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A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
DIFFERENCES V2500 TO
CFM56-5B
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 1
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
TDTI / HAT / ATA 71-80
Austrian Technical Training
School Notes - For Training Purposes Only
Issue: 06/08
Revision: 18.06.2008
Page 2
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
This document must be used for training purpose only.
Under no circumstances should this document be used as a reference.
It will not be updated.
All rights reserved.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 3
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
THIS PAGE INTENTIONALLY LEFT BLANK
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 4
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
DIFFERENCES V2500 TO CFM56-5B ................................ 1
GENERAL........................................................................ 10
INTRODUCTION ............................................................................ 10
POWER PLANT LEVEL 2.................................................................. 14
POWERPLANT DESCRIPTION .......................................................... 16
ENGINE SYSTEM D/O (3)............................................... 20
ENGINE
ENGINE
ENGINE
ENGINE
CHARACTERISTICS ........................................................... 20
GENERAL PARAMETERS ..................................................... 22
CONTROL P/B'S AND SWITCHES ........................................ 24
INDICATING PRESENTATION............................................. 26
BORESCOPE PORTS .......................................................................88
ENGINE FUEL SYSTEM D/O (3)...................................... 94
GENERAL .......................................................................................94
FUEL FEED ....................................................................................94
METERED FUEL ..............................................................................94
SERVO FUEL ..................................................................................94
DIVERTED FUEL.............................................................................94
EEC CONTROL ...............................................................................96
INDICATING ..................................................................................98
FUEL DISTRIBUTION COMPONENTS ............................................. 100
HEAT MANAGEMENT SYSTEM D/O (3) ......................... 124
POWER PLANT DRAIN PRESENTATION (3) ................... 28
GENERAL ..................................................................................... 124
COMPONENTS ............................................................................. 124
CONTROL .................................................................................... 126
POWER PLANT INSTALLATION D/O (3) ........................ 34
GENERAL ..................................................................................... 134
COMPRESSOR AIRFLOW CONTROL ............................................... 136
TURBINE CLEARANCE CONTROL AND TURBINE COOLING ............. 148
N°4 BEARING COMPARTMENT COOLING ....................................... 152
NACELLE VENTILATION................................................................ 154
GENERAL ...................................................................................... 28
PYLON DRAINS .............................................................................. 30
ENGINE DRAINS ............................................................................ 32
AIR INTAKE COWL ......................................................................... 34
FAN COWL DOORS (LH & RH) ........................................................ 36
THRUST REVERSER "C" DUCTS ...................................................... 38
FIREWALLS AND ACOUSTIC PANELS ............................................... 40
COMMON NOZZLE ASSEMBLY ......................................................... 42
EXHAUST CONE ............................................................................. 42
FWD MOUNT ................................................................................. 44
AFT MOUNT .................................................................................. 44
FLUID DISCONNECT PANEL ............................................................ 46
FAN ELECTRICAL CONNECTOR PANEL ............................................ 48
CORE ELECTRICAL JUNCTION BOX ................................................. 50
AERODYNAMIC STATIONS ............................................................. 52
ENGINE BEARINGS ........................................................................ 54
ENGINE SEALS .............................................................................. 62
COMPRESSOR................................................................................ 64
COMBUSTION SECTION ................................................................. 74
TURBINE SECTION ........................................................................ 78
TDTI / HAT / ATA 71-80
Issue: 06/08
AIR SYSTEM PRESENTATION (3) ................................. 134
FADEC PRESENTATION (3) .......................................... 158
PURPOSE..................................................................................... 158
FADEC FUNCTIONS ...................................................................... 158
FADEC BENEFITS ......................................................................... 158
POWER SUPPLY ........................................................................... 158
FADEC ARCHITECTURE (3) .......................................... 160
DUAL CHANNEL ........................................................................... 160
DUAL INPUTS .............................................................................. 160
HARDWIRED INPUTS ................................................................... 160
DUAL OUTPUTS ........................................................................... 160
BITE CAPABILITY ......................................................................... 160
FAULT STRATEGY ........................................................................ 160
FAIL-SAFE CONTROL.................................................................... 160
Revision: 18.06.2008
Page 5
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
MAIN INTERFACES ...................................................................... 166
FADEC PRINCIPLE (3) ................................................. 168
GENERAL .................................................................................... 168
FADEC ........................................................................................ 168
EIU ............................................................................................. 168
POWER MANAGEMENT ................................................................. 168
ENGINE LIMITS ........................................................................... 168
ENGINE SYSTEMS ........................................................................ 168
IGNITION AND STARTING ............................................................ 168
THRUST REVERSER ..................................................................... 170
EEC INTERFACES (3) ................................................... 172
GENERAL .................................................................................... 172
DIGITAL INPUTS ......................................................................... 172
DIGITAL OUTPUTS ...................................................................... 172
DISCRETE/ANALOG SIGNALS ....................................................... 172
BLOCK DIAGRAM ......................................................................... 174
EIU INTERFACES (3) .................................................... 184
GENERAL .................................................................................... 184
EIU COMPOSITION ...................................................................... 184
EIU FUNCTIONS .......................................................................... 184
INPUTS ....................................................................................... 184
OUTPUTS .................................................................................... 184
EEC INTERFACE ........................................................................... 184
ECS INTERFACE ........................................................................... 186
CFDS INTERFACE......................................................................... 186
BMC INTERFACE .......................................................................... 186
OTHER INTERFACES .................................................................... 186
POWER SUPPLY ........................................................................... 186
FMGS .......................................................................................... 188
EIU CFDS DISCRETE OUTPUTS SIMULATION ................................ 190
ENGINE THRUST MANAGEMENT (3) ............................ 196
BASIC INFORMATION .................................................................. 196
AUTOTHRUST CONTROL MODE .................................................... 210
TDTI / HAT / ATA 71-80
Issue: 06/08
MANUAL CONTROL MODE ............................................................ 214
BACK-UP N1 MODE ...................................................................... 216
ENGINE CONTROLS (3) ............................................... 222
THROTTLE CONTROL SYSTEM ...................................................... 222
OIL SYSTEM D/O (3) ................................................... 228
GENERAL ..................................................................................... 228
OIL SUPPLY CIRCUIT ................................................................... 229
OIL SCAVENGE CIRCUIT .............................................................. 230
OIL INDICATING SYSTEM............................................................. 232
VENT CIRCUIT ............................................................................. 234
EEC ELECTRICAL PWR SPLY CONTROL (3) .................. 238
GENERAL ..................................................................................... 238
POWERING N2 < 10% ................................................................. 238
POWERING N2 > 10% ................................................................. 238
AUTO DEPOWERING .................................................................... 238
MANUAL REPOWERING ................................................................ 238
IGNITION & STARTING SYSTEM PRESENTATION (3) . 240
GENERAL ..................................................................................... 240
CONTROL AND INDICATING ......................................................... 240
AUTOMATIC START ..................................................................... 240
MANUAL START ........................................................................... 240
CRANKING ................................................................................... 240
CONTINUOUS IGNITION .............................................................. 240
SAFETY PRECAUTIONS................................................................. 240
STARTING COMPONENTS ............................................................. 242
MAINTENANCE PRACTICES........................................................... 242
IGNITION & STARTING SYSTEM D/O (3) .................... 244
GENERAL ..................................................................................... 244
AUTO START ............................................................................... 246
MANUAL START ........................................................................... 260
ENGINE CRANK ............................................................................ 278
ENGINE INDICATION/ MONITORING D/O (3) ............ 286
POWER INDICATING .................................................................... 288
Revision: 18.06.2008
Page 6
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
TEMPERATURE ............................................................................ 294
POWER ....................................................................................... 298
DEDICATED ALTERNATOR (PMA) ................................................. 304
ANALYZERS ................................................................................. 306
ENGINE WARNINGS (3)............................................... 312
THRUST REVERSER SYSTEM PRESENTATION (3) ........ 316
REVERSER DESIGN ...................................................................... 316
HYDRAULIC SUPPLY .................................................................... 318
ACTUATION ................................................................................ 318
REVERSER CONTROL ................................................................... 318
REVERSER INDICATING ............................................................... 318
MAINTENANCE PRACTICES .......................................................... 320
THRUST REVERSER MANAGEMENT (3) ........................ 322
GENERAL .................................................................................... 322
THRUST REVERSER ACTUATION .................................................. 324
THRUST REVERSER CONTROL ...................................................... 324
THRUST REVERSER INDICATION .................................................. 328
CFDS INTERFACE......................................................................... 328
INITIAL CONDITIONS .................................................................. 330
DEPLOY SEQUENCE ..................................................................... 332
STOW SEQUENCE ........................................................................ 334
COMMAND LIMITATION ............................................................... 338
FAN AND THRUST REVERSER COWL DOORS CLOSING .................. 376
THRUST REVERSER COWL DOORS CLOSING ................................. 376
FAN COWL DOORS CLOSING ........................................................ 378
THRUST REVERSER DEACTIVATION AND LOCKOUT ...................... 380
PRECAUTIONS ............................................................................. 382
MANUAL DEPLOYMENT AND STOWAGE PROCEDURE ..................... 384
PRECAUTIONS ............................................................................. 386
ENGINE REMOVAL AND INSTALLATION (3) ................ 388
BOOTSTRAP SYSTEM INSTALLATION ............................................ 388
ENGINE TRANSPORTATION STAND ATTACHMENT POINTS ............ 392
ENGINE REMOVAL ....................................................................... 394
ENGINE INSTALLATION................................................................ 396
FAN COWL DOOR REMOVAL/INSTALLATION ................................. 398
THRUST REVERSER COWL DOOR REMOVAL/INSTALLATION .......... 402
COMMON NOZZLE ASSEMBLY REMOVAL/INSTALLATION ................ 406
AIR INTAKE COWL - REMOVAL/INSTALLATION ............................. 408
FAN BLADE - REMOVAL/INSTALLATION ........................................ 410
FAN MODULE - REMOVAL/INSTALLATION ..................................... 412
ENGINE BORESCOPE - INSPECTION/CHECK .................................. 414
AIR STARTER - REMOVAL/INSTALLATION ..................................... 416
IDG - REMOVAL/INSTALLATION ................................................... 418
IDG - SERVICING ......................................................................... 420
SERVICING, MAINTENANCE AND MEL ITEMS ............. 340
ENGINE OIL SERVICING ............................................................... 340
MASTER CHIP DETECTOR CHECK ................................................. 342
MEL / DEACTIVATION .................................................................. 346
MAINTENANCE TIPS .................................................................... 354
ENVIRONMENTAL PRECAUTIONS ................................................. 356
OPENING & CLOSING OF ENGINE COWL DOORS (3) .. 358
FAN AND THRUST REVERSER COWL DOORS OPENING .................. 358
FAN COWL DOORS OPENING ....................................................... 358
THRUST REVERSER COWL DOORS ............................................... 362
THRUST REVERSER COWL DOORS OPENING ................................ 374
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 7
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
TDTI / HAT / ATA 71-80
Austrian Technical Training
School Notes - For Training Purposes Only
Issue: 06/08
Revision: 18.06.2008
Page 8
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
THIS PAGE INTENTIONALLY LEFT BLANK
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 9
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
GENERAL
INTRODUCTION
It is produced by International Aero Engines (IAE) corporation. This
corporation consists of the following companies:
ƒ
ƒ
ƒ
ƒ
ƒ
JAEC (Japanese Aero Engines Corporation)
Rolls Royce
Pratt & Whittney
MTU ( Motoren & Turbinen Union )
Fiat Avio
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 10
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
CORPORATION MEMBERS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 11
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
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.
ƒ
ƒ
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. For example: The V2500
– A1 engine is used on A320 and has only a 3 stage booster.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 12
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE MARK NUMBERS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 13
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
POWER PLANT LEVEL 2
SYSTEM OVERVIEW
The IAE V2500-A5 engine is a two spool, axial flow, high bypass ratio
turbo fan power engine. The V2500-A5 powers the complete single
aisle family of aircraft except the A318. V2500-A5 engines are available
in several thrust ratings. All the engines are basically the same. A
programming plug on the Electronic Engine Control (EEC) changes the
available thrust. The power plant installation includes the engine, the
engine inlet, the exhaust, the fan cowls and the reverser assemblies.
The pylon connects the engine to the wing structure. The engine is
attached to the pylon by FWD and AFT mounts.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 14
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
SYSTEM OVERVIEW
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 15
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
POWERPLANT DESCRIPTION
The V2530-A5 engine is a two spool, axial flow, high bypass ratio
turbofan engine.
•
80% of the thrust is produced by the fan.
•
20% of thrust is produced by the engine core.
Its compression system features a single stage fan, a four-stage
booster, and a ten-stage high pressure compressor. The LP
compressor is driven by a five-stage low pressure turbine and the HP
compressor by a two stage HP turbine. The HP turbine also drives a
gearbox, which in turn, drives the engine and aircraft mounted
accessories. The two shafts are supported by five main bearings. The
V2500 incorporates a full authority digital Electronic Engine Control
(EEC). The control system governs all engine functions, including
power management. Reverse thrust is obtained by deflecting the fan
airstream via a hydraulic operated thrust reverser.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 16
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
POWERPLANT DESCRIPTION
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 17
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
THRUST REVERSER SYSTEM (2)
Reverse thrust is controlled by the EEC. Reverse is manually selected
by the flight crew by lifting the latching levers on the throttle control
levers. The reverse thrust command is sent to the EEC and the EIU.
The DEPLOY command from the EEC is routed through the EIU as a
second level of protection against inadvertent deployment. According
to commands from the EEC and the EIU, a Hydraulic Control Unit
(HCU) supplies hydraulic power to operate the thrust reverser. The
thrust reverser assembly has 2 hydraulically actuated translating
sleeves. The translating sleeves are each powered by 2 actuators. As
the translating sleeve moves aft during deployment, it raises blocker
doors to redirect the engine fan airflow.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 18
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
SYSTEM OVERVIEW - THRUST REVERSER SYSTEM
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 19
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE SYSTEM D/O (3)
ENGINE CHARACTERISTICS
The Airbus A319/A320/A321 are powered by two International Aero
Engines (IAE) V2500- A5 turbofan engines. These engines can produce
a thrust ranging from 22,000 lbs (9,980 kg) to 33,000 lbs (14,970 kg)
depending on the aircraft version set by the engine data-programming
plug.
