREL
INLET
VENTILATION
P2/T2 PROBE
ACCESS PANEL
DOOR LOCATOR
DOOR LOCATORS
ALIGNMENT PIN
(4 PLACES)
HOiST POINTS
(4 PLACES)
RING
STRUTBRACKET
(2 PLACES)
INNER
BARREL
INTERPHONE JACK
DOOR LOCATOR
TAI DISCHARGE
GRILLE
AFT BULKHEAD
INLET COWL ASSEMBLY - DETAILS
NACELLE
Mechanical Arrangement
P2/T2 Probe Access
An access panel is provided maintenance of the P2/T2
probe
ACCESS PAEL
PYLON CAP
VENTILATION
INTAKE
DOOR
ALIGNMENT
FEATURE
ZONE VENTILATION
OUTLET
P2/T2 PROBE ACCESS
PROBE ELECTRICAL
CONNECTIONS
DOOR ALIGNMENT
FEATURE
PROBE PRESSURE
CONNECTION
P2/T2 PROBE ACCESS PANEL
NACELLE
Mechanical Arrangement
Inlet Cowl Removal/Replacement
Note: Refer to the AMM Task 71-11-11-000-010.
The procedure to remove/replace the Inlet Cowl is
summarised below.
•
Open the L and R fan cowl doors
•
Attach the sling to the inlet cowl and the hoist.
•
Remove the coupling at the anti-ice duct joint and
discard the seal.
Fit now seal on installation.
Disconnect the four electrical connectors at the top
RH side of the cowl aft bulkhead.
Disconnect the P2 signal pipe.
Take the weight of the cowl on the sling with the hoist.
Remove the cowl securing bolts.
Move cowl forward carefully and lower onto dolly.
•
•
•
•
•
Replacement
This is a reversal of the removal procedure. When offering
up the inlet cowl use the 4 location spigots to ensure
correct alignment.
Tests Required
•
•
•
Test 30-21-00-710-010 - TAI system functional
Test 71-00-00-790-010 - Leak test TAI system
Test 73-22-00-710-010 - FADEC Operational test
1
3
2
1 SLING
2 PLASTIC SHIM
3 INLET COWL
INLET COWL - SLIN6
NACELLE
Mechanical Arrangement
Fan Cowl Doors
The doors extend rearwards from the inlet cowl to the
leading edge of the 'C ducts.
Purpose
To provide access to the fan case and gearbox mounted
accessories.
Construction
Graphite skins enclosing, aluminium honeycomb.
Aluminium reinforcement at each corner minimises
handling/impact damage and wear.
Door Fitting
4 hinges on each door locate to the bottom of the pylon.
Doors abut along bottom centre line and are secured to
each other by 4 quick release - adjustable - latches.
Hold-open Struts
Doors can be propped open, for maintenance operations,
using 2 swivel, telescopic struts which are stowed inside
the doors when not in use.
Warnings
The fan cowl hold open struts must be in the extended
position and both struts must always be used to hold
the doors open.
• Be careful when opening the doors in winds of more
than 26 knots (30 mph) .
• The fan cowl doors must not be opened in winds of
more than 52 knots (60 mph).
•
HINGES
DOOR ALIGNMENT
FEATURES
ECC COOLING
INTAKE
HOISTING
POINTS
ACOC
COOLING
OUTLET
QUICK ACCESS
MASTER CHIP
DETECTOR
QUICK ACCESS AIR
STARTER VALVE
AND BLOW OUT DOOR
WHEELCASE
BREATHER
OUTLET
QUICK ACCESS
OIL FILL AND
SIGHT GLASS
BLOW OUT DOOR
STRUTS
LATCHES
FAN COWL DOORS
ADJUSTABLE
KEEPERS
NACELLE
Mechanical Arrangement
Fan Cowl Doors - Removal/Replacement
Refer to the AMM
•
•
Task 71-ll3-ll-000- 010 (Removal)
Task 71-13-11-400-010 (Replacement)
The procedure is summarised below.
1. Remove the blanking caps from the cowl slinging
points.
2. Attach sling to cowl door and heist.
3. Open cowl door to gain access to hinges.
4. Remove split pins from hinge bolts.
5. Remove nuts and shouldered bolts.
6. Remove cowl door and lower onto dolly.
Replacement
This is the reversal of the removal sequence.
On completion, check the cowl door alignment and latch
tension.
PYLON
FAN COWL DOOR
TYPICAL 4 PLACES
1AE-1N20404
BLANK CAP
FAN COWL DOOR-REMOVAL SLING
Mechanical Arrangement
Fan Cowl Doors
Latch Adjustment and Alignment
The mismatch between the two cowl doors can be
adjusted by fitting/removing shims as shown below.
Latch tension is adjusted by use of the adjusting nut at the
back of the latch keeper as shown below.
The latch closing load should be between 45-60 lbs f / in.
SHIM
LEFT HAND FAISI
COWL DOOR
SHIM
RIGHT HAND
FAN COWL DOOR
ALLOWABLE
MISMATCH
VIEW LOOKING FWD
SHIM
HEXAGONAL
WRENCH
SHIM
KEEPER
ASSEMBLY
FAN COWL DOORS - LATCH ADJUSTMENT
ADJUSTING
NUT
NACELLE
Mechanical Arrangement
Inlet Cowl - Handling/Transportation
The arrangement for storage/transportation of the inlet
cowl is shown below.
IAE-1N20401
FAN COWL -TRANSPORT AND WORKSTAND
NACELLE
Mechanical Arrangement
Fan Cowls - Hold Open Struts
When in the open position the fan cowls are
supported by two telescopic hold-open struts, using
anchorage points provided on the fan case (rear) and
inlet cowl (front). Stowage brackets are provided to
securely locate the struts when they are not in use.
The arrangement for stowing and anchoring the hold
open struts are shown below.
ANCHORAGE BRACKET
(TYPICAL 2 PLACES)
STOW
BRACKET
STOW
BRACKET
ATTACH POINT
BRACKET
ATTACH POINT
BRACKET
LATCH HANDLE
{TYPICAL 4 PLACES)
RH FAN COWL DO IR-HOLD-OPEN ROD BRACKETS
NACELLE
Mechanical Arrangement
Fan Cowl Doors Detail
The L.H. fan cowl door detail is shown below.
HINGES (4
PLACES)
DOOR
ALIGNMENT
FEATURES
HOISTING FEATURE
(TYPICAL 2 PLACES)
DOOR STRUT
ATTACHMENT
DOOR STRUT
ATTACHMENT
OIL FILLER
ACCESS DOOR
MASTER CHIP
DETECTOR
ACCESS DOOR
STRUT
STOWAGE
STRUT
STOWAGE
VENTILATION
GRILLE
ADJUSTABLE
KEEPERS
LH FAN COWL DOOR-HOLD-DETAILS
DOOR
ALIGNMENT
FEATURE
HINGES (4
PLACES)
DOOR
ALIGNMENT
FEATURES
EEC COOLING
AIR DISCHARGE
EEC COOLING
INTAKE
HOISTING
FEATURE
DOOR STRUT
ATTACHMENT
STARTER VALVE
ACCESS/PRESSURE
RELIEF DOOR
ACOC AIR
GEARBOX
OVERBOARD
BREATHER
OUTLET
DOOR
ALIGNMENT
FEATURE
DOOR STRUT
ATTACHMENT
STRUT
STOWAGE
VENTILATION
GRILLE
RH FAN COWL DOOR-HOLD-DETAILS
LATCHES
(4 PLACES)
NACELLE
Nechanical Arrangement
Fan Cowl Doors - Storage/Transportation
The arrangement for storing / transporting the cowl doors is
shown below.
IAE-1N20402
FAN COWL DOOR-HANDLING DOLLY
NACELLE
Mechanical Arrangement
C – Ducts / Thrust Revecser
The C - ducts extend rearwards from the fan cowls to the
Combined Nozzle Assembly (CAN). An overview of the
C – ducts / reverser assembly is shown below.
Purpose
The C - ducts:
•
form the cowling around the core engine (inner barrel)
•
form the fan air duct between the fan case exit and the
entrance to the C.N.A.
•
house the thrust reverser operating mechanism and
cascades
•
form the outer cowling between the fan cowls and
C.N.A.
Construction
Mostly composites but some sections are metallic - mainly
aluminium - e.g. inner barrel, Mocker doors and links.
C - Duct fitting
Each C-duct is located to bottom of the strut by 4 hinge
brackets.
The 'C ducts abut along the bottom cent-re line and are
secured to each other by a series of latches which are
located under a hinged access panel.
Opening/Closing C-Ducts
Access to the core engine, for maintenance operation is
gained by opening the hinged C ducts.
The 'C ducts are supported/locked in the open position by
integral support struts.
Opening is effected through an integral, self contained,
hand pump operated, hydraulic system.
UPPER TRACK AND
HINGE FITTING
PRECOOLER
DUCT
DOOR OPENING ACTUATOR
COOLING DUCT
TRANSLATING
SLEEVE
UPPER
ACTUATOR
ACTUATOR
ACCESS
RIGHT
CDUCT
INNER
BARREL
STOW LOCKOUT
(TYPICAL 2 PLACES)
HOLD OPEN
ROD
STOW LOCKOUT/PIN STOWAGE
(TYPICAL 2 PLACES)
MANUAL C-DUCT
HYDRAULIC LINE
C-DUCT - OVERALL VIEW
NACELLE
Mechanical Arrangement
LH Reverser ( C - Duct)
An overview of the ' C' duct is shown below.
Latch Assemblies
Hold - open struts
Four latch hooks engage with four latch keepers, on the
RH C - duct to secure the two C - ducts together. A further
double latch hook, at the rear, is used to lock the two
translating sleeves together.
Two struts on each C - duct support the ducts in the open
position to facilitate maintenance operations on the core
engine.
Bumpers
5 lower bumpers and 4 upper bumpers absorb the fan air
compressive loads. The bumpers, incorporate adjusters shim packs - to provide rigging adjustments.
Heatshield
The whole of the inner barrel is lined with heat
reflective / insulating material.
Pre-Coolor Ducts
Supply fan air to the pre-cooler of the ECS located in the
aircraft strut.
Note For added safety during maintenance operations and
to support the C - ducts during an engine change a GSE
safety rod is inserted.
Take Up Device
This device is used to pull the two 'C ducts together to
facilitate engagement of the main latch assemblies.
UPPER
BUMPER
(4 PLACES)
HEAT
SHIELD
UPPER
BIFURCATION
DOOR OPENING
ACTUATOR
COOLING DUCT
PRECOOLER
DUCT
TORQUE
RING
LOAD SHARE
FITTING
7TH STAGE
BLEED
C-DUCT OPENING
ACTUATOR
OUTER
V BLADE
ONE PIECE
INNER BARREL/
BIFURCATION
AFT HOLD
OPEN ROD
7TH STAGE
BLEED
HEAT
SHIELD
LOWER
BUMPER
(5 PLACES)
LOWER
BIFURCATION
ACAC INLET
FORWARD
HOLD OPEN
ROD
LH THRUST REVERSER ASSEMBLY
FORWARD BUMPER
AND LATCH
NACELLE
Mechanical Arrangement
RH reverser (C – Duct)
An overview of the RH C - duct is shown below.
Description is same as LH C - duct .
7th & 10th Stage Bleed Ports
Ports in both 'C ducts discharge air from the 7 & 10 stage
compressor handling bleed valves, when open, into the
fan duct.
DOOR OPENING
ACTUATOR
COOLING DUCT
TORQUE
RING.
UPPER
BIFURCATION
HEATSHiELD
PRECOOLER
DUCT
UPPER
BUMPER
(4 PLACES)
7TH STAGE
BLEED
LOAD SHARE
FITTING
10TH STAGE
BLEED
C-DUCT OPENING
ACTUATOR
7THSTAGE
BLEED
AFT LOWER
BUMPER
AFT HOLD OPEN ROD
AOAC
BLEED
ONE PIECE
INNER BARREL/
BIFURCATION
OUTER
'V BLADE
HEATSHIELD
LOWER BUMPER
(5 PLACES)
LOWER
BIFURCATION
FORWARD HOLD
OPEN ROD
RH THRUST REVERSER ASSEMBLY
NACELLE
Mechanical Arrangement
C - Duct Latches
A total of six latches arc used to secure the two C - ducts
to each other.
Access to latch A is through the L. and R. hand fan cowls.
3 latches B are located under a hinged access panel.
Latch C is a double latch assembly but the two latches
must be released / latched individually.
LATCH DETAIL A
LATCH DETAIL B
LATCH DETAIL C
C DUCTS - LATCHES
Mechanical Arrangement
Latch Access Panel and Take Up Device
An access panel is provided to gain access to three of the
C-duct latches and the C-duct take-up device.
The take-up device is a turnbuckle arrangement which is
used to draw the two C-ducts together. This is necessary
to compress the C-duct seals far enough to enable the
latch hooks to engage with the latch keepers.
The take up device is used when closing and opening the
C-ducts.
The take up device must be disengaged and returned to
its stowage bracket, inside the LH C-duct, when not in
use.
CNA
TRANSLATING SLEEVE
DOUBLE LATCH
THRUST REVERSER
C DUCT LATCHES
C DUCT
TAKE UP DEVICE
DOOR LATCH
LATCH ACCESS PANEL
BRACKET
ACCESS DOOR INSTALLATION
HOOK
NACELLE
Mechanical Arrangement
C - Duct Hold Open Struts
Two hold open struts are provided on each C - duct to
support the C-ducts in the open position.
The struts engage with anchorage points located on the
engine as shown below.
When not in use the struts are located in stowage brackets
provided inside the C - duct.
The front strut is a fixed length strut.
The rear strut is a telescopic strut and must be extended
before use.
The arrangement for the LH C - duct is shown below, the
RH C-duct is similar.
Warning
Both struts must always be used to support the 'C ducts in
the open position. The C-ducts weigh approx. 578 lbs
each. Serious injury to personnel working under the
C - ducts can occur if the C-duct is suddenly released.
ROD END
FITTING
C DUCT
ROD
ANCHORAGE
BRACKET
ROD
ANCHORAGE
BRACKET
LH T/R DOOR HOLD-OPEN RODS-INSTALLED
Mechanical Arrangement
C-Ducts - Maintenance
Each C-duct is attached to the aircraft pylon by four
hinges.