PYLON
The engines are attached to the lower surface of the wings by pylons.
The pylons supply an interface between the engine and the aircraft for
electrics and fluids.
NACELLE
The engine is enclosed in the nacelle, which supplies aerodynamic
airflow around the engine and ensures protection for the accessories.
ENGINE CONTROL
The engine includes a Full Authority Digital Engine Control (FADEC),
which supplies engine control, engine monitoring and help for
maintenance and trouble shooting.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 20
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE CHARACTERISTICS - PYLON ... ENGINE CONTROL
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 21
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE GENERAL PARAMETERS
There is a different kind of thrust depending on the engine installed on
the aircraft. For the A319, 22,000 lbs (9,980 kg) can be achieved
during take off conditions with the V2522-A5, and 24,000 lbs (10,890
kg) with V2524-A5 version. For the A320, 27,000 lbs (12,250 kg) can
be achieved during take off conditions with the V2527-A5 version. For
the A321, 31,400 lbs (14,240 kg) can be achieved during take off
conditions with the V2530-A5 and 33,000 lbs (14,970 kg) with the
V2533-A5 version. Other general parameters of the power plant are
shown in the table too.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 22
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE GENERAL PARAMETERS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 23
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
N1 MODE. P/B
ENGINE CONTROL P/B'S AND SWITCHES
Position ON:
•
ENGINE MODE SELECTOR
switches EEC from EPR Mode to N1 Mode
Position CRANK:
• selects FADEC power.
• allows dry and wet motoring (ignition is not available)
Position IGNITION/ START:
• selects FADEC power
• allows engine starting (manual and auto).
Position NORM:
•
FADEC power selected OFF (Engine not running)
ENGINE MASTER LEVER
Position OFF:
•
closes the HP fuel valve in the FMU and the LP fuel valve and
resets the EEC.
Position ON:
•
•
starts the engine in automatic mode (when the mode selector
is in IGNITION / START).
selects fuel and ignition on during manual start procedure.
MANUAL START P/B
ƒ
controls the start valve (when the mode selector is in
IGNITION/ START or CRANK position).
FADEC GND PWR P/B B POSITION ON :
•
selects FADEC power
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 24
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE CONTROL P/B’S AND SWITCHES
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 25
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
SECONDARY ENGINE DISPLAY
ENGINE INDICATING PRESENTATION
The lower display shows the secondary engine parameters listed
below. The engine page is available for display by command, manually
or automatically during engine start or in case of system fault:
INDICATION GENERAL
PRIMARY ENGINE DISPLAY
ƒ
The primary engine parameters listed below are permanently displayed
on the Engine and Warning display E/ WD
ƒ
ƒ Engine Pressure Ratio EPR
ƒ Exhaust Gas Temperature (EGT
ƒ N1 (low rotor speed )
ƒ N2 (high rotor speed
ƒ FF (fuel flow)
After 5 min of the power up test the indication is displayed in amber
and figures are crossed (XX). Normal indication can be achieved by
using the FADEC GRD power switches, one for each engine at the
maintenance panel or by the MODE selector switch on the Engine
panel at the pedestal in CRANK or IGN/ START position for both
engine.
If a failure occurs on any indication displayed, the indication is
replaced by amber crosses, the analog indicator and the marks on the
circle disappear, the circle becomes amber.
Only in case of certain system faults and flight phases a warning
message appears on the Engine Warning Display.
TDTI / HAT / ATA 71-80
Issue: 06/08
ƒ
ƒ
ƒ
ƒ
ƒ
Total FUEL USED
For further info see ATA 73
OIL quantity
For further info see ATA 79
OIL pressure
For further info see ATA 79
OIL temperature
For further info see ATA 79
Starter valve positions, the starter duct pressure and during
eng start up, that operating Ignition system (ONLY ON ENGINE
START PAGE)
In case of high nacelle temperature a indication is provided
below the engine oil temp. indication.
Engine Vibration - of N1 and N2
As warnings by system problems only:
ƒ
ƒ
ƒ
Some
OIL FILTER CLOG
Fuel FILTER CLOG
No. 4 BRG SCAV VALVE with valve position
engine parameters also displayed on the CRUISE page
Revision: 18.06.2008
Page 26
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE PARAMETERS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 27
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
POWER PLANT DRAIN PRESENTATION (3)
GENERAL
The power plant drain system collects fluids which can leak from the
pylon, the engine accessories and drives. The fluids collected from the
power plant can be fuel, oil, hydraulic or water. They are discharged
overboard through the pylon drains and the engine drains.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 28
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
GENERAL
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 29
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
PYLON DRAINS
The engine pylon is divided into 7 compartments; various systems are
routed through these areas. Any leakage from fluid lines is drained
overboard through separate lines in the rear of the pylon.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 30
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
PYLON DRAINS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 31
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE DRAINS
Fluid drained from the oil tank scupper, fuel diverter valve and gear
box mounted accessories, is independently routed to the drain mast.
The fuel drains from the core engine accessories, are routed through a
separate drain line which passes through the bifurcation panel.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 32
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
ENGINE DRAINS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 33
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
POWER PLANT INSTALLATION D/O (3)
AIR INTAKE COWL
The air intake cowl is bolted onto the front of the fan case flange. It
includes an anti-ice system, an interphone jack and a P2/ T2 probe.
For removal and installation, the components that follow supply the
inlet cowl:
•
•
•
4 hoisting points,
36 identical attach fittings,
4 alignment dowels.
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 34
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
AIR INTAKE COWL
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
Page 35
A318/A319/A320/A321
DIFF IAE V2500 / Cat B1
Austrian Technical Training
School Notes - For Training Purposes Only
FAN COWL DOORS (LH & RH)
There are two fan cowl doors to give access to the fan case and
gearbox installed accessories. Four hinges hold each door at the pylon.
Four latches latch the door assembly along the bottom centerline. Each
door is installed with:
•
•
2 hoisting points, for removal and installation,
2 hold-open struts, to open. Access doors are also installed for
the start valve and to service the oil tank.
NOTE: THAT AERODYNAMIC STRAKES ARE INSTALLED ON THE
INBOARD NACELLE SIDE (A320).
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Revision: 18.06.2008
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FAN COWL DOORS (LH & RH)
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Revision: 18.06.2008
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THRUST REVERSER "C" DUCTS
The thrust reverser "C" ducts are in two halves installed with cascades,
blocker doors and translating sleeves. Four hinges hold each half at
the pylon. 6 latches latch the halves assembly along the bottom
centerline.
Each half is installed with:
• 3 attachment points for handling,
• 1 opening actuator operated with a hand pump,
• 2 hold-open rods for opening.
The latch assembly includes:
•
•
•
1 FWD bumper latch,
3 center latches, with access through a hinged access panel,
1 aft twin latch.
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Revision: 18.06.2008
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THRUST REVERSER "C" DUCTS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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FIREWALLS AND ACOUSTIC PANELS
The firewalls and fire seals protect from fire (to a fireproof standard)
between the power plant designated fire zones. The fire seals divide
the space in the engine into compartments. This means isolation limits
propagation, if a fire occurs. The aft bulkhead of the intake cowl and
the rear of the inner barrel are made of carbon composite sandwich
and the two supplies a firewall barrier to the fan case compartment
(zone 1). The inner barrel in the air intake cowl is made of carbon fiber
composite/Nomex honeycomb for acoustic purpose and is bolted to
the engine fan casing front flange. The acoustic panels are structural
and transmit air-intake cowl loads. The mixed exhaust system
structure has acoustic panels. The engine exhaust cone makes the
inner contour of the common nozzle exhaust collector. This exhaust
cone is made of a welded inco 625 honeycomb perforated panel for
sound attenuation. In the thrust reverser, the translating sleeves are
also installed with acoustic panels. The acoustic panels are sandwich
panels.
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Revision: 18.06.2008
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FIREWALLS AND ACOUSTIC PANELS
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
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COMMON NOZZLE ASSEMBLY
The Common Nozzle Assembly (CNA) mixes the exhaust gases from
the secondary and primary airflows. It is bolted to the rear flange of
the turbine exhaust case. 56 bolts attach the CNA to the LP turbine
frame.
EXHAUST CONE
The exhaust cone supplies the inner contour of the common exhaust
stream flow. It is attached to the inner flange of the turbine exhaust
case. 13 bolts attach the exhaust cone to the inner LP turbine frame.
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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COMMON NOZZLE ASSEMBLY & EXHAUST CONE
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Revision: 18.06.2008
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FWD MOUNT
The FWD mount transmits the engine thrust, vertical and side loads. It
is installed on the intermediate case, at the top center of the LP
compressor, with the FWD support shaft of the engine. The FWD
mount is made of 4 main parts:
•
•
•
•
two thrust link assemblies,
a beam assembly,
a cross beam assembly,
a support bearing assembly. The FWD mount is made to be
fail-safe. 4 bolts and self-locking nuts link it to the pylon FWD
part.
AFT MOUNT
The aft mount transmits the engine torque, vertical and side loads. It
is free in FWD and aft directions to let engine thermal expand. It is
installed on the case lugs of the turbine exhaust. The aft mount is
made of 3 main parts:
•
•
•
a beam assembly,
two side link assemblies,
a center link assembly. The aft mount is made to be fail-safe. 4
bolts link it to the pylon aft part.
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Revision: 18.06.2008
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FWD MOUNT & AFT MOUNT
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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FLUID DISCONNECT PANEL
The fluid disconnect panel supplies the fluid connection between
engine and pylon. It is installed on the LH side of the fan-case upper
part. The fluid connection lines for the fuel system are:
• fuel supply,
• fuel return to tank.
The fluid connection lines for the hydraulic system are:
•
•
•
hydraulic pump suction,
hydraulic pump pressure delivery,
hydraulic pump case drain.
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Revision: 18.06.2008
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FLUID DISCONNECT PANEL
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
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FAN ELECTRICAL CONNECTOR PANEL
The fan electrical connector panel supplies the interface between the
fan electrical harnesses and the pylon. It is installed on the RH side of
the fan-case upper part.
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Revision: 18.06.2008
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FAN ELECTRICAL CONNECTOR PANEL
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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CORE ELECTRICAL JUNCTION BOX
The core electrical junction box supplies the interface between the
core electrical harnesses and the pylon. It is installed in the FWD
mount zone.
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Revision: 18.06.2008
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CORE ELECTRICAL JUNCTION BOX
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Revision: 18.06.2008
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AERODYNAMIC STATIONS
Here are the main aerodynamic stations corresponding to the pressure
and temperature sensors installed on the engine:
•
STA 1: intake /engine inlet interfaces,
•
STA 2: fan inlet,
•
STA 12.5: fan exit,
•
STA 2.5: LP compressor exit,
•
STA 3: HP compressor exit,
•
STA 4: combustion section exit,
•
STA 4.5: HP turbine exit,
•
STA 4.9: LP turbine exit.
Here is the compressor stage numbering:
•
Stage 1.fan,
•
Stages 1.5 to 2.5: booster (LP compressor),
•
Stages 3 to 12: HP compressor.
Here is the turbine stage numbering:
•
Stages 1 and 2: HP turbine,
•
Stages 3 to 7: LP turbine.
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Revision: 18.06.2008
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AERODYNAMIC STATIONS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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ENGINE BEARINGS
The No.1, 2 and 3 (front) bearing compartment is built into the case
module and contains the support bearings for the low spool and high
spool stubshafts. The No.4 bearing compartment is part of the
diffuser/combustor module and the No 5 is located on the exhaust
case.
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Revision: 18.06.2008
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ENGINE BEARINGS - NO.1 AND NO.2 BEARINGS ... NO.4 AND NO.5 BEARINGS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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FRONT BEARING COMPARTMENT
The bearings No. 1, 2 and 3 are located in the front bearing
compartment, which is at the center of the intermediate module 32.
The compartment is sealed using air supported carbon seals, and oil
filled (hydraulic) seal between the two shafts. This seal is supported by
8th stage air.
the rods and any radial movement is dampened by oil pressure fed to
an annulus around the bearing outer race. The gearbox drive gear is
splined onto the HP shaft and retained by No.3 bearing nut.
Adequate pressure drops across the seals to ensure satisfactory
sealing. This is achieved by venting the compartment, by an external
tube, to the de-oiler.
GEARBOX DRIVE
The HP stubshaft, which is located, axially by No 3 bearing, has at its
front end a bevel drive gear, which provides the drive for the main
accessory gearbox, through the tower shaft. The HP stubshaft
separates from the HP compressor module at the curvic coupling and
remains as part of the intermediate case module.
DESCRIPTION
The drawing below shows details of No.2 and No.3 bearings.
A phonic wheel is fitted to the LP stubshaft, this interacts with speed
probes to provide LP shaft Speed signals (N1) to the EEC and the
Engine Vibration Monitoring Unit (EVMU), which is aircraft mounted.
The hydraulic seal prevents oil leakage from the compartment passing
rearwards between the HP and LP shafts.
No.3 bearing is hydraulically damped. The oil flow to the No.3 bearing
damper is maintained at the full oil feed pressure whilst the rest of the
flow passes through a restrictor to drop the pressure. This allows
larger jet diameters to facilitate flow tolerance control. The outer race
is supported by a series of eighteen spring rods which allow some,
slight radial movement of the bearing. The bearing is centralised by
TDTI / HAT / ATA 71-80
Issue: 06/08
Revision: 18.06.2008
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FWD BEARING COMPARTMENT
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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NO.4 BEARING COMPARTMENT
The No. 4 bearing compartment is situated in a high temperature, high
pressure environment at the centre of the combustion section.
The bearing compartment is shielded from radiated heat by a heat
shield and air.
The No. 4 bearing compartment is cooled by 12th stage air.
LOCATED IN THE EXHAUST AT 5 O' CLOCK POSITION (AFT
LOOKING FORWARD
12TH STAGE AIR COOLER (BUFFER AIR)
The No. 4 bearing compartment air cooler is installed on the turbine
casing. The exchanger is held by its coolant air duct flanges.