The three front attachment points ate provided by beams
located on the bottom of the pylon. The beams are not
rigidly attached to the pylon and this provides a degree of
self alignment when closing the 'C ducts.
The rear hinge point is a solid location on the side of the
pylon.
SLEEVE
BOLT
WASHER
WASHER
BOLT
NUT
WASHER
NUT
NUT
WASHER
(TYPICAL LHAND RH
INSTALLATION)
PYLON
GSE PIN
(AIRBUS
FURNISMFH)
BOLT
BOLT
THRUST REVERSER C-DUCT
Mechanical Arrangement
C-Duct Opening/Closing System
Purpose
To provide a mechanical method of to raising /l owering
the C-ducts facilitate one man operation.
Features
On each 'C duct
•
•
•
•
single acting hydraulic actuator
self sealing/quick release hydraulic connection
rigid and flexible hydraulic hoses
pylon and C duct hydraulic actuator attachment
brackets
Aircraft carried
•
hand operated hydraulic pump
Note
The hydraulic fluid used in the system is Engine
Lubricating Oil.
CAUTION
• WING SLATS MUST BE
RETRACTED AND DEACTIVATED
• ALL 6 LATCHES AND
TAKE UP DEVICES
MUST BE RELEASED
• IF REVERSER IS OE
PLOYED. PYLON FAIRING
MUST BE REMOVED
CAUTION
SEE DECAL ABOVE
BEFORE OPENING
C DUCT
'C DUCT OPEN
POSITION
C DUCT CLOSED
POSITION
DOOR OPENING
ACTUATORS
MANIFOLD/PRESS
RELIEF VALVE
IAE-1N20009
HAND PUMP
FLEX HOSE QUICK
DISCONNECT FITTING
THRUST REVERSER D00R OPENING SYSTEM - OVERVIEW
NACELLE
Mechanical Arrangement
C-Duct Opening Actuator
The installation of the RH C-duct actuator is shown below.
LH actuator is similar.
The actuator hydraulic supply tube is continuously cooled
by air taken from the fan duct. The arrangement of the
cooling jacket is shown below.
The actuator has an integral, one way, check valve. This
restricts the fluid return when the C-duct is closing and
thus controls the speed at which the C-duct closes.
Note The C-duct closes owing to its own weight, thus it is
not necessary to use the hand pump to pressurise the
actuator during the closing operations.
HINGE
BEAM
PYLON
CDUCT OPENING
ACT UAT OR
A C T UA T O R
CHECK VALVE
CLAMP
HYDRAULIC
TUBE
HOSE COVER
(EXPANDED)
TUBING AND
HOSES
RELIEF VALVE
WITH I N T E G R A L
FILTER
FILTER
VENT
MANIFOLD
QUICK
DISCONNECT FITTING
RH T/R - DOOR OPENING ACTUATOR COOLIN6 JACKETS - INSTALLATION
NACELLE
Mechanical Arrangement
Combined Nozzle Assembly (CAN)
The CNA is bolted to the rear flange of the turbine exhaust
casing.
Purpose
•
forms the exhaust unit
•
mixes the hot and cold gas streams and ejects the
combined flow to atmosphere through a single
propelling nozzle.
•
completes the engine nacelle
EXIT NOZZLE
INCONEL 625
BRAZED SANDWICH
FAIRING/UPPER STRUT
TITANIUM SKIN AND
FRAME
INNER ANNULUS
LUMINUM SKIN
TITANIUM FRAME
TYPICAL
LOWER STRUT
TITANIUM SKIN
COMMON NOZZLE ASSEMBLY - INSTALLATION
NACELLE
Mechanical Arrangement
C-Duct – Maintenance
Slinging & Hoisting
The slinging and hoisting arrangements are shown below.
HOIST
IAE:1N20002
HINGE ACCESS
PANEL
SCREW
(11 PLACES
SLING - TRUST REVERSER HALF-REMOVAL/INSTLLATION
NACELLE
Mechanical Arrangement
CNA Handling
The arrangements for slinging, hoisting, stowage and
transportation are shown below.
PYLON
IUPPER STRUT
EXHAUST PLUG
(.75 IMCH)(19.5mm)
CLEARANCES
TYPICAL
COMMON NOZZLE
EXHAUST COLLECTOR
NUT
(56 PLACES)
IAE-1N20001
COMMON NOZZLE
FIXTURE
BOLT
(56 PLACES)
IAE-1N20004
CNA DOLLY
COMMON NOZZLE ASSEMBLY-REMOVAL/INSTLLATION
Mechanical Arrangement
C-Ducts - Maintenance
The arrangement for storage/transportation of the
C-duct(s) is shown below.
STRAP
C DUCT
DOLLY AND WORKSTAND
IAE 1N20005 (LH) IAE
1N20006 (RH)
FIGURE 5
DOLLY AND WORKSTAND
Exhaust Plug
The exhaust, plug is located inside the CNA as shown
below.
Removal of the exhaust plug provides access to the cover
plate of the rear bearing compartment (No5 bearing).
TYPICAL
13 PLACES
EXHAUST PLUG - INSTALLATION
NACELLE
Mechanical Arrangement
Engine Mounts
The engine is mounted to the pylon at two places.
Front Mount
Locates to the engine intermediate casing at 3 points,
2 brackets and a Monoball mount.
Located to pylon by 5 bolts aligned by 2 shear pins.
Loads
Transfers vertical, lateral and thrust loads.
CROSS BEAM
MOUNT BEAM ASSEMBY
SHEAR PIN
BEAM SHEAR PIN
(2 places)
MONOBALL STOP
PLATE
THRUST LINK
SHEAR PIN
BEAM JOINING
BOLTS
THRUST LINK
ANTI-ROTATION AND
RETAINING PLATES
(10 places)
FORWARD ENGINE MOUNT
Mechanical Arrangement
Engine Mounts
Rear Mount
Locates to the LP turbine exhaust casing.
Transfers vertical, lateral and torque loads.
Located to the pylon by 4 bolts aligned by 2 shear pins.
BEAM ASSY
JOINING BOLTS
MOUNT BEAM ASSY
SHEAR PIN
(2 places)
LINK ASSY
ANTI-ROTATION AND
RETAINING PLATES
TURBINE EXIT CASE
RETAINING
PLATE
SLEEVF
LINK SHEAR PIN
SOLID PIN AND
SLEEVE
AFT ENGINE MOUNT
V2500 NACELLE
PART TWO - SECTION 3
THRUST REVERSER
NACELLE
Thrust Reverser Unit (T.R.U.)
Introduction
Purpose
The thrust reverser provides deceleration forces to slow
the aircraft on landing or during an abandoned take off. It
is incorporated in the 'C ducts and forms an integral part of
the fan stream exhaust duct.
The thrust reverser comprises a fixed inner and a movable
outer (translating} assembly.
Controls
Selection of reverse thrust, and control of engine power in
reverse thrust is effected by the gated throttle lever {thrust
lever). All signals to and from the T.R.U. are through the
E.E.C. and the Engine Interface Unit (E.I.U.).
Features
•
electrical control circuit
•
hydraulic actuation system
•
positional information feedback system
•
actuator lock position sensors and feed back
•
automatic restow system
•
manual deployment / stow for maintenance
•
manual lock out allows aircraft dispatch with
inoperative thrust reverser
An overview of the thrust reverser is shown below.
GRAPHITE
COMPOSITE
CASCADES
UPPER TRACK
AND HINGE
FITTING
PRECOOLER
DUCT
UPPER ACTUATOR
ACTUATOR
ACCESS
TRANSLATING
COWL
ONE PIECE
INNER
BARREL
STOW LOCKOUT
BLOCKER
DOORS
PIN
STOWAGE
LOWER ACTUATOR
REVERSER OVERVIEW
Introduction - Operation
A sectional view through the Nacelle showing the
translating cowl in the stowed (forward thrust) position and
the deployed (reverse thrust) position is shown below.
GRAPHITE PANEL
DYNAROHR ACOUSTIC
PANEL
BLOCKER DOOR
STOWED - FORWARD THRUST
ALUMINIUM
OUTER
PANEL
CASCADE
ACTUATOR
TORQUE RING
CASCADE
AFT RING
ALUMINIUM
BLOCKER
DOOR LINK
ACCESS
DOOR
21 in STROKE
TRANSLATING
SLEEVE
INNER BARREL
ENGINE
VEE GROOVE
DEPLOYED - REVERSE THRUST
THRUST REVERSER OPERATION
Thrust Reverser
Introduction - Controls and Indications
Selection of reverse thrust and control of engine power in
reverse thrust is effected by the normal thrust lever
(throttle lever).
Movement of the thrust lever rearwards is limited to
forward idle by the latches carried on the thrust levers.
When the latches are lifted the thrust lever is allowed to
move further rearwards (maximum movement approx. 12
degrees). This initiates the following sequence of events
which are all controlled by the E.E.C.
•
the thrust reverser begins to deploy, engine power
commanded to idle.
•
when the thrust reverser has deployed approx.
78%, EEC commands engine to accelerate to
reverse power selected by thrust lever position.
•
thrust reverser continues to deploy to fully deployed
position.
Indications
NO INDICATION
• the thrust reverser is fully stowed
• both locks are fully engaged.
REV
• both locks are disengaged and
• the reverser is between fully stowed and fully deployed
i.e. in transit.
REV
• thrust reverser is fully deployed.
FORWARD
REVERSE
UPPER ECAM DISPLAY
THRUST REVERSE
LATCHING LEVERS
THROTTLE CONTROL
LEVERS
REVERSE POWER RANGE
ON QUADRANT
ENGINE PANEL
THRUST REVERSER CONTROL INDICATING
NACELLE
Thrust Reverser - Description
Hydraulic Actuation System
Purpose
To provide the force required to move the translating cowl
during thrust reverser operation.
Features
•
Two linear hydraulic actuators per translating cowl.
•
One non-locking (upper) actuator which incorporates
a Linear Voltage Differential Transformer (LVDT) to
provide actuator positional feedback signals.
•
One locking (lower) actuator which includes a
locking mechanism to hold the reverser in the
stowed position, the locks incorporate sensors which
signal lock position to the EEC.
•
The Hydraulic Control Unit (HCU) which incorporates
the isolation valve, the directional control valve
(DCV) and the pressure switch.
•
Flexible hydraulic hose assemblies which link the
two upper actuators and incorporate the hydraulic
feed tubes (hydraulic T unions).
•
Rigid hydraulic tubes which link the upper and lower
actuators.
•
Actuator synchronisation system which utilizes
flexible synchronising cables running inside the
hydraulic deploy tubes.
•
Manual
drive
system
(which
utilises
the
synchronising cables) used to stow or deploy the
translating cowl for maintenance operations.
TRANSLATING SLEEVE
UPPER ACTUATOR
HYDRAULIC "T"
AIRBUS
FURNISHED
HYDRAULIC
CONTROL UNIT
RETRACT
HYDRAULIC
HOSE ASSY
- EXTEND HYDRAULIC
HOSE/FLEXSHAFT
ASSEMBLY
- EXTEND TUBE
LOWER
LOCKING
ACTUATOR
RETRAC
TUBE
T/R HYDRAULIC SYSTEM ACTUATION
NACELLE
Thrust Reverser - Description
Hydraulic Control Unit
The Hydraulic Control Unit (HCU) is a self contained line
replaceable unit (LRU), providing safe control of the thrust
reverser actuators in response to electrical position
demands from either channel of the EEC.
The HCU comprises the following:•
One Isolation valve, which can be mechanically
latched in the closed position during maintenance.
•
One direction control valve to port fluid to the
actuators in response to stow or deploy demands.
•
switch
to
detect
system
One pressure
pressurisation downstream of the isolation valve.
•
Two dual-coil solenoid valves to control operation of
the isolation and direction control valves in
response to electrical signal from either channel of
the EEC.
•
One filter with clogging indicator.
•
One bleed valve.
Location
The hydraulic control unit is bolted to the bottom of the
engine pylon in the fan case area. Access is gained by
opening the L.H. fan cowl.
RETURN
SUPPLY
FILTER
HYDRAULIC
ISOLATION
VALVE
DEACTIVATING
LEVER
SOLENOIDS
PRESSURE
SWITCH
HCU
PYLON
CLOGGING
INDICATOR
BYPASS
VALVE
DEPLOY
LOCK PIN
STORAGE
BRACKET
HYDRAULIC
CONTROL UNIT
DEACTIVATION
(Lock out)
PIN
LOCKOUT LEVER
POSITION
HYDRAULIC
FILTER
ELECTRICAL
CONNECTORS
UNLOCK LEVER
POSITION
T/R HYDRAULIC CONTROL UNIT
DIRECTION CONTROL
VALVE
NACELLE
Thrust Reverser - Description
Hydraulic Actuators
Four actuators are used for each thrust reverser, two
actuators are used for each translating cowl.
• the lower actuators incorporate an integral lock
mechanism which holds the piston in the fully stowed
position.
• the upper actuators incorporate an integral Linear
Variable Directional Transformer (LVDT) to indicate
piston position, and thus translating cowl position, to
the EEC.
All actuators use hydraulic snubbing at the end of the
deploy stroke to slow down the actuators over the final
part of the deploy stroke. All actuators also incorporate the
necessary deploy stroke mechanical stops.
Upper Actuators
The two upper actuators are identical Line Replaceable
Units (LRU's), and, in conjunction with the two lower
locking actuators, control movement of the fan reverser
translating elements in response to hydraulic inputs from
the hydraulic control unit.
EYE END
MOUNTING GIMBAL
UNION (Stow)
ADAPTER
(Deploy)
UPPER NONLOCKING ACTUATOR
BALL
ADJUSTABLE EYE END
FLOW CONTROL
VALVE
SEALS
SPLIT GUIDE
CLUTCH SHAFT
PIVOT
END CAP
BEARING
RINGS
SEAL
PIN
SCRAPER
ROD GUIDE
THRUST NEEDLE
BEARING
NUT
PISTON
JACKHEAD ASSEMBLY
SCREWSHAFT
PISTON
WORMWHEEL
UNION (Stow)
MOUNTING BUSH
CLUTCH SHAFT
GIMBAL
JACKHEAD ASSEMBLY
SCREWSHAFT
WORMSHAFT
LOCKNUT
LVDT
ADAPTER (Deploy)
END CAP
UPPER NONLOCKING ACTUATOR - DETAIL
Thrust Reverser
Lower Locking Actuators
The two lower locking actuators are identical Line
Replaceable Units (L.R.U's) and, in conjunction with the
two upper actuators, control movement of the fan reverser
translating elements in response to hydraulic inputs from
the hydraulic control unit.