12TH STAGE AIR ( BUFFER AIR)
This supply of cooled 12th stage air (called "buffer) air” is admitted to
the space between the chamber and first heat shield. The 12th stage
air is cooled by fan air via the buffer air cooler, located on the rear left
hand side of the engine.
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.
NOTE: 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/
SCAVENGE LINE.
NOTE: A DRAIN HOLE IS PROVIDED TO INDICATE A POSSIBLE
LECKAGE AT THE NO.4 BEARING COMPARTMENT. IT IS
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Revision: 18.06.2008
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NO. 4 BEARING COMPARTMENT
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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REAR BEARING COMPARTMENT
The rear bearing compartment is located at the center of the LP
turbine module (module 50 ) and houses No.5 bearing which supports
the LP turbine rotor.
The compartment is sealed at the front end by an 8th stage air
supported carbon seal. At the rear is a simple cover plate, with an Oring and a thermally insulated heat shield, both secured by the same
twelve bolts. Inside the LP shaft there is a small disc type plug with an
O-ring seal, secured by a spring clip. There are no air or oil flows
down the LP shaft.
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.
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Revision: 18.06.2008
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REAR BEARING COMPARTMENT
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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STAGE 2 HPT AIR SEALS
ENGINE SEALS
The most important seals in this engine are:
•
•
•
•
•
front bearing compartment seals,
diffuser case and combustor assembly seal.
stage 1 High Pressure Turbine (HPT) air seals (inner and
outer).
stage 2 HPT air seal,
LPT seals.
FRONT BEARING COMPARTMENT SEALS
The sealing of the compartment is done by the following components:
•
•
•
•
The stage 2 HPT air seal is located between the stage 1 and 2 turbine
rotor assemblies. The stage 2 air seal also acts as a spacer for proper
axial positioning of both rotors.
LPT SEALS
The outer static stages 3, 4, 5 and 7 rotor seals are separated seal
segments with braze in honeycombs installed in the LPT case. The
bearing No. 5 compartment is sealed from the LPT cavity air by two
element radial carbon seals. The carbon elements are contained in a
housing assembly and sealed against the No. 5 bearing compartment
outer wall and the seal housing walls.
No.1 bearing front carbon oil seal,
hydraulic seal assembly for center seal,
bearing compartment rear brush air seal,
and No.3 bearing rear carbon oil seal.
DIFFUSER CASE AND COMBUSTOR SEALS
A double seal including a brush and knife edge seals is bolted onto the
front inner diameter of the case. The seals prevent leakage of HPC
discharge air into the diffuser case. A seal is supplied at the transition
from the rear of the combustor outer liner to the turbine nozzle guide
vanes. The front lip of the combustor inner liner makes a seal with the
inner diameter of the combustor hood.
STAGE 1 HPT AIR SEALS
In the stage 1 turbine rotor assembly, the blades fit into axial slots in
the hub and are held in place at the front by the stage 1 air seal
(outer). The stage 1 HPT air seal (inner) is located at the front of the
stage 1 turbine hub. The stage 1 HPT inner and outer air seals control
leakage of stage 1 HPT blade cooling air from the cooling duct.
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Revision: 18.06.2008
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ENGINE SEALS - FRONT BEARING COMPARTMENT SEALS ... LPT SEALS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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COMPRESSOR
The compressor section includes three modules:
•
•
•
LPC (fan) module assembly,
LPC/intermediate case module,
HPC module.
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COMPRESSOR
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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LPC FAN MODULE ASSEMBLY
The LPC (fan) module is a rotor assembly, which includes twenty-two
blades and a disk. Rotation of the rotor causes air to be ingested into
the front of the engine and compressed. A large proportion of the
compressed air is delivered through the fan discharge duct to the
exhaust nozzle to supply the majority of engine thrust. The remainder
of the compressed air goes into the booster section for further
compression by the booster.
FAN MODULE
Module 31 is the complete Fan assembly and comprises:
ƒ
ƒ
ƒ
22 wide-cord titanium shroudless hollow fan blades
22 annulus fillers
the titanium fan disc
ƒ
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 achieved by removing the front blade
retaining ring, and sliding the blade along the dovetail slot in the disc.
The fan inner annulus is formed by 22 annulus fillers.
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Revision: 18.06.2008
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FAN ASSEMBLY AND ANNULUS FILLER
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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FAN DISC
NOSE CONE
The class-fibre 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. lce 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 THE WEIGHTS MUST
BE MARKED BEFORE REMOVAL TO ENSURE THEY ARE
REFITTED IN THE SAME POSITION.
ANNULUS FILLERS
The blades do not have integral platforms to form the gas-path inner
annulus boundary. This function is fulfilled by annulus fillers, which are
located between neighbouring pairs of blades. The material of the
fillers is aluminium. Each annulus filler has a hooked trunnion at the
rear and a dowel pin and a pin at the front. The rear trunnion is
inserted in a hole in the rear blade retaining ring. The front pins are
inserted in holes in the front blade retaining ring. The fillers are
radially located by the front and rear blade retaining rings. Each filler is
secured to the front blade retaining ring by a bolt. In order to minimize
the leakage of air between the fillers and the aerofoils, there is a
rubber seal bonded to each side of each filler.
TDTI / HAT / ATA 71-80
Issue: 06/08
The fan disk is driven through a curvic coupling, which attaches it to
the LP stub shaft. The curvic coupling radially locates and drives the
fan disk. During manufacture of the fan disk, it is dynamically balanced
by removal of metal from a land on the disk.
INTERMEDIATE CASE
FAN CASE
The fan case provides a titanium shroud around the fan rotor and
forms the outer annulus of the cold stream duct.
LP COMPRESSOR OUTLET GUIDE VANES
Aerodynamic control air flow within the cold air steam duct is achieved
by 60 vanes manufactured in aluminium. The vanes consist of 20
segments, each containing 3 vanes. Both sides of the vanes are
attached to the outer and inner platforms. The outer platform is bolted
to the fan case and the inner platform is pinned to the outer shroud
ring of the LP compressor stage 2.5 stator assembly.
BOOSTER STAGE BLEED VALVE (BSBV)
The bleed valve mechanism is supported by the intermediate structure
and the outer ring of the stage 2.5 vanes. Two actuating rods, which
are each motivated by actuators, allow an axial motion to the valve
ring via 2 power arms.
Revision: 18.06.2008
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BOOSTER STAGE BLEED VALVE ACTUATION
LPC INTERMEDIATE CASE
TDTI / HAT / ATA 71-80
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HP COMPRESSOR
The HP compressor has 10 stages. It utilises variable inlet guide vanes
at the inlet to stage 3 and variable stator vanes at stages 3, 4 and 5.
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 7 and
10 stage air offtakes.
NOTE: ON THE V2500-A1 THE INLET GUIDE VANES AND
STAGES 3, 4, 5 & 6 ARE VARIABLE.
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HIGH PRESSURE COMPRESSOR
TDTI / HAT / ATA 71-80
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COMPRESSOR OPERATION
Rotation of the fan rotor causes air to be ingested into the front of the
engine and to be compressed. Compressed air is then divided into two
separate airflows:
•
•
a large portion is delivered to the exhaust nozzle,
the remainder is compressed in the booster before being
compressed again by the HPC.
The fan outlet inner vane assembly, compressor stage 1.5, 2, 2.3 and
2.5 assemblies and compressor stage 2.5 vanes compress the air
progressively. Then the air is delivered at a suitable pressure to the
HPC. Then the HPC accepts air from the LPC and booster system,
further compress it, and direct the air to the diffuser/combustion.
LPC bleed valve and actuating mechanism are incorporated into the
bleed air system. The modulated bleed makes sure that the booster
has an adequate surge margin under all operating conditions.
Due to the compression ratio that supports each stage, the airflow
tries to expand in counter direction. If the entire compressor is not
capable to compress the airflow.
Stall is a local effect where the airflow is not compressed. Stall effects
can bring the compressor to surge. To prevent the compressor surge
the stall effects are controlled through the methods of airflow control.
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COMPRESSOR OPERATION
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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COMBUSTION SECTION
The diameter of the diffuser section is larger at the rear than at the
front. This diametral difference decreases the speed of the air and
changes the energy of the speed into pressure.
The combustion section includes primarily the diffuser case,
combustor, fuel injector and igniters. The high compressor exit guide
vanes, the turbine nozzle assembly and the No.4 bearing compartment
are also part of the module.
The diffuser case has twenty mounting pads, where the fuel injectors
are installed and two mounting pads where the igniter plugs are
installed and six borescope bosses located around the case. The
combustor is an annular type combustor, which includes an outer liner
assembly and an inner liner assembly.
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Revision: 18.06.2008
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COMBUSTION SECTION
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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COMBUSTION SECTION OPERATION
The combustion section has four primary functions:
•
•
straighten the flow of air from the HPC,
change the flow of air characteristics to get the best speed and
pressure for combustion,
• mix fuel with the air and supply ignition to make the fuel burn,
• hold the No.4 bearing in position.
In the combustion section the gases increase their energy with a
controlled combustion process. These gases release their energy to the
turbine nozzles driving the engine rotors.
The air is guided and diffused around the combustor through
diffuser case inner and outer walls. Air, which surrounds
combustor, is used in the combustion process for dilution and in
temperature profile control of the combustion gases and for
cooling of the combustor walls.
TDTI / HAT / ATA 71-80
the
the
exit
the
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COMBUSTION SECTION OPERATION
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TURBINE SECTION
The turbine section includes the HPT and LPT modules. The HPT is a
two stages turbine and drives the HPC and the accessory gearbox.
The LPT is a five stages module attached at the rear flange of the HPT
module.
HP TURBINE
10TH STAGE MAKE UP AIR VALVE
The two position stage 10 ON/ OFF valve is bolted to the 10th stage
manifold at the top of the engine compressor case.
PURPOSE
The make up air discharges into the area around No 4 bearing
housing and supplements the normal airflows in this area and
increases the cooling flow passing to the H.P. turbine, stage 2. All of
the HPT airfoils are cooled by secondary air flow. The first stage HPT
blades are cooled by the HPC discharge air, which flows through the
fist stage HPT duct assembly. The second stage vane clusters are
permanent cooled by 10th stage compressor air mixed with thrust
balance seal vent air supplied externally. The 10th stage air is supplied
through 4 tubes (2 tubes on each engine side). Second stage HPT
cooling air is a mixture of HPC discharge air and 10th stage
compressor (make up air). This air moves through holes in the first
stage HPT air seal and the turbine front hub into the area between the
hubs. The air then goes into the second blade root and out the cooling
holes,
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TURBINE COOLING
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10TH STAGE "MAKE-UP" AIR SYSTEM
INTRODUCTION
The make up air discharges into the area around No4 bearing housing
and supplements the normal airflows in this area and increases the
cooling flow passing to the H.P. turbine, stage 2. 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 E.E.C. as a function of corrected N2 and altitude.
OPERATION
Signals from the E.E.C. will energize/ deenergize 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 (solenoid
deenergized) 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 No 4 bearing housing.
NOTE: THE E.E.C. WILL KEEP THE AIR VALVE OPEN AT ALL
ENGINE OPERATING PHASES EXCEPT CRUISE. THE VALVE
INCORPORATES 2 MICRO SWITCHES FOR TRANSMITTING
VALVE POSITION TO THE E.E.C CHANNELS A & B.
The "fail safe" position is valve open, solenoid deenergized.
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Revision: 18.06.2008
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1OTH STAGE AIR SYSTEM
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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ACTIVE CLEARANCE CONTROL SYSTEM (ACC)
The ACC system uses air to control the tip clearances of the HPT and
LPT blades. Active Clearance Control (ACC) is used to control seal
clearances and to supply structural cooling.
HPT SECTION
The HPT rotor and stator assembly includes stage 1 turbine rotor
assembly; a HPT case and vane assembly; a stage 2 HPT airseal; and
a stage 2 turbine rotor assembly.
LPT CASE, VANES AND STATIC SEALS
The LPT case, stator vanes and seals are the static parts of the LPT
assembly. This assembly includes five stages of vanes, static air seals,
and diffuser of inner and outer segments assembled in a casing.
LPT ROTOR
The LPT rotor is a bolted configuration including five disks and
associated blades and rotating air seals.
LPT SHAFT
The LPT Shaft joins the LPT assembly with the LPC and the Fan. The
shaft is bolted to the LPT at the LPT rotor disk stage 6 and extends
forward through the engine.
TEC AND Nº5 BEARING ASSEMBLY
The Turbine Exhaust Case (TEC) is used to straighten the turbine
exhaust gases, and supplies structural support for the No.5 bearing.
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TURBINE SECTION - HPT SECTION ... ACTIVE CLEARANCE CONTROL SYSTEM (ACC)
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TURBINE SECTION OPERATION
The HPT rotor and stator assembly supplies the rotational driving force
to the HPC and accessory gearbox by extracting energy from the hot
combustion gases, and releases a gas stream to the LPT in order to
drive the LPC and the fan through the LPT shaft. Exhaust gas from the
LPT goes through a nozzle to supply propulsive thrust. Blade tip
clearance and LPT case heat expansion are controlled by an external
ACC system. Fan discharge air is directed externally to the LPT case
via the ACC tubes. This controls the heat expansion of the LPT case
and optimizes the blade tip clearances.
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TURBINE SECTION OPERATION
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ANGLE AND MAIN GEARBOX
The cast aluminium 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 hand cranking for the HP rotor (N2) for maintenance
operations. The gearbox is mounted by 4 flexible links to the bottom
of the fan case, main gearbox 3 links angle gearbox 1 link
FEATURES:
FRONT FACE
ƒ Individually replaceable drive units
ƒ Magnetic chip detectors
ƒ Main gearbox 2 magnetic chip detectors
ƒ Angle gearbox 1 magnetic chip detector
ƒ De-oiler
ƒ Pneumatic starter
ƒ Dedicated generator/ alternator
ƒ Hydraulic pump
ƒ Oil Pressure pump
REAR FACE
ƒ Fuel pumps (and Fuel Metering Unit FMU)
ƒ Oil scavenge pumps unit
ƒ Integrated Drive Generator System (I.D.G.)
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ANGLE / MAIN GEARBOX
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BORESCOPE PORTS
Several ports are provided on the engine for borescope inspection.
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BORESCOPE PORTS
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COMPONENT ARRANGEMENT
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ENGINE R/H SIDE
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ENGINE L/H SIDE
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ENGINE FUEL SYSTEM D/O (3)
GENERAL
The engine fuel system is designed to provide fuel flow into the
combustion chamber, servo fuel for actuation of the compressor
airflow control and turbine clearance control systems and cooling for
engine oil and Integrated Drive Generator (IDG) oil.
FUEL FEED
The fuel coming from the aircraft tanks supplies the Low Pressure (LP)
fuel pump then provides engine oil cooling, through the fuel cooled oil
cooler. It then passes through the filter before entering into the High
Pressure (HP) pump then into the Fuel Metering Unit (FMU). A fuel
Differential Pressure (DELTA P) Switch provides indication to the
cockpit if the filter is clogged. A fuel temperature sensor is installed at
the fuel filter outlet for the fuel diverter and return valve operation.
NOTE: THE LP SOV, PRESSURE RAISING AND SOV CLOSE
WHEN THE ENGINE MASTER LEVER IS SET TO OFF. THE
PRESSURE RAISING AND SOV COULD BE AUTOMATICALLY
CLOSED BY THE EEC, DURING AUTO START ONLY, TO ABORT
IN CASE OF AN INCIDENT.
SERVO FUEL
In the FMU, a servo regulator provides the HP fuel to:
• the FMU TM servo valves,
• the Air Cooled Oil Cooler (ACOC) actuator,
• the booster stage bleed valve actuators,
• the Variable Stator Vane (VSV) actuator,
• the Active Clearance Control (ACC) actuator.
The servo regulator of the FMU regulates a fuel pressure to the
compressor airflow control systems, i.e. the booster stage bleed valve
and VSV and the pressure of the turbine ACC system.
DIVERTED FUEL
METERED FUEL
The fuel from the fuel pump assembly passes through a fuel metering
valve, an overspeed valve and a pressure raising and Shut-Off Valve
(SOV) included in the FMU. The fuel flow is then routed to a fuel
distribution valve, which supplies 20 nozzles through 10 manifolds. The
valves included in the FMU are controlled by the Electronic Engine
Control (EEC) through Torque Motors (TMs) to ensure fuel metering,
overspeed protection, pressure raising and shut-off functions. A flow
adjuster installed below the metering valve is used for the fuel flow
adjustment between "high flow" and "normal flow" by a maintenance
action.
TDTI / HAT / ATA 71-80
Issue: 06/08
Part of the fuel is used to provide adequate cooling of the engine oil
and IDG oil, and to maintain engine fuel and oil temperatures within
specified limits. These functions are controlled through a fuel diverter
and return valve, which incorporates a module to permit fuel to be
returned to the aircraft tanks under certain conditions. The EEC
processes the operation modes of the fuel diverter and return valve by
software logic. The logic is generated around the limiting temperatures
of fuel and oil, to provide the heat management system.
Revision: 18.06.2008
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GENERAL ... EEC CONTROL
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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EEC CONTROL
The EEC controls the operation of the FMU TM servo valves, the fuel
diverter and return valve, the ACOC actuator, the actuators of the
booster stage bleed valve, VSV and ACC systems. The EEC performs
control functions and fault analysis required to regulate the fuel and to
maintain the engine operation in all conditions. In the event of loss of
control functions on both channels, each servo valve and actuator has
a fail-safe position.
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Revision: 18.06.2008
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FUEL FILTER CLOGGING INDICATION
INDICATING
GENERAL
The fuel filter clog indication is provided on the lower ECAM display
unit. When the pressure loss in the fuel filter exceeds 5 plus or minus 2
psid, the pressure switch is energized. This causes:
GENERAL
INDICATING
The engine fuel system is monitored from:
• the ECAM display,
• the warning and caution lights.
The indications cover all the main engine parameters through the
FADEC.
The warning and cautions reflect:
•
•
•
the engine health and status through the FADEC,
the FADEC health & status,
the fuel fitter condition through a dedicated hardwired pressure
switch.
The fuel system is monitored by:
•
•
Triggering of the MASTER CAUT light and single chime.
The engine page to come on the lower ECAM DU with the
caution signal FUEL CLOG.
• The associated caution message to come on the upper ECAM
DU.
When the pressure loss in the filter decreases between 0 and -1.5 psid
from the filter clog energizing pressure, the pressure switch is deenergized which causes the caution to go off. The differential pressure
switch signal is fed directly to the SDAC through the hardware.
The fuel flow indication on the upper ECAM display unit permanently
displayed in green and under numerical form.
The fuel filter clogging caution (amber) on the lower ECAM display unit
associated with the MASTER CAUT light and the aural warning (singlechime).
Fuel flow indication, Fuel Used:
The Fuel Flow Transmitter is installed near the FMU. The signals are
routed to the EEC and via the DMCs to the ECAM.
The Fuel Used-is calculated in the DMCs .
The fuel flow transmitter signal is fed to the FADEC, which processes it
and transmits the information to the ECAM system for display.
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FUEL INDICATION
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Revision: 18.06.2008
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FUEL DIVERTER & RETURN VALVE
FUEL DISTRIBUTION COMPONENTS
GENERAL
FUEL FILTER
DESCRIPTION
The fuel filter element is a low pressure filter which removes all
contamination from fuel to go through it.
The filter element is installed in the lower housing of a fuel cooled oil
cooler (FCOC). The FCOC includes the following components:
A fitter cap, which has a pressure plate to keep the filter element in
position once installed. The filter cap of the FCOC also includes a fuel
Drain plug to drain the fuel for maintenance purposes.
The fuel diverter and return valve (FD & RV) is a primary unit in the
heat management system (HMS) of the engine. The FD & RV has two
valves in one body. They are a fuel diverter valve (FDV) and a fuel
return valve (FRV).The FDV operates to change the direction of the fuel
metering unit (FMU) spill flow to:
• The fuel cooled oil cooler (FCOC) or,
• the fuel filter ( element ) inlet or,
• the fuel cooled IDG oil cooler (IDG FCOC).
The FRV operates to control fuel flow, which goes back to the aircraft
fuel tank acting as a fuel cooler.
A filter bypass valve to let the fuel go around the filter element when
it be comes clogged.
FUEL FILTER DIFF. PRESS, SWITCH
The fuel filter clog indication is provided on the lower ECAM display
unit. When the pressure loss in the fuel filter exceeds 5 plus or minus
2 psid, the pressure switch is energized. When the pressure loss in the
filter decreases between 0 and -1.5 psid from the filter clog energizing
pressure, the pressure switch is deenergized which causes the caution
to go off. The differential pressure switch signal is fed directly to the
SDAC
FUEL TEMPERATURE THERMOCOUPLE
The measured temperature is transmitted to the EEC (Electronic
Engine Control) and used for the Heat Management System.
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FUEL DISTRIBUTION COMPONENTS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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FUEL DISTRIBUTION VALVE
GENERAL
The fuel distribution valve (FDV) subdivides scheduled engine fuel
flow from the fuel metering unit (FMU) equally to ten fuel manifolds,
each of which in turn feeds two nozzles.
DESCRIPTION
The fuel distribution valve is installed at the 4:00 o'clock location, at
the front flange of the diffuser case. The fuel distribution valve
receives fuel through a fuel line from the fuel metering unit. The fuel
goes through a 200 micron strainer, and then into ten internal
discharge ports. The ten discharge ports are connected to the ten fuel
manifolds. Eight of the ten internal discharge ports in the valve are
connected after an engine shutdown. Eight of the fuel manifolds are
drained into the engine through the lowest fuel nozzle.
The two fuel manifolds, which remain full, held supply fuel for the
next engine start.
LOCATION FUEL DISTRIBUTION VALVE
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FUEL DISTRIBUTION VALVE
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on areas where fuel touches metal. Heatshields installed also prevent
formation of coke.
FUEL MANIFOLD AND TUBES
DESCRIPTION
The fuel manifold and fuel tubes consist of several single wall tubes,
which carry fuel between components in the fuel system. Fuel
supplied to the fuel nozzles is carried by a large tube from the fuel
metering unit to the fuel distribution valve. At the fuel distribution
valve the fuel supply is split and carried to twenty fuel nozzles by ten
manifolds. Each fuel manifold feeds two fuel nozzles. Fuel pressure for
actuating various valves is supplied by small tubes from the fuel
metering unit mounted on the fuel pump. All the brackets and tubings
are fire proof.
FUEL NOZZLE
GENERAL
The fuel nozzles receive fuel from the fuel manifolds. The fuel nozzles
mix the fuel with air, and send the mixture into the combustion
chamber in a controlled pattern.
DESCRIPTION/OPERATION
There are 20 fuel nozzles equally spaced around the diffuser case
assembly. The fuel nozzles are installed through the wall of the case,
and each nozzle is held, in position by three bolts. The fuel nozzles
carry the fuel through a single orifice. The fuel is vaporized by highvelocity air as it enters the combustion chamber. The fuel nozzle
forms the atomized mixture of fuel and air into the correct pattern for
satisfactory combustion. The design of the fuel nozzle results in fast
vaporization of the fuel through the full range of operation. This
results in decreased emissions, high combustion efficiency, and good
start quality. The high-velocity flow of fuel prevents formation of coke
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FUEL NOZZLES
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IDG FUEL COOLED OIL COOLER
The IDG oil cooler is installed at the left hand side on the fan case,
near the FCOC. The IDG oil cooler has two sets of inlet and outlet
ports. One set of ports is used for the flow of the fuel to or from the
fuel diverter and return valve. The other set of ports is used for the
flow of oil from and to the IDG. The hot scavenge oil, which has been
used to lubricate and cool the IDG, flows from the IDG to the oil
cooler. As the oil goes through the oil cooler, the heat in the oil is
transmitted to the fuel. The cooled oil then returns to the IDG. Two
drain plugs are also installed in the oil cooler, one for the fuel and one
for the oil.
IDG OIL COOLER TEMP. THERMOCOUPLE
This temperature information is send to the EEC and is used for the
heat management system.
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IDG FUEL COOLED OIL COOLER AND IDG OIL TEMP. THERMOCOUPLE
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FUEL PUMP
General The LP 1 HP fuel pumps are housed in a single pump unit,
which is driven by a common gearbox output shaft. A low pressure
(LP) stage and a high pressure (HP) stage provide fuel at the flows
and pressures required for operation of hydromechanical components
and for combustion in the burner. The unit consists of a LP centrifugal
boost stage which feeds an HP single stage, two gear pump. The
housing has provision for mounting the fuel metering unit (FMU)
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 isolation of the fuel supplies at engine stop
NOTE: THERE ARE NO MECHANICAL INPUTS TO, OR OUTPUTS
FROM THE FMU.
The LP stage receives fuel from aircraft tanks through the aircraft
pumps. The LP pump is designed to provide fuel to the HP gear stage
with the aircraft pumps inoperative. After passing through the LP
boost stage, fuel proceeds through the fuel filter to the HP gear stage.
A coarse mesh strainer is provided at the inlet to the HP gear stage.
This stage is protected from overpressure by a relief valve. Exceeding
flow from the gearstage pump is recirculated through the FMU bypass
loop to the low pressure side of the PUMP.
FUEL METERING UNIT
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 F.M.U. The FMU, under the control of the EEC
meters the fuel supply to the 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
Pressure Raising and Shut Off Valve
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FMU FUEL METERING UNIT
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PRESSURE RAISING AND SHUT OFF VALVE
OVERSPEED VALVE
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 (109,1 % N1, 105,4% N2) 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 shut down the
engine.
To shut down the engine is the only way to release the hydraulic
latching.
The PRSOV torque motor is commanded open by the EEC during AUTO
starts.
It is commanded open by the MASTER SWITCH in the cockpit 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.
NOTE: THE EEC'S ABILITY TO CLOSE THE SHUT OFF VALVE IS
INHIBITED ABOVE 43% N2. ABOVE 43% N2, AND IN FLIGHT,
THE PRSOV CAN ONLY BE CLOSED BY THE MASTER SWITCH
IN THE COCKPIT.
FAIL SAFE POSITION OF THE PRSOV: " LAST COMMANDED
POSITION”
The engine fuel supply system has two fuel shut off valves.
ƒ
one PRSOV in the FMU
ONE LP - FUEL SHUT OFF VALVE, ON THE FRONT WING SPAR.
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.
FAIL SAFE POSITION: " NORMAL FUEL METERING"
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FMU OPERATION
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ENGINE MASTER LEVER OPENING COMMAND
During the start sequence, the EEC controls the opening of the Fuel
Metering Valve (FMV), which causes the pressure raising and SOV to
open, provided its latching torque motor is de-energized. When the
MASTER lever is set to ON, the pressure raising and SOV will only open
if fuel pressure from the FMV is available.
NOTE: DURING ENGINE START IN AUTOMATIC MODE, THE
EEC CAN CLOSE THE PRESSURE RAISING AND SOV TO ABORT
THE START SEQUENCE UP TO 50% HIGH PRESSURE ROTOR
SPEED (N2).
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GENERAL & ENGINE MASTER LEVER OPENING COMMAND
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ENGINE MASTER LEVER CLOSURE COMMAND
The closure of the pressure raising and SOV is controlled directly from
the MASTER lever. When it is set at the OFF position, it energizes the
latching torque motor which closes the valve. A time-delay relay
automatically de-energizes the pressure raising and SOV latching
torque motor 2 minutes after engine shutdown. This device avoids
damage due to high temperature induced by a long time power supply
on ground.
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ENGINE MASTER LEVER CLOSURE COMMAND
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MONITORING
The pressure raising and SOV is monitored by two microswitches,
which send signals to the EEC and then to the Engine Interface Unit
(EIU). In case of disagreement between control and position, an ECAM
warning is triggered and the FAULT light comes on.
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MONITORING
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ENGINE LP FUEL SHUT-OFF VALVE CONTROL (3)
GENERAL
The LP fuel - valve 12QM (13QM) is in the fuel supply line to its related
engine. The LP fuel - valve is usually open and in this configuration
let’s fuel through to its related engine. When one of the LP fuel valves is closed, the fuel is isolated from that LP fuel valve's related
engine. The LP fuel - valve is installed between the engine pylon and
the front face of the wing front spar (between RIB8 and RIB 9)
• a valve spindle
• a mounting flange.