The actuators incorporate an integral lock mechanism to
hold the piston rod when the actuator is in the fully stowed
position. The lock releases on rising hydraulic pressure
when deploy is commanded via the hydraulic control unit.
The lock mechanism incorporates a manual release facility
and proximity switch for electrical lock position feedback to
the EEC.
Thrust Reverser - Description
Lower Locking Actuators - Locking Mechanism
In the stowed position (forward thrust) the stow and deploy
ports are both connected to return pressure. The piston
assembly is positively, mechanically, held in the stow
position by the claws of the tine lock.
On a deploy command from the E.E.C. the H.C.U. ports
hydraulic fluid to both the stow and deploy ports.
As the flow into the actuator rod side is not restricted, the
pressure rise in the rod side of the piston is greater than
that in the head side. The net effect is to initially move the
piston in the stow direction, which relieves the locking load
on the tine lock. As the pressure in the head side
subsequently builds up, the differential area unlock sleeve
moves forward against the lock spring preload to release
the radial restriction to tine lock movement. Thus, the lock
is released, the resultant pressure acting on both sides of
the differential area piston drives the piston and
synchronising nut through the tine lock in the deploy
sense. The release sleeve spring holds the sleeve in its
disengaged position.
Thrust Reverser - Description
Synchronisation System
The system is used to provide inter-actuator
synchronisation using flexible synchronisation cables
inside the deploy hydraulic tubing.
The system comprises:• One T-piece connector, to split the deploy hydraulic
pipe connection from the hydraulic control unit to each
of the upper feedback actuators.
• Two flexible tubes to carry the deploy hydraulics from
the T-piece to the upper actuators and guide the
flexible synchronising shafting running between the two
upper actuators.
• Two rigid tubes to carry the deploy hydraulics between
upper and lower actuators on each side and guide the
flexible synchronising shafting running between the
upper and lower actuators.
• Three flexible shafts with square male end fittings to
interconnect the synchronising mechanism of each
actuator
The two deploy tubes incorporate a telescopic coupling at
one end to permit simple removal and replacement without
disturbing actuator installation.
T-PIECE HOUSING
SLEEVE
FLEXIBLE GUIDE TUBE
O RING SEAL
FLEXIBLE
DRIVE SHAFT
(TOP)
FLEXIBLE
DRIVE SHAFT
(TOP)
SHAFTING GUIDE T-
PIECE HOUSING ASSEMBLY
FLEXIBLE
DRIVE SHAFT
(SIDE)
GUIDE TUBE
FLEXIBLE
DRIVE SHAFT
(SIDE)
GUIDE TUBE
LOCKNUT
R.NGSEAL
END FITTING
FIGURE 2A
END FITTING
SLEEVE ACTUATOR - FLEXSHAFT INSTALLATION
Thrust Reverser
Cascades (Deflector Boxes)
In the reverse thrust mode all the fan air is directed
through the cascades which eject the air in a forward
direction.
A total of 16 cascades are fitted as shown below.
THURST REVERSER CASCADE
Thrust Reverser - Operation
The thrust reverser is shown schematically, below, in the
forward thrust (flight) position.
In this position:•
hydraulic pressure from the aircraft system is
available as far as the control unit only
•
the isolation valve is in the closed position - control
solenoids de-energised
•
both sides of the actuator pistons are ported to
hydraulic return
•
the thrust reverser is maintained in the forward
thrust position by mechanical locks which are an
integral part of the locking (lower) actuators
•
the directional control valve is in the stow position,
control solenoids de-energised
THRUST REVERSER TRU - SCHEMATIC
Thrust Reverser - Operation
Deploy/Stow
Reverse Thrust Selection
Selection of reverse thrust will - via the EEC - provides
signals to:
•
open the isolation valve to allow hydraulic pressure
to the thrust reverser
•
move the directional control valve (DCV) to the
deploy position
Hydraulic pressure is then felt on both sides of the
actuator pistons but, because of the differential piston
areas, the actuators will extend to move the translating
cowls to the reverse thrust position.
Note
the signal from the EEC to the DCV is routed via a relay
which is closed by the Engine Interface Unit (EIU) when
reverse thrust is selected within the permitted operating
envelope.
Re-selection of Forward Thrust
•
moves DCV to stow position
•
the open signal to the isolation valve is maintained
providing hydraulic pressure to the stow side of the
actuator pistons
•
the extend side of the actuator pistons is ported, via
the DCV to hydraulic return
•
the EEC will cancel the open signal to the isolation
valve 5 seconds after the translating cowls reach
the fully stowed position, to ensure full lock
engagement.
THRUST REVERSER OPERATION
NACELLE
Thrust Reverser - Maintenance
Manual Deploy/Stow
The thrust reverser may be deployed/stowed manually for
maintenance - trouble shooting operations.
The procedure is summarised below, the full procedure,
warnings and cautions may be found in the M.M. CH 78.
•
•
open and tag the following circuit breakers for the
appropriate engine T.B.D. by Airbus
open the L and R hand fan cowls
•
move the thrust reverser hydraulic control unit deactivation lever to the de-activated position and
insert lockout pin
•
disengage the locks on the two locking (lower)
actuators - insert pins to ensure locks remain
disengaged - see below
•
position t lie non return valve in the hydraulic return
line to the by-pass position, see next page (deploy
only -not necessary for stow operation)
•
insert 3/8 inch square drive speed brace into
external socket - see below - push to engage drive
and rotate speed brace to extend /retract translating
cowl as required.
Note do not exceed max. indicated torque loading.
THRUST REVERSER MANUAL DEPLOY
NACELLE
Thrust Reverser - Maintenance
Manual Deploy
During manual deploy operations it is necessary to draw
some hydraulic fluid from the aircraft system. This is done
by moving the non return by-pass lever, located in the
hydraulic return line, to the by-pass position as shown
below.
Access to the non-return valve is gained by removing the
pylon access panel.
On completion of the manual deploy operation the by-pass
valve must be re-positioned to the normal position and the
access panel replaced. A baulking feature on the access
panel prevents the panel being closed if the by-pass lever
is in the by-pass position.
TO HYDRAULIC
RESERVOIR
FROM HCU
VIEW ON A
T/R MANUAL DEPLOY
NON RETURN VALVE (By-pass)
Thrust Reverser - Maintenance
De-activation
An inoperative thrust reverser may be locked in the
forward thrust position for flight, as permitted by M.E.L.
requirements).
Method
The procedure is summarised below, the full procedure is
described in the AMM CH 78 Task 78-32-00-040-011.
•
if the thrust reverser is deployed -stowed manually
as previously described
•
install the lock out pin in the de-activation lever of
the hydraulic control unit. Task 78-32-00-041-010.