The LP fuel - valve actuator has two electrical motors, which drive the
same differential - gear to turn the ball valve through 90 deg. The limit
switches in the actuator control this 90 deg. movement and set the
electrical circuit for the next operation. One of the two motors can
open or close the valve if the other motor does not operate. The
actuator drive shaft has a see/ feel indicator where it goes through the
actuator body. The see/ feel indicator gives an indication of the valve
position
without
removal
of
the
fuel
LP
fuel
valve.
Each LP valve has an actuator 9QG (10 QG). The interface between
the actuator and the LP valve is a valve spindle. When the actuator is
energized, it moves the LP valve to the open or closed position. A V band clamp 80QM(81QM) attaches the actuator to the LP valve.
Each actuator has two motors, which get their power supply from
different sources :
• the 28VDC BATT BUS supplies, the motor
• the 28VDC BUS 2 supplies the motor 2.
If damage occurs to the electrical circuit, it is necessary to make sure
that the valve can still operate. Thus the electrical supply to each
motor goes through a different routing. The routing for motor 1 is
along the front spar. The routing for motor 2 is along the rear spar
and then forward through the flap track fairing at RIB 6.
The actuators send position data to the System Data Acquisition
Concentrators (SDAC1 and SDAC2). The SDACs process the data and
send it to the ECAM, which shows the information on the FUEL page.
COMPONENT DESCRIPTION
The LP fuel - valve has:
•
•
a valve body
a ball valve
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GENERAL
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ENGINE MASTER LEVER CONTROL
When the MASTER lever is set to OFF, both electrical motors drive the
LP SOV to the closed position.
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ENGINE MASTER LEVER CONTROL
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ENGINE FIRE P/B COMMAND
When the ENG FIRE P/B is released out, both electrical motors drive
the LP SOV to the closed position.
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ENGINE FIRE P/B COMMAND
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HEAT MANAGEMENT SYSTEM D/O (3)
GENERAL
Heating and cooling of the fuel, engine oil and Integrated Drive
Generator (IDG) oil is performed by the fuel cooled oil cooler, the Air
Cooled Oil Cooler (ACOC) and the IDG fuel cooled oil cooler under the
control of the Electronic Engine Control (EEC). The EEC acts on the
heat management system through the fuel diverter and return to tank
valves and the ACOC modulating valve.
AIR MODULATING VALVE
PURPOSE
To govern the flow of cooling (fan) air through the air/ oil heat
exchanger (ACOC), 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
COMPONENTS
Bolted to the outlet face of the air/oil heat exchanger.
ACOC MODULATING VALVE
Oil heated by the engine passes through the ACOC and then to the
fuel cooled oil cooler. The ACOC modulating valve regulates a bleed
part of fan airflow crossing the ACOC to maintain both oil and fuel
temperatures within acceptable minimum and maximum limits.
FUEL DIVERTER AND RETURN TO TANK VALVE
The fuel diverter valve and the return to tank valve are enclosed in the
same housing. The fuel diverter valve and the return to tank valve
maintain fuel, engine oil and IDG oil temperatures within limits by
minimizing ACOC cooling air usage. The two-position diverter valve
works by managing the fuel recirculation inside the engine fuel system.
The return to tank valve will divert a modulated proportion of the LP
fuel back to the aircraft tanks.
FEATURES
•
fire seal forms an air tight seal between the unit outlet and the
cowling orifices
• controlled by either channel A or B of EEC
• valve positioned by fuel servo pressure acting on a control
piston
• valve position feed back signal via LVDT to each channel of EEC
• fuel servo pressure directed by the Electro - Hydraulic Servo
Valve assembly which incorporates a Torque motor
FAIL SAVE POSITION: AIR VALVE SPRING LOADED FULLY OPEN
(maximum cooling position)
In case of malfunction the warning
“ENG 1 (2) AIR EXCHANGER FAULT” is displayed on the ECAM E/ WD.
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GENERAL ... FUEL DIVERTER AND RETURN TO TANK VALVE
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CONTROL
The EEC controls the heat management system through 4 modes of
operation taking into account 3 parameters of temperature:
• engine fuel,
• engine oil,
• IDG oil.
The temperature of the aircraft fuel tanks and the engine power
setting are parameters used for inhibition of fuel return to aircraft
tanks.
NOTE: THE 4 MODES OF OPERATION ARE MODES 1, 3, 4 AND
5:
•
MODES 1 AND 4 ARE RETURN TO TANK MODES,
•
MODES 3 AND 5 ARE NO RETURN TO TANK MODES.
MODE 1
Normal return to tank mode: When the engine is not at high power
setting, some of the fuel is returned to the tank. The heat is absorbed
and dissipated within the tank. The ACOC modulating valve is fully
closed.
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CONTROL - MODE 1
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MODE 4
Mode selected when in normal mode, i.e. mode 1, there is a high
engine fuel temperature. In this mode the oil system is used to
achieve a supplemental cooling of the fuel. The ACOC modulating
valve is fully open.
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CONTROL - MODE 4
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MODE 3
Engine at high power setting: In this condition all the heat is absorbed
by the burned fuel. If however the fuel flow is too low to provide
adequate cooling the ACOC valve could be modulated.
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CONTROL - MODE 3
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MODE 5
Mode selected when system condition demand as in mode 3 but this is
not permitted because IDG oil temperature is excessive or return to
tank is not permissible due to the high return fuel temperature. The
ACOC valve is fully open.
NOTE: THIS IS THE FAIL-SAFE MODE OF OPERATION.
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CONTROL - MODE 5
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AIR SYSTEM PRESENTATION (3)
GENERAL
The engine air system makes sure that the compressor airflow and
turbine clearances are controlled. The system also deals with the
cooling and pressurizing airflows.
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GENERAL
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HP COMPRESSOR BLEED VALVES
COMPRESSOR AIRFLOW CONTROL
The 7th and 10th stages bleed valves maintain a more stable operation
of the compressor.
GENERAL
The booster stage bleed valve, the variable stator vane and HP
compressor bleed valves systems are controlled by the EEC. The
booster stage bleed valve controls the LP compressor airflow. The
variable stator vane and the 7th, and 10th stage bleed valves control
the HP compressor airflow.
COMPRESSOR AND CLEARANCE CONTROL LRU’S:
ƒ
ƒ
ƒ
ƒ
ƒ
ƒ
BOOSTER STAGE BLEED VALVE (BSBV) CONTROL
The EEC controls the BSBV position. The EEC uses the BSBV feedback
signal from the LVDT to adjust the actual BSBV position. At low LP
spool speeds the booster provides more air than the core engine can
utilize. To match the booster discharge airflow to the core engine
requirements at low speed, excess air is bled off through booster stage
bleed valves (BSBV) into the fan discharge air stream. At higher engine
speeds the BSBV are closed so that all the booster discharge (primary
air flow) enters the core engine.
ƒ
ƒ
ƒ
ƒ
ƒ
ƒ
ƒ
BSBV Master Actuator
Servo Valve
Feedback for EEC
BSBV Slave Actuator
Servo Valve
Feedback for EEC
VSV Actuator
Servo Valve
Feedback for EEC
7th Stage Bleed Valves (3)
7th Stage Solenoids (3)
10th Stage Bleed Valve
10th Stage Solenoid
VARIABLE STATOR VANE (VSV) CONTROL
The VSV position is controlled by the EEC. The EEC uses the VSV
feedback signal from the LVDT's to adjust the actual VSV position. The
VSV system maintains a satisfactory compressor performance over a
wide range of operating conditions. The system varies the angle of the
inlet guide vanes and stator vanes to aerodynamically match the low
pressure stages of compression with the high pressure stages. This
variation of vane position changes the effective angle at which the air
flows across the compressor blades and vanes. The VSV angle
determines the compression characteristics (direction and velocity) for
any particular stage at compression.
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COMPRESSOR AIRFLOW CONTROL
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BOOSTER BLEED VALVE AND ACTUATING MECHANISM
BOOSTER BLEED SYSTEM
GENERAL
The primary function of the LP compressor airflow control system is to
control the airflow thus ensuring compressor stable operation during:
ƒ
ƒ
ƒ
ƒ
ƒ
ƒ
Engine start.
Engine transient operation.
DESCRIPTION
GENERAL
The airflow control system includes:
•
Two bleed-valve actuating rods
•
Piston Jack Fork End
•
An LPC bleed-master actuator
•
An LPC bleed-slave actuator
•
Intermediate Structure
•
A booster bleed valve and actuating mechanism
DESCRIPTION
The bleed valve and actuating mechanism is a sub - assembly that
includes:
The support ring.
The ring valve
The two upper arms, the lower arms and the eight mid arms.
The two actuating rods connect the two upper power arms to
the two actuators. The bleed valve and actuating mechanism
operates to make each bleed valve synchronized, in relation to
the positions of the two actuators.
The airflow control system automatically operates to control the air
bled from the LP compressor. The two actuators are mechanically
attached to each actuating rod and, the bleed - valve and actuating
mechanism. The two actuators are connected hydraulically and
operate together by command and feedback signals from/ to the EEC.
FAIL.SAFE POSITION: “BSBV OPEN"
In case of a malfunction “ENG 1 (2) COMPRESSOR VANE” is displayed
on the ECAM E/WD.
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BOOSTER BLEED SYSTEM
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VSV SYSTEM
The four stages of variable incidence stators comprise inlet guide
vanes to stage 3 and stages 3, 4 and 5 stator vanes.
GENERAL
The purpose of this system is to position the Inlet Guide Vanes (IGV)
and stator vanes, using a fuel driven hydraulic actuator, in response to
electrical signals provided by the EEC.
VARIABLE STATOR VANE (VSV) CONTROL
The VSV position is controlled by the EEC as a function of N2/ square
root of theta T2.6 (synthesised value). The EEC uses the VSV feedback
signal from the LVDT's to adjust the actual VSV position.
DESCRIPTION
VARIABLE STATOR VANE ACTUATOR
The stator vane actuator accurately controls vane movement with
respect to a torque motor current supplied by the EEC. Operation of
the stator vanes in regulated by accurate control of high pressure fuel
flow to one or other side of a differential area piston. The piston has
an externally adjustable low speed stop at the extended end of its
travel. The high speed stop is formed by a collar which limits piston
retraction. Provision is made to lock the piston with a rigging pin for
setting purposes.
LINEAR VARIABLE DIFFERENTIAL TRANSFORMER (LVDT)
A Dual Wound Linear Variable Differential Transformer (LVDT) is
located in the center of the actuator piston rod. The LVDT completes
TDTI / HAT / ATA 71-80
Issue: 06/08
the electronic control loop by providing a signal of actuator position to
the Engine Electronic Control.
ENGINE LINKAGE WITH THE VSV ACTUATOR
The engine IGV and Stator Vane linkage is connected to a fork end on
the piston rod of the VSVA unit. The securing pin of link on to fork
end.
OPERATION OF THE MSV ACTUATOR
Dual wound torque motors convert electrically isolated drive signals
from each channel of the Electronics Engine Control (EEC) into
hydraulic drive signals to position the actuator piston. If power to the
stator vane actuator torque motor is lost, the stator vane actuator will
go to the full open position.
VARIABLE STATOR VANE ACTUATION MECHANISM
The variable geometry operating mechanism for the compressor
comprises the following elements
ƒ
ƒ
ƒ
ƒ
actuator/crankshaft drag link
crankshaft (steel)
four crankshaft/unison ring drag links
four unison rings spindle levers (titanium) variable IGV’s and
stage 3, 4, and 5 variable stators
FAIL SAFE POSITION: “VANES OPEN“
In case of a malfunction "ENG 1 (2) COMPRESSOR VANE” is displayed
on the ECAM E/ WD.
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VSV SYSTEM
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HANDLING BLEED VALVES
Handling bleed valves are fitted to the HP 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 EEC.
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 HP compressor
delivery air known as P3 or Pb.
Silencers are 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.
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HANDLING BLEED VALVE SYSTEM
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HANDLING BLEED VALVES FUNKTION
Description 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 the engine start. When the engine is started the bleed air
from the engine 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 deenergized). The bleed valves will be closed at the
correct time during an engine acceleration by the EEC 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,
OPENING-SCHEDULE
The schedule for one bleed valve - 7C - is shown, in detail, below.
During decelerations the valve will be commanded open whenever N2
is below the transient opening speed. The valve remains open until the
deceleration ceases and a deceleration time, 2 seconds, expires.
NOTE: THE TRANSIENT REGIME IS SLIGHTLY MODIFIED FOR
OPERATION ABOVE 15000 FT BUT OPERATES IN THE SAME
WAY.
SURGE/ REVERSE
If the engine is operating in reverse thrust operation is the same as
Transient but different speeds apply. 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 restabilizes.
During an engine deceleration the reverse operation occurs and the
bleed valve opens.
STEADY STATE
It can be seen that the valve will be commanded closed at stabilised
min idle, 8600 N2, and will not be opened again in Steady state.
TRANSIENT
The valve will be commanded open during engine acceleration
whenever N2 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).
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HANDLING BLEED VALVE OPERATION
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HANDLING BLEED VALVES (SURGE BLEED)
BLEED VALVE OPERATING SCHEDULE
The bleed valves and the solenoid control valves all operate in the
same manner.
Bleed Valve
Regime
Open (RPM)
Close (RPM)
7A
Steady State
11400
11800
FAIL SAFE POSITION: "7th and 10th OPEN".
(35000Ft&below)
11800
12250
(42000Ft&above)
Surge
Reverse
& 12562
12772
7B
Steady State
7650
8000
7C
Steady State
6800
7000
Transient
11600
12050
& 12352
12562
Surge
Reverse
10
Steady State
Surge
Reverse
IDLE 57% = 8800 RPM
7650
& 10667
8000
10667
(open below)
Max. TO 100% = 14950
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ENGINE OPERABILITY (SURGE FREE OPERATION)
HANDLING BLEED VALVE MALFUNCTIONS
Stuck closed
A engineering order (010169) is released to cover this problems.
7TH/ 10TH STAGE HANDLING BLEED VALVES STICKING
Hung starts or starting stalls experienced due to 7th and 10th stage
handling bleed valves failing to open or close. The consequences of the
malfunction of one or more handling bleed valves on:
ƒ the ground and airstart capability,
ƒ the engine operability (surge free operation)
ƒ the engine performance ( EGT, fuel consumption)
have been assessed and are summarized in the following tables:
NOTE: A BLEED TEST SET IS PROVIDED TO CHECK THE BLEED
VALVES AND SOLENOID VALVES FOR PROPER FUNCTION.
one or more 7th stage possible surge upon no effect
valve(s)
fast transient
10th stage valve
no effect
no effect
CLIMB AND CRUISE PARAMETERS
The effect of one 7th stage handling bleed valve being stuck opened,
has been assessed to be as follows:
typical climb
typical cruise
delta EGT
+30 to 40
+20 to 30
delta FF (%)
+5 to 6
+3 to 5
GROUND AND AIRSTART CAPABILITY
Stuck closed
Stuck open
7th stage valve(s)
hung start or sub- no effect
idle stall/ rol-back
10th stage valve
hung
stall)
start
TDTI / HAT / ATA 71-80
(rotating rising EGT
Issue: 06/08
Stuck open
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TURBINE CLEARANCE
COOLING
CONTROL
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AND
TURBINE
The EEC controls the actuation of an Active Clearance Control (ACC)
valve for the High Pressure (HP) and LP turbine active clearance
control. The ACC system fulfills the blade tip clearances of the turbines
for better performance. The EEC controls a dual ACC valve which
discharges fan air through manifolds to cool the surfaces of the HP and
LP turbine cases during climb and cruise power operations.
The continuous flow adapter supplies supplemental air, from the HPC
10th stage, to cool the stage 2 blades and HP 1 turbine disc bore.
The HP/ LP Turbine Active Clearance Control (ACC) system uses fan air
to cool the HP and LP cases for blade tip clearance control in order to
improve engine performance and maximize the turbine cases life time.
Fan air is drawn from a common HP/ LP turbine ACC air scoop in the
fan duct. This air is divided into HP and LP cooling air and passes
through individual short ducts to the Active Clearance Control Valves,
which direct air for both HP and LP turbine case cooling.
NOTE: THE HP TURBINE CLEARANCE CONTROL VALVE IS
EQUIPPED WITH 4 PLUGS IN THE VALVE VANE. THIS PLUGS
CAN BE REMOVED ACCORDING TO A SERVICE BULLETIN TO
ALLOW A PERMANENT COOLING OF THE HP TURBINE. IN
CASE OF A VALVE REMOVAL/ INSTALLATION THE SAME
CONFIGURATION MUST BE PROVIDED ON THE NEW VALVE. IF
THE PLUGS MUST BE REMOVED, THERE IS A STORAGE
BRACKET PROVIDED ON THE ACTUATOR ROD. DO NOT
THROW THE PLUGS AWAY!
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TURBINE CLEARANCE CONTROL AND TURBINE COOLING
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HPT/ LPT ACTIVE CLEARANCE CONTROL (HPT/ LPT ACC).
Take - off
The active clearance control (ACC) system ensures the blade tip
clearances of the turbines for better performance. The HPT/ LPT ACC
valve modulates fan air flow to the HP and LP turbine cases. The EEC
controls the valve position as a function of thrust level and altitude.
The LVDT's transmit the valve position to the EEC.
During take - off, the position of the actuator piston is at point C. At
this point:
HP TURBINE (10TH STAGE) COOLING AIR CONTROL
NOTE: THE ACTUATOR POSITION BETWEEN POINT C AND
POINT E DEPENDS ON ALTITUDE.
The HP turbine cooling air valve (make up air valve) supplies
supplemental air (from HP compressor 10th stage) to cool the 2nd
stage vanes, hubs and discs of the HP. The, valve operates as a
function of high rotor speed and altitude and incorporates a 2 position switch to provide a feedback signal to the EEC (channels A
and B). During cruise the valve is closed.
OPERATING SCHEDULE
ƒ
ƒ
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.
FAIL SAFE:
When there is no torque motor current or no fuel servo pressure, the
actuator piston moves to point A. LP valve will be partially open (-44
deg). The actuator piston remains at this point at all defective
conditions. (HP valve closed)
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:
ƒ
ƒ
The control valve for the HP turbine ACC is closed.
The control valve for the LP turbine ACC is not less than 44
percent opened.
Engine Operation
During engine operation, the EEC controls the position of the actuator
piston between point B and point E.
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HPT/ LPT ACTIVE CLEARANCE OPERATION
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OPERATION SCHEDULE
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N°4 BEARING COMPARTMENT COOLING
The number 4 bearing compartment is cooled by 12th stage air of the
HPC. An external line carries this air through an air-cooled air cooler
before entering into the Nº4 bearing compartment. The air-cooled air
cooler cools the 12th stage air of the HPC with fan air before going into
the Nº4 bearing compartment. The fan air is then discharged
overboard.
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N°4 BEARING COMPARTMENT COOLING
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NACELLE VENTILATION
Ventilation is provided for the fan compartment Zone 1, and the core
compartment Zone 2 to:
ƒ
ƒ
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 L.H.
side of the air intake cowl. The air circulates through the fan
compartment and exits at the exhaust located an 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 (A.C.C.) 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 fan case - Zone 2. Zone 2 ventilation
is still effected in the same way as when the engine is running.
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FAN AND CORE VENTILATION ZONES
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NACELLE TEMPERATURE
GENERAL
The Nacelle Temperature Sensor has a Measurement Range of –54°C
to 330°C
This Signal is fed to the EIU, which transforms the Information to
digital Form.
The EW Transmits the Data to the ECAM System.
The nacelle temperature is displayed if the system is not in engine
starting mode and one of the two temperatures reaches the advisory
threshold. A advisory indication will be created on the engine system
page when the temperature reaches approx. 300 – 320°C.
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NACELLE TEMPERATURE
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FADEC BENEFITS
FADEC PRESENTATION (3)
The application of a FADEC system provides multiple benefits:
•
PURPOSE
The Full Authority Digital Engine Control (FADEC) system provides full
range engine control throughout all flight and operational phases. It
consists of a dual channel Electronic Engine Control (EEC) and its
peripheral components and sensors.
FADEC FUNCTIONS
The FADEC provides the engine system regulation and scheduling to
control the thrust and optimize the engine operation.
The FADEC provides:
•
•
•
•
•
•
•
•
•
•
•
•
•
•
•
it saves weight and fuel by a full range control of the gas
generator,
it reduces pilot workload and maintenance cost,
it allows the optimum adaptation of thrust rating schedules to
the A/C needs.
POWER SUPPLY
The FADEC system is self-powered by a dedicated Permanent Magnet
Alternator (PMA) when N2 is greater than 10%. The EEC is powered
by the aircraft 28 VDC electrical network for starting, as a backup and
for testing with the engine not running. 115 VAC is used for the power
supply of the ignition system and the P2/T2 probe heating.
power setting with EPR or N1 back-up mode,
P2/T2 heating,
acceleration and deceleration times,
idle speed governing,
overspeed limits for N1 and N2,
Fuel Flow (FF) control,
Variable Stator Vane system (VSV) control,
compressor handling bleed valves control,
booster stage bleed valve system control,
High Pressure (HP)/Low Pressure (LP) turbine Active Clearance
Control (ACC),
automatic and manual engine starting,
thrust reverser control,
oil and fuel temperature management through the heat
management system.
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PURPOSE ... POWER SUPPLY
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motors and solenoids. The other channel calculation is used for
crosschecking.
FADEC ARCHITECTURE (3)
DUAL CHANNEL
BITE CAPABILITY
The Full Authority Digital Engine Control (FADEC) system is fully
redundant and built around two independent control channels. Dual
inputs, dual outputs and automatic switchover from one channel to the
other eliminate any dormant failure. The Electronic Engine Control
(EEC) consists of two channels A and B. Each channel can control the
different components of the engine systems. Channels A and B are
permanently operational. The channel in control manages the system.
FAULT STRATEGY
DUAL INPUTS
All control inputs to the FADEC system are doubled. Only some
secondary parameters used for monitoring and indicating are single.
To increase the fault tolerant design, the parameters are exchanged
between the two control channels inside the EEC via the cross channel
data link.
Using the BITE system, the EEC can detect and isolate failures. It also
allows the EEC to switch engine control from the faulty channel to the
healthy one. Depending on the nature of the failure, the EEC will
behave differently in case of:
•
•
•
•
HARDWIRED INPUTS
Most of the information exchanged between the A/C and the EEC is
transmitted over digital data buses, some signals over a single line. In
addition, some signals are hard-wired directly from the A/C to the EEC.
The Throttle Lever Angle (TLA) signal is sent directly from the throttle
resolvers to the EEC.
DUAL OUTPUTS
All the EEC outputs are double, but only the channel in control supplies
the engine control signals to the various receptors such as torque
TDTI / HAT / ATA 71-80
The EEC is equipped with a BITE system which provides maintenance
information and test capabilities via the MCDU.
Issue: 06/08
single input signal failure, there is no channel changeover, the
channel in control uses the inputs from the other channel
through the cross channel data link,
dual input signal failure, the system runs on synthesized values
of the healthiest channel,
single output signal failure, there is an automatic switchover to
the standby active channel,
complete output signal failure, there is no longer any current to
drive the torque motors or solenoids, the related component
will go to the "fail-safe" position.
FAIL-SAFE CONTROL
If a channel is faulty and the channel in control is unable to ensure
one engine function, this control is moved to a fail-safe position.
Example: if the channel is faulty and the remaining channel in control
is unable to control the Variable Stator Vane (VSV) position, the vanes
are set to the fail-safe open position.
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DUAL CHANNEL ... MAIN INTERFACES
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FADEC FAULT STRATEGY
The selected channel is one having the least significant failure.
GENERAL
SINGLE OUTPUT, SIGNAL FAILURE
If an output failure occurs, there is an automatic switchover to the
standby active channel.
The Electronic Engine control (EEC) system is dual, the two channels
are equal. Failures are classified as class 1, 2, and 3. According to the
failure class, the system can use data from the other channel, or
switch to the other channel. Faults are memorized in the system BITE
as they occur.
INPUT FAULT STRATEGY
All sensors and feedback signals are dual. Each parameter sensor as
well as feedback sensors used by each channel come from two
different sources:
T/ S ACTION
One Channel - most likely LRU failure.
COMPLETE OUTPUT SIGNAL FAILURE
In case of complete output failure there will be no current flow
through torque motors or solenoids. The associated component will be
the “FAIL-SAFE” position.
Local or cross-channel through the Cross-channel Data Link
NOTE: SOME SENSORS CAN DIRECTLY BE SYNTHESIZED BY
THE CORRESPONDING CHANNEL
NOTE: IF THE EEC POWER SUPPLY IS LOST,
COMPONENTS WILL GO INTO “FAILE-SAFE” POSITION.
SINGLE INPUT SIGNAL FAILURE
There is no channel changeover for input signal failure, as long as the
Cross-Channel Data Link is operativ.
NOTE: FAULTS ARE NOT LATCHED. AUTOMATIC RECOVERY IS
POSSIBLE.
DUAL INPUT SIGNAL FAILURE
If dual input signal failure occurs, the system runs on synthesized
values of the healthiest channel.
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CONTROL UNIT
COMPONENT FAIL SAFE STATES
COMPONENTS:
METERING VALVE
FAIL SAFE STATE:
NOTE: IF THERE IS A FAILURE OF THE THRUST REVERSER
HYDRAULIC
CONTROL
UNIT
DIRECTIONAL
VALVE WHILE THE REVERSER IS DEPLOYED, THE REVERSER
WILL REMAIN DEPLOYED.
MIN FLOW
VARIABLE
STATOR VANES OPEN
VANE ACTUATOR
2.6 BLEED ACTUATOR BLEED OPEN
(BSBV)
7TH STAGE HANDLING BLEED OPEN
BLEED VALVES
10TH STAGE HANDLING BLEED OPEN
BLEED VALVE
HPT ACC VALVE
VALVE CLOSED
LPT ACC VALVE
VALVE PARTIALLY OPEN - 45%
ACOC AIR VALVE
OPEN
10TH STAGE "MAKEUP" OPEN
AIR VALVE
FUEL DIVERTER VALVE
RETURN
VALVE
TO
FMU RETURN FLOW THROUGH FCOC
(MODE
4
OR
5)
SOLENOID
DE-ENERGIZED
TANK CLOSED (MODE 3 OR 4)
IGNITION
ON
STARTER AIR VALVE
CLOSED
P2/ T2 PROBE HEAT
ON
THRUST
REVERSER REVERSER STOWED
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COMPONENTS:
EIU SIGNALS
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DISAGREEMENT
BETWEEN TRA:
FAIL SAFE STATE:
NO ENGINE STARTING
ON GROUND: SET FORWARD IDLE
IN FLIGHT: SELECT LARGER VALUE
BUT LIMIT THIS, TO MCT
ON REVERSE: SELECT REVERSE IDLE.
NO AUTOTHRUST
ON BOTH ENGINES
NO REVERSE THRUST
MODULATED IDLE NOT AVAILABLE
CONTINUOUS IGNITION
ADC SIGNALS:
BOTH
TLA:
REVERSE:
EEC USES ENGINE SENSORS.
IN IF
REVERSER
INADVERTENTLY
DEPLOYS AND BOTH REVERSER
FEEDBACKS ARE INVALID, POWER IS
SET TO IDLE.
ON GROUND: SET IDLE
IN FLIGHT: AT TAKE OFF FREEZE LAST
VALID TLA,THEN SELECT MCT AT SLAT
RETRACTION
AUTOTHRUST CAPABILITY.
ONE TLA:
THE EEC
SENSOR.
USES
BOTH 115V AC:
NO IGNITION
THE
REDUNDANT
NO P2/T2 PROBE HEATING
BOTH 28V DC:
NO START
RUN ON ALTERNATOR ABOVE 10% N2
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MAIN INTERFACES
To perform all its tasks, the EEC interfaces with A/C computers, either
directly or via the Engine Interface Unit (EIU), which is an interface
concentrator between the A/C systems and the FADEC system. There
is one EIU for each engine, located in the avionics bay.