•
remove the translating cowl de-activation pins (2)
from their stowage and insert them in the deactivation position
Note
When fully inserted in the de-activation position the pins
will protrude approx. 0.8" to provide visual indication of
"lock out".
THRUST REVERSER -DEACTIVATION
Thrust Reverser - Maintenance
Cascades – Removal & Installation
•
The cascades are not all interchangeable.
•
The cascade positions are identified by numbering.
•
No 1 position is the top position on the RH C-duct.
•
numbering is clock-wise when viewed from the rear.
•
The cascades (deflector boxes) are identified by
part No, as shown below.
VIEW LOOKING FORWARD
LH NACELLE
VIEW LOOKING FORWARD
RH NACELLE
# BLANK CASCADE
THRUST REVERSER CASCADE POSITIONS
Thrust Reverser - Maintenance
Cascades - Removal/Installation
The procedure to remove/install a cascade is outlined below.
Reference must be made to the M.M. Task 78-32-19-000010.
•
deploy the thrust reverser, hydraulically or manually
•
de-activate the hydraulic control unit (HCU).
•
remove the cascade (deflector box)
Note
blank deflector boxes have 12 bolts, all others have eight bolts
•
installation of cascade boxes is a . reversal of the
above procedure.
LH AFT
SUPPORT RING
THESE TWO (2)
HOLES ARE 0.10
INCH FURTHER AFT
TO PROVIDE
BULKING FEATURE
THRUST REVERSER CASCADE - INSTALLATION
Thrust Reverser - Electrical Harness
The thrust reverses electrical harness connections
are as shown below.
HYDRAULIC CONTROL UNIT
1 Directional control valve
2 Isolation valve
3 Hydraulic pressure switch
LH LVDT
RH LVDT
LH LOCK SENSOR
RH LOCK SENSOR
RH CDUCT
4005VC (9O0J)
THRUST REVERSER - HARNESS
V2500 GENERAL
PART THREE - SECTION 2
COMPONENT LOCATION GUIDE
HYDRAULIC
PUMP
NOSE CONE
FAN CASE
FUEL COOLED
OIL COOLER
HP COMPRESSOR
SECTION
REAR
ENGINE MOUNT
FUEL
FILTER
OIL
TANK
HYDRAULIC
PUMP
COMBUSTION
SECTION
OIL
PUMP
STAGE 10
BLEED VALVE
COMMON
NOZZLE
FUEL
PUMP
GEARBOX
ENGINE- LEFT HAND SIDE- COMPONENT LOCATION
RELAY BOX
STAGE 7
BLEED VALVES
LP COMPRESSOR
(FAN)
EEC
TURBINE SECTION
No. 4 BEARING
COMPARTMENT AIR
COOLER
STARTER
DE-OILER
AIR COOLED
OIL COOLER
IDG
BLEED VALVE
CONTROL VALVES
ENGINE- RIGHT HAND SIDE- COMPONENT LOCATION
V2500 GENERAL
PART THREE - SECTION 1
TROUBLE SHOOTING
TROUBLESHOOTING
A320
Introduction
Menu Mode
The A3 20 utilises Electronic systems to detect, categorise
and display faults which occur during aircraft operation.
This mode is only available when the aircraft is on the
ground and the engines are not running.
One of these systems is the Centralised Fault Display:
System (CFDS). The CFDS consists basically of;-
It allows maintenance crews to establish an interactive
dialogue through one MCDU with any of' the connected
aircraft systems.
•
all (Bite) Built In Test Equipment portions of. the
electronic control systems
•
the central computer - the Centralised Fault Display
Interface Unit (CFDIU)
•
two Multi-Purpose Control and Display Units
(MCDU)
The CFDS has two operating modes:•
normal mode (or reporting mode)
•
menu mode (or interactive mode)
Normal Mode
The CFDS continuously receives failure and status
information from the Bite portions of all the connected
systems.
The location of the units is shown below.
MULTI PUHPOSE
PRINTER
2 MCDU
TO MCDU's
AND PRINTER
CFDIU
bite
bite
bite
bite
ELECTRONIC
SYSTEM
AVIONICS BAY - 80 VU
A 320 - CFDS - COMPONENTS LOCATION
A320 TROUBLESHOOTING
Centralised Fault Display System (CFDS)
The CFDS is used to provide maintenance information
and can also be used to initiate various functional tests,
e.g. thrust reverser, ignition system, etc.
The diagram below shows how the CFDS interfaces with
the Engine Electronic Control (EEC) and the Electronic
Centralised Aircraft Monitoring System (ECAM).
Engine Systems which are controlled by the EEC are
continuously monitored by fault detection logic circuitry
(BITE), within the EEC.
When the EEC detects a fault it generates fault data which
is then transmitted to three user systems:•
the Flight Warning Computer, which generates
ECAM displays to alert the flight crew and provide
advice on recovery / handling procedures.
•
through the Central Maintenance System (CMS) to
the CFDIU.
•
to its own non volatile memory Maintenance
personnel can interrogate this memory through the
CFDS to obtain more detailed fault data.
Printed copies of all fault data can be obtained from the
EEC memory, using the CFDS printer.
ATA Ref.'
Language Message
MCDU
Prinloul of
In-Hight messages
761100
EGAM
ENG W.D.
TRA sense/HC/EEC
Ptinlout of EEC
fault memory
CFDIU
Maintenance crew
can use CMS menu
mode to obtain
more Information
Flight Warning Compuler
ARINC output bus
CMS logic
Faull delectfon logic
Memory
EEC
Faulted componanl
E.G. TRA crosscheclt
FAULT MESSAGE ANNUNCITION ECAM/CFDS
A3 20
CFDS Failure Classification
Three classes of failures have been defined, according to:• how critical the failure is
• its operational impact
• the scheduled maintenance policy
The failures are categorised as:Class 1
These are failures which have
consequence on the flight in progress.
an
operational
Class 2
These are failures which do not have an operational
consequence on the flight in progress but may affect
subsequent flights - refer to Minimum Equipment List
(MEL) for functions lost.
Class 3
Failures which have no operational effect and can be left
until next scheduled maintenance check.
The failure classification is summarized below.
Note: The EEC does not transmit class three messages.
1
2
Operational consequences
on the current flight
YES
NO
NO
Indicated to the Pilots
YES
YES
NO
Failure
Classes
Dispatch consequences
Maintenance information
Warnings / Flags
System pages
REFER TO MMEL
may be
"GO" "GO IF" "NO GO"
On the system Display
"STATUS"page
FUNCTIONS LOST
INDICATED IN MMEL
"GO" without conditions
HAVE TO BE REPORTED BY THE PILOTS IN
THE LOG BOOK
INDICATED AT THE END OF EACH FLIGHT
LEG : "LAST LEG REPORT
FAILURE
CLASSIFICATION
3
NO REFERENCE
IN MMEL
AVAILABLE ON REQUEST
CAN BE LEFT UNCORRECTED
UNTIL THE NEXT ADEQUATE
MAINTENANCE OPPORTUNITY
A320 TROUBLESHOOTING
CFDS Normal Mode - Health Monitoring
Wraparound Checks
When operating in the Normal mode the Bite portion of the
EEC continuously monitors the operation of the systems,
under its control, by carrying out:-
This checks out the system electrical circuitry, detects
faults such as - loose connectors, chaffed harness, bent
pins, high resistance, broken cable etc.