EEC INPUTS
The EEC receives inputs from:
•
•
•
•
•
•
the Landing Gear Control and Interface Unit (LGCIU),
the Air Data Inertial Reference Units (ADIRUs),
the Flight Control Unit (FCU),
the Environmental Control System (ECS) computers,
the Centralized Fault Display Interface Unit (CFDIU),
the cockpit engine controls including TLA, fire and anti-ice
systems.
EEC OUPUTS
The EEC sends outputs to:
•
the Bleed air Monitoring Computers (BMCs),
•
the Electronic Control Box (ECB),
•
the Flight Warning Computers (FWCs),
•
the Display Management Computers (DMCs),
•
the Flight Management and Guidance Computers (FMGCs),
•
the CFDIU,
•
the Data Management Unit (DMU), in option.
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•
FADEC PRINCIPLE (3)
auto thrust mode depending on the auto thrust function
generated by the Auto Flight System (AFS).
The FADEC also provides two idle mode selections:
The Full Authority Digital Engine Control (FADEC) system manages the
engine thrust and optimizes the performance.
• approach idle,
• minimum idle.
Approach idle is obtained when the slats are extended. Minimum idle
can be modulated up to approach idle depending on air conditioning,
engine anti-ice and wing anti-ice demands.
GENERAL
FADEC
The FADEC consists of the Electronic Engine Control (EEC) and its
peripheral components and sensors used for control and monitoring.
The EEC is in relation with the other A/C systems through the Engine
Interface Unit (EIU). The primary parameters exhaust pressure ratio,
Low Pressure Rotor Speed (N1), High Pressure Rotor Speed (N2),
Exhaust Gas Temperature (EGT), and Fuel Flow (FF) are sent directly
by the EEC to the ECAM. The secondary parameters are sent to the
ECAM through the EIU.
The FADEC provides overspeed protection for N1 and N2, in order to
prevent the engine from exceeding limits, and also monitors the EGT
and exhaust pressure ratio.
ENGINE SYSTEMS
The FADEC provides optimal engine operation by controlling the:
•
•
EIU
Each EIU, located in the avionics bay, is an interface concentrator
between the airframe and the corresponding FADEC located on the
engine. There is one EIU for each engine.
POWER MANAGEMENT
The FADEC provides automatic engine thrust control and thrust
parameter limit computation. The thrust is computed according to the
exhaust pressure ratio in normal mode or N1 in back-up mode. In fact,
when the exhaust pressure ratio mode is no longer operational the
FADEC automatically reverts to the N1 alternate control mode. The
FADEC manages power according to two thrust modes:
•
ENGINE LIMITS
FF,
compressor airflow and turbine clearance.
IGNITION AND STARTING
The FADEC controls the engine start sequence. It monitors exhaust
pressure ratio, N1, N2, and EGT parameters and can abort or recycle
an engine start. The FADEC controls the ignition and starting in
automatic or manual mode when initiated from the ENGine start or
ENGine MANual START panels.
manual mode depending on the Throttle Lever Angle (TLA),
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GENERAL ... THRUST REVERSER
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THRUST REVERSER
The FADEC supervises the thrust reverser operation entirely. In case of
inadvertent deployment, the FADEC will command the automatic
restowing sequence.
NOTE: DURING REVERSE OPERATION THE
CONTROLLED AS A FUNCTION OF N1.
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THRUST IS
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DIGITAL OUTPUTS
EEC INTERFACES (3)
Each channel has 2 output ports and each bus has a separated line
driver, i.e. A1, A2, B1, B2, in such a way that propagation of failures is
prevented.
GENERAL
In order to provide a full range of engine control and monitoring, the
Electronic Engine Control (EEC) interfaces with the following sub
systems:
•
•
•
•
air data computers, which transmit air data for engine control,
Engine Interface Unit (EIU), which concentrates A/C signals
and transmits them to the EEC,
for cockpit indication and control, and autothrust control, the
EEC sends signals to computers,
engine sensors and controls.
DIGITAL INPUTS
Inputs of each channel are isolated in order to prevent failure
propagation. Each channel receives inputs for both the Air
Data/Inertial Reference Units (ADIRUs) and the EIU. The EEC performs
a fault detection on its input parameters by determining if they are
valid. This check is made by applying a range of tests to each input.
Faults detected by the EEC are annunciated and recorded for
maintenance or crew action, if required. The ADIRUs send air data
parameters to the EEC for engine control. The EEC performs validation
tests and selection logic between air data signals from the ADIRUs and
the engine sensors. ADIRU data is preferred over engine data. The air
data used to validate Altitude (P0), Total Air Temperature (TAT),
TOTal PRESSure, mach for the power management and engine
controls are:
•
•
TAT, TOTal PRESSure, P0, mach number, from the ADIRUs,
P2, T2 and P0 from engine sensors.
TDTI / HAT / ATA 71-80
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The EEC output buses provide:
•
•
•
•
•
engine control parameters,
engine condition monitoring parameters,
EEC status and fault indication,
engine rating parameters,
Full Authority Digital Engine Control (FADEC) system
maintenance data. ´
The EEC transmits outputs continuously on all buses in normal
operation, irrespective of whether the given channel is in active control
or not. The parameter values on the 2 output buses are normally
identical as long as the cross channel data link is functioning.
NOTE: CHANNEL SWITCHOVER DOES NOT AFFECT THE
OUTPUT DATA OF THE EEC.
DISCRETE/ANALOG SIGNALS
The EEC receives the Throttle Resolver Angle (TRA) signals, the
autothrust instinctive disconnect switch signals and engine sensor
signals. They are validated by the EEC. The EEC also sends signals to
the engine controls. The resolvers of the Throttle Control Unit (TCU)
receive an excitation current from the EEC. The EEC checks the range
limits, the rate limits and performs an interface fault detection test.
NOTE: 1 DEGREE TLA CORRESPONDS TO 1.9 DEGREE TRA.
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GENERAL ... DISCRETE/ANALOG SIGNALS
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BLOCK DIAGRAM
Here are the interfaces between the EEC and the A/C in the form of a
block diagram.
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BLOCK DIAGRAM
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DATA ENTRY PLUG
The Data Entry Plug (DEP) provides discrete inputs to the EEC.
Located to the Junction 6 of the EEC it provides unique engine data to
channel A and B.
NOTE: THE PART NUMBER IS WRITTEN ON THE DEP. THE
PARTNUMBER CAN ALSO BE FOUND ON THE ENGINE DATA
PLATE, WHICH IS LOCATED AT THE LEFT HAND SIDE OF THE
FAN CASE.
The data transmitted by the DEP is:
ƒ
ƒ
ƒ
EPR Modifier (Used for power setting)
Engine Rating
Engine Serial No.
NOTE: IF THE DATA INPUTS OF THE DATA ENTRY PLUG J6 ARE
LOST, THEN AN AUTOMATIC RE-VISION FROM EPR MODE TO
UNRATED N1 MODE OCCURS.
DATA ENTRY PLUG MODIFICATION
DESCRIPTION
The DEP links the coded data inputs through the EEC by the use of
shorting jumper leads which are used to select the plug pins in a
unique combination.
During a life of an engine, it may be necessary to change the DEP
configuration either during incorporation of Service Bulletins or after
engine overhaul , when the EPR modifier code may need to be
changed. This is accomplished by changing the configuration of the
jumper leads in accordance with the relevant instructions.
During removal/ replacement of the DEP it is necessary to use an EEC
Harness Wrench as it is imperative that the connectors are fight. On
fitment of the DEP to the EEC align the main key of the connector with
the EEC and hand tighten the connector. Then using the EEC Harness
Wrench torque tighten the DEP connector to 32 lbf in.
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DATA ENTRY PLUG AND EEC PORTS
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HARNESS (ELECTRICAL) AND PRESSURE CONNECTIONS
Two identical, but separate electrical harnesses provide the input/
output circuits between the E.E.C. and the relevant sensor/control
actuator, and the aircraft interface. The harness connectors are
“keyed” to prevent misconnection.
NOTE: SINGLE PRESSURE SIGNALS ARE DIRECTED TO
PRESSURE TRANSDUCERS - LOCATED WITHIN THE E.E.C. THE PRESSURE TRANSDUCERS THEN SUPPLY DIGITAL
ELECTRONIC SIGNALS TO CHANNELS A AND B.
J11
Engine D211 P
Rear Face
J5
Engine D205P
J6
Data Entry Plug
J7
E.B.U. 4000 KS13
J8
Engine D208P
J19
Engine D209P
J10
Engine D210P
The following pressures are sensed:
Pamb
ambient air pressure (fan case sensor)
Pb
burner pressure (air pressure) P3/ T3 probe
P2
pressure (P2/ T2 fan inlet probe)
P2.5
booster stage outlet pressure
P5 (P4.9)
L.P Turbine exhausts pressure (P5 (P4.9) rake)
P12.5
fan outlet pressure (fan rake)
ELECTRICAL CONNECTIONS
FRONT FACE
Harness Connector Plug Identification
J1
E.B.U. 4000 KSA
J2
Engine D202P
J3
Engine D203P
J4
Engine D204P
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EEC CONNECTION PORTS
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FADEC SENSORS
FADEC LRUS SENSORS
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T4.9 (EGT) Sensor
N1 Sensor
N2 Sensor
Engine Oil Temperature Sensor
P2/T2 Sensor
P3/T3 Sensor
P4.9 (P5)
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ENGINE SENSOR LOCATIONS
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OVERSPEED
FAILURES AND REDUNDANCY
Improved reliability is achieved by utilising dual sensors dual feedback.
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Dual sensors are used to supply all EEC inputs except
pressures, (single pressure transducers within the EEC provide
signals to each channel -A and B) .
The EEC 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 if
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 pre-determined
hierarchy decides, which channel is more capable of control
and utilises that channel.
In the event of loss of both channels, or loss of electrical
power, the systems are designed to go to their failsafe
positions.
Overspeed protection logic consists of overspeed limiting loops, for
both the low and high speed rotors, which act directly upon the fuel
flow command. Supplementary electronic circuitry for overspeed
protection is also incorporated in the EEC. Trip signals for hardware
and software are combined to activate a torque motor, which drives a
separate overspeed valve in the fuel metering unit to reduce fuel flow
to a minimum value. The engine can be shut down to reset the
overspeed system to allow a restart if desired.
ENGINE SURGE
Engine surge is detected by a rapid decrease, in burner pressure or
the value of rate of change of burner pressure, which indicates that
surge varies with engine power level. Once detected, the EEC will reset
the stator vanes by several degrees in the closed direction, open the
booster 7th and 10th stage bleeds, and lower the maximum Wf/ Pb
schedule. Recovery of burner pressure to its steady state level or the
elapse of a timer will release the resets on the schedules and allow the
bleeds to close.
ENGINE LIMITS PROTECTION
GENERAL
The FADEC prevents inadvertent overboosting of the expected rating
(EPR limit and EPR target) during power setting. It also prevents
exceedance of rotor speeds (N1 and N2) and burner pressure limits. In
addition, the FADEC unit monitors EGT and sends an appropriate
indication to the cockpit display in case of exceedance of the limit. The
FADEC unit also provides surge recovery.
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FALIURES AND REDUNDANCY
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•
EIU INTERFACES (3)
fault detection logic carried out by an internal BITE and
transmission of the result to the Centralized Fault Display
System (CFDS).
GENERAL
INPUTS
There are 2 Engine Interface Units (EIUs), one for each engine. The
EIU is an interface concentrator between the A/C and the Full
Authority Digital Engine Control (FADEC) system.
The EIU receives the following:
The main functions of the EIU are:
•
•
•
•
to concentrate data from the cockpit panels,
to ensure the segregation of the 2 engines,
to provide the Electronic Engine Control (EEC) with an electrical
power supply,
to give the necessary logic and information from the engine to
the aircraft systems.
EIU COMPOSITION
The EIU is composed of 5 main parts:
•
•
•
•
•
discrete and analog inputs,
digital inputs,
digital outputs,
discrete outputs,
power supply switching.
The EIU performs the following:
acquisition of information,
transmission of messages,
logics for Low Oil Pressure (LOP) and Auxiliary Power Unit
(APU) boost,
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•
discrete signals which are of the ground/open circuit type,
analog inputs which are of the differential type with a working
range of 1 to 9 volts,
digital inputs on ARINC 429 lines.
OUTPUTS
The EIU sends the following:
•
•
digital output signals on ARINC 429 buses,
discrete signals which are of the 28V DC/open circuit or
ground/open circuit types.
EEC INTERFACE
EIU FUNCTIONS
•
•
•
•
•
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The EIU receives 2 ARINC 429 output data buses from EEC channels A
and B and it uses data from the channel in control. If some data is not
available from the control channel, data from the other channel is
used. The EIU looks at particular engine data on the EEC digital data
flow to interface this with other aircraft computers and with cockpit
panels for control and monitoring. The EIU sends information, coming
from all aircraft computers, which have to communicate with the EEC,
through an output ARINC 429 data bus. The EIU does not deal with Air
Data/Inertial Reference Unit (ADIRU) and thrust lever information as
they communicate directly with the EEC.
NOTE: THERE IS NO DATA FLOW DURING THE EIU INTERNAL
TEST OR INITIALIZATION.
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GENERAL ... POWER SUPPLY
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signal, the BMC closes the pneumatic system Pressure Regulating
Valve (PRV).
ECS INTERFACE
The EIU receives two input buses from the Environmental Control
System (ECS) primary and secondary computers of the Zone Controller
(ZC). The ECS determines the various air bleed configurations
according to logics of the air conditioning, wing anti-ice and nacelle
anti-ice. This information is transmitted by the EIU to the EEC to
compute the bleed air demand required at the engine customer bleed
ports.
The EIU receives and generates all starting signals from the cockpit
engine panels. Therefore engine starting is not possible in case of EIU
failure. The control panels provide the EIU with the following signals:
• engine start mode selector position,
• master lever position,
• manual start P/B.
The EIU provides all starting signals to the EEC and to the ENGine
start panel FAULT light.