1. track checks
2. cross checks
3. wraparound checks
Track Check
The EEC compares the commanded position (output
signal) and the actual position (feedback signal) of the
output device. An error between these two signals shows
that the output device has not gone to the commanded
position. If the error is in one channel only ;the fault is
most likely in the output device, if the fault is in both
channels it indicates either a failure or a mechanical
problem, fouling, bent control rod etc*
Cross Check
The ESC compares the outputs of the Channel A and
Channel B feedback devices (LVDT's, microswitches). A
difference indicates and internal unit fault.
Fault Indications
A fault detected by the Bite generates two messages:•
a clear language message displayed on CFDS
•
a unique alphanumeric fault code
CHANNEL A
DATA
LINK
CHANNEL B
CONDITlON (FEEDBACK) SIGNAL
EEC
1. TRACK CHECK
√ or X
2. CROSS CHECK
√ or X
3. WSAPAROUND
CHECK
√ or X
LVDT
CONTROL SIGNAL
CFDS- HEALTH MONITORING
A320 TROUBLESHOOTING
CFDS - Menu Mode
Detailed information about failures can be obtained from
the CFDS, operating in the Menu mode, through either of
the MCDU's.
Menus are displayed on the MCDU. Selection of the
appropriate item is made by pushing the keys located
alongside the display.
The CFDS menu mode functions are:Last Leg Report:
Displays all the class 1 and class 2 failures of the last leg,
up to a maximum of 20.
Last Leg ECAM Report:
Displays the ECAM warnings experienced during the last
leg - up to a maximum of 20.
Previous Leg Report
This is a copy of the 63 previous legs
maximum capacity of 200 failures, (whichever is first).
Avionics Status
Provides a real time display of all systems affected by
internal or external failures.
Post Flight Print
Prints a copy of the Last Leg Report plus the Last Leg
ECAM Report.
System Report/Test
Allows maintenance personnel to interact with the chosen
system, through the CFDIU.
CFDS- MENU MODE
A3 20 TROUBLESHOOTING
Fault Diagnosis
The following pages describe the use of the CFDS to
confirm and diagnose faults.
For example, a fault message which reads:ENG 1 COMPRESSOR VANE
has appeared on the ECAM during flight and has been
written in the tech. log by the pilot.
Several faults could generate the same message.
By utilising the CFDS the specific problem can be
identified and further fault data, to assist in fault diagnosis,
can be obtained.
The troubleshooting sequence is as follows:Gain access to the flight deck, then:
1 turn on the FADEC power using the switch on the
overhead panel - RH side.
2 press MCDU menu key
3 select CFDS
4 select POST FLIGHT PRINT
The post flight report, see the example below, records the
ECAM messages and the fault messages. The fault
message is identified by cross checking the times, and is a
clear language message - in this case
2.5 BLEED ACT/HC/EECl
This tells us that the specific fault is in the 2.5 (BSBV)
bleed system.
POST FLIGHT REPORT
A320 TROUBLESHOOTING
Fault Diagnosis (Coiit)
The clear language message obtained from the post flight
report provides the basis for further troubleshooting.
At this point reference is
Troubleshooting Manual (TSM).
made
to
the
A320
An extract from the TSM is shown below. When using the
TSM use the clear language CFDS message to locate the
correct page.
The TSM presents 3 options for the cause of this particular
fault:•
•
•
the 2.5 bleed actuator
a harness or connector fault
the EEC.
Additional data, to identify which option to take, can be
obtained from the CFDS.
This data is in the form of an alpha / numeric fault code
which comprises two letters/digits e.g. D9, 35, 3C or 5F.
The procedure used to obtain the fault code is detailed,
step by step, in the following pages.
A 320 TSM EXTRACT
A320 TROUBLESHOOTING
Fault Codes
To obtain a fault code for a CFDS fault indication proceed
as follows:Gain access to the flight deck, then:1. turn on the FADEC power using the switch on the
overhead panel RH side
2. Press MCDU menu key
3. select CFDS
4. select SYSTEMREPORT/TEST
5. press NEXT PAGE key on MCDU
6. select ENG
7. select FADEC I A
Note:
Always interrogate FADEC A and B for the appropriate
engine
8. select LAST LEG REPORT
The sequence is shown below and is continued on next
page.
CFDS- FAULT CODE ACCESS
A320 TROUBLESHOOTING
Fault Codes (Cant)
9. an example of a LAST LEG REPORT is shown below.
Look for the clear language message which was shown
on the CFDS i.e.
2.5 BLD ACT/HC/EECl
identify, and note, the fault cell number (30)
10. press return key to go back to FADEC IA menu
11. (cont on next page)
FAULT CELL-IDENTIFICATION
A320 TROUBLESHOOTING
Fault Codes (Cont)
11 select TROUBLESHOOTING
12 select FLIGHT DATA
13 look for a fault cell which has the same number as the
fault in question (found in step 9) i.e. CELL 30
NB there may be more than one page, if the fault cell
number does not appear, use the NEXT PAGE key on
the MCDU to move through all the fault cells
14 identify word Number 1, as shown below, and note the
two right hand letter / digit(s) i.e. 3C
Note On EECS with a software code prior to SCN11E
(approx January 1991) word six was used to obtain this
data.
FAULT CODE IDENTIFICATION
A320 TROUBLESHOOTING
Fault Code Identification
The fault code 3C
troubleshooting data.
is
used
to
provide
further
Reference to fault code lists shows that the fault code 3C
means that the 2.5 bleed actuator has failed the track
check.
This indicates a mechanical fault, the actuator has not
gone to the commanded position.
An examination of the 2.5 bleed actuator may reveal an
actuator foul or mechanical damage, if this is not the case,
change the 2.5 bleed master actuator.
Note
•
Three other faults could have produced the same
clear language message, these fault codes are:-
•
5F 2.5 bleed cross check failed. To rectify this fault
change the 2.5 bleed valve master actuator
•
35 2.5 bleed torque motor wrap around failure usually indicates a harness or connector problem
•
D9 Local 2.5 LVDT latched failed change actuator
FAULT CODE
A3 20 TROUBLESHOOTING
Harness / Electrical Faults
These faults usually cause the system to fail the
Wraparound check.
Reference to the troubleshooting manual will detail the
investigation (method and sequence) to be carried out.
This can be seen below.
In our example the fault code was 3C.
The TSM details the specific harness checks to be carried
out for this fault code and also provides the locator
reference for the Aircraft Schematic Manual (ASM>, i.e.
ASM 73-25-00
The ASM provides a schematic circuit diagram which
identifies and locates the connectors, cables, etc for this
circuit.
The Aircraft Maintenance Manual (AMM) locator reference
is also quoted i.e. AMM 71-50-00, This details the specific
harness checks to be carried out i.e. continuity checks,
visual checks, resistance checks etc and also explains
how these should be accomplished.
Should any defective circuits or damaged parts be found
the AMM will provide the locator reference for the Aircraft
Wiring Manual (AWM) i.e. 20-71-00.
The AWM provides explicit instructions on how to carry out
the necessary repairs/replacements.
DATA PATH - HARNESS FAULT
A3 20 TROUBLESHOOTING
FADEC Test
The FADEC system self test if carried out from the flight
deck through the MCDU using the CFDS in the MENU
mode.
The procedure is shown, step by step, below.
With regard to the FADEC self test there are two points
which must be clearly understood:1. if fault rectification has involved a component change
or a repair the FADEC self test alone does not satisfy
the test requirements. Additional test, some of which
involve an engine ground run, may also be required.
This is explained in more detail on the next page - see
Engine Testing After Fault Rectification.
2. if the FADEC self test shows "Test Failed", it does not
mean the fault is within the EEC. For example, part of
the FADEC self test are wraparound checks, and a
harness fault, a loose plug, for instance, could register
a "Test Failed" indication.
FADIC SELF TEST
A320 TROUBLESHOOTING
Engine Testing After Fault Rectification
On completion of Power Plant, Module, Component,
repairs or replacement procedures some further tests may
be required before the aircraft is returned to service.
The tests required after specific actions are detailed in the
Aircraft Maintenance Manual (AMM) in:CH 71-00-00 Page Block 200
Maintenance Practices
as shown below.
In our example the 2.5 Bleed, master actuator has been
changed.
Reference to AMM 71-00-00 PB 200 shows that tests 3 or
1, and 11 are required.
test 1 : Dry motor leak check.
test 3 : Idle leak check.
test 11: High Power assurance test ()
The test procedures are also detailed under the same
AMM reference. An extract from the AMM is shown below
and on the next page.
AMM EXTRACT (I)
AMM EXTRACT (2)
V2500 NACELLE
PART TWO - SECTION 4
NACELLE
VENTILATION / OVERHEAT &
FIRE PROTECTION
NACELLE
Nacelle Ventilation
Ventilation is provided for the fan compartment (Zone 1),
and the core compartment (Zone 2), to:
•
prevent accessory and component overheating
•
prevent the accumulation of flammable vapours
Zone 1 Ventilation
Ram air enters the zone through an inlet located on the
upper LH side of the air intake cowl. The air circulates
through the fan compartment and exits at the exhaust
located on the bottom rear centre line of the fan cowl doors.
Zone 2 Ventilation
The ventilation of Zone 2 is provided by air exhausting from
the Active Clearance Control (ACC.) system around the
turbine area.
The air circulates through the core compartment and exits
through the lower bifurcation of the 'C ducts.
Ventilation during Ground Running
During ground running local pockets of natural convection
exist providing some ventilation of the fancase - Zone 2.
Zone 2 ventilation is still effected in the same way as when
the engine is running.
PNEUMATIC
DUCT LEAKS
FAN CASE COMPARTMENT
VENTILATION EXIT
VENTILATION
INL£T
PRESSURE RELIEF
DOOR (TYPICAL 2
PLACES)
THRUST REVERSER
SEAL LEAKS
ZONE 1
ZONE 2
CORE ENGINE
COMPARTMENT
PRESSURE
RELIEF DOOR
VENTILATION
EXIT
FAN CASE AND CORE COMPARTMENT VENTILATION
NACELLE
Fire Detection System
The fire detection system monitors the air temperature in
Zone 1 and Zone 2. When the air temperature increases
to a predetermined level the system provides flight deck
warning by:
• master warning light
• audible warning
• specific fire indications
Zone 1 and Zone 2 fire detectors function independently of
each other.
Each zone has two detector units which are mounted as a
pair, each unit gives an output signal when a fire or
overheat condition occurs. The two detector units are
attached to support tubes by clips.
ENGINE FIRE DETECTION SYSTEM - OVERALL VIEW
NACELLE
Fire Detection System
Detector Units
Responder
Each detector unit comprises:-
The responder has two pressure switches, one normally
open, the other normally closed.
• a hollow sensor tube
• a responder assembly
Sensor Tube
The sensor tube is closed and sealed at one end, the
other open end is connected to the responder. The tube is
filled with Helium and carries a central core of ceramic
material impregnated with Hydrogen.
An increase in the air temperature around the sensor tube
causes the Helium to expand and increase the pressure
within the tubes, further increases in temperature cause
the core material to expel Hydrogen to increase the
pressure within the tube.
Both pressure switches sense the gas pressure in the
sensor tube. The responder is connected via its electrical
receptacle and the harness to an electronic fire detectLon
module.
ELECTRICAL
CONNECTOR
NORMAL OPEN
SWITCH
RESPONDER
CORE ELEMENT
CERAMIC IMPREGNATED
WITH HYDROGEN
SENSOR TUBE
NORMALLY CLOSED
SWITCH
TEST
AIRCRAFT
ELECTRICAL
SYSTEM
28V DC
FIRE DETECTOR UNIT - SCHEMATIC
GAS-HELIUM
NACELLE
Fire Detection System
Sensor Setting Limits
The operating temperature limits for Zone 1 and Zone 2
are shown below.
ACCESSORY ZONE
CORE ZONE
12 INCHES
IMMERSED
371 ± 55° C
620 ± 55° C
FULLY
IMMERSED
235 ± 14° C
370 ± 22° C
MARGIN
MARGIN
125° C
ZONE MAXIMUM TEMPERATURE
260° C
ZONE MAXIMUM TEMPERATURE
FIRE DETECTOR UNIT - SETTING RANGE
NACELLE
Fire Detection System
Indications and Controls
The fire indications and controls are located as shown
below.
FIRE PROTECTION-INDICATIONS/CONTROLS
NACELLE
Fire Detection and Nacelle
Temperature Monitoring
Electrical Harness
The fire detection nacelle temperature indication electrical
harness connections are as shown below.
NACELLE TEMPERATURE SENSING
AND FIRE DETECTION HARNESS
V2S00 NACELLE
PART TWO - SECTION 5
ENGINE
REMOVAL/INSTALLATION
EngIne Change
The arrangements for slinging/hoisting the engine are
shown below.
NOTE
During this operation the C-ducts are supported by (GSE)
Ground Service Equipment rods which are positioned
between the C-duct and the aircraft pylon.
ENGINE REMOVAL AND INSTALLATION
V2500 GENERAL
PART THREE - SECTION 4
BORESCOPING
LP COMPRESSOR
OUTLET GUIDE VANE
STAGE 1
FAN BLADE
FAN FRAME
STRUT
FLEXIBLE BORESCOF
FAN OUTLET
INNER VANE
LP COMPRESSOR
STAGE 1,5 BLADE
Examine the Front Surfaces of the Stage ? LP Compressor Blades Fig
602/TASK 72-00-00-991-156
INSPECTION/CHECK
R EFFECTIVITY: V2500
72-00-00
Page 612
INLET GUIDE—
VANE
LP COMPRESSOR
STAGE 2.5 BLADE
FAN FRAME
STRUT
BORESCOPE FLEXIBLE
IP Compressor Stage 2.5 Blades - Inspection/Check
Fig 603/TASK 72-O0-OO-991-157
INSPECTION/CHECK
R EFFECTIVITY: V2500
72-00-00
Page 613
INTERNATIONAL AERO ENGINES
V2500 Propulsion System
ENGINE MAINTENANCE MANUAL