OTHER INTERFACES
The EIU also receives other signals from various aircraft systems for
control and monitoring purposes. The EIU also generates signals for
various aircraft systems.
POWER SUPPLY
The EIU receives the following:
•
•
28V DC for its own power supply and for the FADEC power
supply,
115V AC for engine ignition system power supply.
NOTE: IN CASE OF EIU FAILURE, THE POWER SUPPLY FOR
FADEC AND IGNITION IS PRESERVED (FAIL SAFE POSITION).
CFDS INTERFACE
The Centralized Fault Display Interface Unit (CFDIU) communicates
with the BITE memory of the EIU and with the BITE memory of the
EEC via the EIU. The EIU interfaces with the CFDIU to generate the
EIU and the FADEC fault messages on the MCDU. To access the
FADEC menu function, the CFDS interrogates the EEC BITE memory
through the EIU.
BMC INTERFACE
There is an interface between the EIU and the pneumatic system
Bleed Monitoring Computer (BMC) for engines 1 and 2. During engine
start, the EIU generates a ground signal for the BMC when the start
valve moves away from the closed position. On receipt of this ground
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THIS PAGE INTENTIONALLY LEFT BLANK
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FMGS
The AutoTHRust (A/THR)
Management and Guidance
sends all command signals
Flight Control Unit (FCU) and
function is provided by the Flight
System (FMGS). Therefore, the FMGS
for engine power management via the
both EIUs.
The FMGS assumes the following functions:
•
•
•
•
•
computation of Engine Pressure Ratio (EPR) target,
selection of A/THR modes,
alpha floor protection,
flexible takeoff,
A/THR engagement.
FEEDBACK
The EEC directly sends specific feedback outputs to the FMGS without
going through the EIU. The EIU also receives this data but does not
transmit it to FMGS. The EEC feedback output data to Flight
Management and Guidance Computer (FMGC) is as follows:
•
•
•
•
•
•
•
•
•
•
•
•
Thrust Resolver Angle (TRA),
EPR command,
EPR actual,
EPR target in feedback,
EPR limit,
EPR max,
EPR reference throttle,
thrust mode,
Fuel Flow (FF),
bleed configuration in feedback,
engine rating identification,
flexible temperature in feedback.
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FMGS - FEEDBACK
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EIU CFDS DISCRETE OUTPUTS SIMULATION
The Purpose of this Menu is to Simulate some Engine Interface Unit
(EIU). Discrete Outputs by Setting their Status to 0 or 1
SIMULATION: “LOP GND 1”
To simulate “OIL LOW PRESS & GND” for the following systems
through the MCDU:
PHC1, PHC3, WHC1, AEVC, DFDR and CVR.
WARNING:
CAUTION:
The DISCRETE OUTPUT SIMULATION can operate systems and
components without special indication on the MCDU. Make always sure
that the working areas are clear! For the simulation refer to AMM
73-25-34, (TASK 73-25-34-860-041).
REMOVE THE PROTECTIVE COVERS FROM THE PROBES
BEFORE YOU DO THE TEST.
The Discrete Outputs are listed on two Pages, one for the Positive
Type and one for the Negative Type.
SIMULATION: “APU BOOST”
To simulate an APU BOOST command through the MCDU.
Push the line key adjacent “APU BOOST” discrete output status:
“APU BOOST” becomes “1” and the EIU sends the APU BOOST
command to the 59KD ECB.
If the line key adjacent to LOP is pushed, LOP GND1 discrete output
status becomes GND1 “0”.
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The PHC1 (3) commands a low probes heating level.
The WHC1 commands a low captain windshield heating level.
The CVR and DFDR are switched on.
NOTE: WHEN "LOP GND1" IS SIMULATED TO "0" THE HORN
WILL BE INHIBITED IN CASE OF LOW AVIONIC BAY EXTRACT
AIRFLOW.
APU. BOOST 1 simulates a not closed starter air valve. The APU is
boosted (if running)
APU BOOST 2 simulates a energized starter air valve solenoid.
SIMULATION: “FAULT”
To simulate a disagree between the position and the command of the
HP fuel valve through the WDU the line key adjacent to FAULT discrete
output status is must be pushed. The FAULT becomes “1” and the
FAULT legend of the 5KS1(2) annunciator light comes on.
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DISCRETE OUTPU SIMULATION
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SIMULATION: “LOP GND 2”
To simulate "OIL LOW PRESS & GND" for the following systems
through the MCDU:
Blue/ yellow main hydraulic pressure power warning indicating WHC2,
PHC2, green main hydraulic PWR RVSR indicating, FCDC1, FCDC2.
When the line key adjacent to LOP "LOP GND 2 " discrete output
status becomes GND2 "0".
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“B (Y) ELEC PUMP LO PR” warning message is no longer
inhibited.
The PHC2 commands a low probes heating level.
The WHC2 commands a low windshield (F/O) heating level.
The 3DB1 and 3DB2 rain repellent valve opening is authorized
NOTE: THE "LOP GND2" DISCRETE IS USED TO INHIBIT THE
FLIGHT CONTROL SYSTEM TEST THROUGH THE CFDS. ACCESS
TO THIS MENU IS PROHIBITED BY THE CFDS ARCHITECTURE
AS LONG AS YOU WORK ON THE EIU DISCRETE OUTPUTS
MENU.
status FUEL PN becomes “1” and the zone controller 8HK will receive
the HP FUEL VALVE 1(2) open condition.
NOTE: THE ZONE CONTROLLER USES THE HP FUEL VALVE
POSITION TO ELABORATE THE BLEED STATUS ON LABEL 061
AND SENDS IT TO THE EEC THROUGH THE EIU (LABEL 030).
THE BLEED STATUS CAN ONLY BE MODIFIED BY THIS INPUT
IF THE PRV OPENS (ENGINE RUNNING).
SIMULATION “PACKS OFF”
To simulate the PACK FLOW control valve closure command through
the MCDU push the line key adjacent to “PACKS OFF” discrete output
status. PACKS OFF becomes "1" and the PACK FLOW control valve
closure solenoid is energized.
NOTE: THE PACK FLOW CONTROL VALVE 1(2) REQUIRES A
MUSCLE AIR PRESSURE TO OPEN.
SIMULATION: “T/R INHIB”
To simulate the authorization of the T/R directional control valve
solenoid closure (through the 14KS1 (2) relay) through the MCDU.
When the line key adjacent to T/R is pushed, "T/R INHIB" discrete
output status INHIB becomes “1” and the 14KS1 (2) inhibition relay is
energized, authorizing the directional control valve solenoid
energisation
SIMULATION: “HP FUEL PN”
To simulate a HP FUEL VALVE 1(2) in open position through the
MCDU. Push the line key adjacent to HP “HP FUEL PN” discrete output
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DISCRETE OUTPU SIMULATION
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SIMULATION OF “N2 > IDLE”
To simulate “N2 > IDLE” for the following systems:
XMR radio altimeter 25A
Blue main hydraulic power
WARNING: MAKE SURE THAT THE TRAVEL RANGES OF THE
FLIGHT CONTROL SURFACES ARE CLEAR BEFORE YOU
PRESSURIZE/ DEPRESSURIZE A HYDRAULIC SYSTEM.
Push the line key adjacent to N2. N2 > IDLE DISCRETE OUTPUT
becomes “1”> IDLE.
The electric pump of the blue hydraulic system start and the blue
hydraulic system is pressurized (approximately 3000PSI)
NOTE: THE N2 > IDLE DISCRETE IS USED TO INHIBIT THE
"RAMP TEST” OF THE
RADIO ALTIMETER 1(2). Access to radio altimeter RAMP TEST menu is
prohibited by the CFDS architecture as long as you work on the EIU
DISCRETE OUTPUTS menu.
SIMULATION OF “TLA > MCT”
To simulate “TLA > MCT” for the following systems
AEVC, PACK CONTROLLERS
CABIN PRESSURE CONTROLLERS.
Push the line key adjacent to TLA "TLA > MCT” discrete output status
> MCT becomes “1”.
On the ECAM PRESS page check that the inlet and extract skin air
valves close.
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DISCRETE OUTPU SIMULATION
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ENGINE THRUST MANAGEMENT (3)
BASIC INFORMATION
Some basic information about engine thrust management is shown in
this module.
PREDICTED EPR
The predicted Engine Pressure Ratio (EPR) is indicated by a white
circle on the EPR indicator and corresponds to the value determined by
the Throttle Lever Angle (TLA).
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BASIC INFORMATION - PREDICTED EPR
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EPR SETTING REQUIREMENTS
EPR
The EEC uses closed loop control based on EPR or, if EPR is
unobtainable, on N1.
Under EPR control, 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
(FF) which is summed with the measured fuel flow (FF actual) to
produce the FF error.
The FF error is converted to a current (I), which is sent to the dual
torque motor. The torque motor repositions the Fuel Metering Valve
(FMV) to change the fuel flow.
The inputs required for EPR control are:
ƒ Ambient temperature (Tamb)
ƒ Engine air inlet temperature (T2 - Altitude (ALT)
ƒ Mach number (Mn)
ƒ Throttle Resolver Angle (TRA)
ƒ Service Bleeds
It is possible to re-select the primary control mode (EPR) through the
N1 mode P/B switch following an automatic reversion to rated or
unrated N1 mode. If the fault is still present, the EEC will remain in its
current thrust setting mode. If the fault is no longer present, the EEC
will switch to the primary control mode (EPR). If the fault later
reoccurs, reversion back to N1 mode (rated or unrated) will result.
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EPR SETTING
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THRUST LIMIT MODES
The throttle levers are used as thrust limit mode selectors. Depending
on the throttle lever position, a thrust limit mode is selected and
appears on the upper ECAM display. If the throttle levers are set
between two detent points, the upper detent will determine the thrust
limit mode. The thrust limit modes are:
•
•
•
•
•
CL: Climb,
FLX: Flexible take-off,
MCT: Maximum Continuous Thrust,
TOGA: Take Off/Go Around.
MREV: Maximum Reverser
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BASIC INFORMATION - THRUST LIMIT MODES
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EPR LIMIT
For each thrust limit mode selection, an EPR limit is computed
according to the Air Data Reference (ADR) and appears on the upper
ECAM display beside the thrust limit mode indication.
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BASIC INFORMATION - EPR LIMIT
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EPR TARGET
For its auto thrust function, the Flight Management and Guidance
System (FMGS) computes an EPR target according to ADR and engine
parameters and sends it to the Electronic Engine Control (EEC).
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BASIC INFORMATION - EPR TARGET
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EPR COMMAND
The EPR command, used to regulate the fuel flow, is the Flight
Management and Guidance Computer (FMGC) EPR target when the
auto thrust function is active. When the auto thrust function is not
active, the EPR command is the EPR corresponding to the TLA. EPR
command is either:
•
•
EPR target,
EPR corresponding to TLA.
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BASIC INFORMATION - EPR COMMAND
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ACTUAL EPR
The actual EPR is the actual value given by the ratio of the LP turbine
exhaust pressure (P4.9) to the engine inlet pressure (P2). The actual
EPR is displayed in green on the EPR indicator. The actual EPR signal
is also compared to the EPR command.
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BASIC INFORMATION - ACTUAL EPR
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AUTOTHRUST CONTROL MODE
The auto thrust function is engaged when the A/THR P/B is on.
The auto thrust engages:
•
•
when the A/THR P/B is pressed in,
at take-off power application.
AUTOTHRUST ACTIVE
When engaged, the auto thrust function becomes active when the
throttle levers are set to the Climb detent after take-off. The EPR
command is the FMGC EPR target. The auto thrust function is active
when the throttle levers are set between IDLE and CL, including CL,
with 2 engines running.
NOTE: THE AUTO THRUST FUNCTION ACTIVE RANGE IS
EXTENDED TO MCT IN THE CASE OF SINGLE ENGINE
OPERATION. WHEN THE THROTTLE LEVERS ARE SET
BETWEEN TWO DETENT POINTS, THE EPR COMMAND IS
LIMITED BY THE THROTTLE LEVER POSITION.
NOTE: IN ALPHA FLOOR CONDITION THE AUTO THRUST
FUNCTION BECOMES ACTIVE AUTOMATICALLY. THE EPR
TARGET IS TOGA.
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AUTOTHRUST CONTROL MODE - AUTOTHRUST ACTIVE
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AUTOTHRUST NOT ACTIVE
When engaged, the auto thrust function becomes inactive when the
throttle levers are set above CL with 2 engines running. The EPR
command corresponds to the TLA.
NOTE: THE AUTO THRUST FUNCTION IS NOT ACTIVE ABOVE
MCT IN CASE OF SINGLE ENGINE OPERATION. THE AUTO
THRUST FUNCTION IS DISENGAGED WHEN THE THROTTLE
LEVERS ARE SET AT THE IDLE STOP.
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AUTOTHRUST CONTROL MODE - AUTOTHRUST NOT ACTIVE
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MANUAL CONTROL MODE
When the auto thrust function is not engaged, the EEC processes the
EPR command signal according to the TLA. In manual mode the auto
thrust is not engaged.
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MANUAL CONTROL MODE
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BACK-UP N1 MODE
Here are presented two back-up N1 modes:
•
•
rated N1 mode,
unrated N1 mode.
RATED N1 MODE
In case of EPR sensor failure, i.e. P2 or P4.9, the EEC automatically
reverts to the rated N1 mode. On the ECAM, the EPR indicator is
crossed amber. The EEC uses TLA, ADR and T2. T2 is used in N1 rated
mode to limit the engine thrust. To select N1 mode on both engines,
the Engine N1 MODE P/Bs must be pressed in.
NOTE: THE AUTO THRUST FUNCTION IS NOT AVAILABLE IN
N1 MODE. WHEN THE N1 MODE P/BS ARE PRESSED IN, THE
LIGHTS COME ON AND A SIGNAL IS SENT TO THEIR
CORRESPONDING FULL AUTHORITY DIGITAL ENGINE
CONTROL (FADEC) SYSTEM TO CONFIRM OR TO FORCE THE
N1 MODE SELECTION. ON THE ECAM, THE EPR INDICATOR IS
CROSSED AMBER AND THE N1 MODE LIMIT IS DISPLAYED
INSTEAD OF THE EPR LIMIT. ON THE N1 INDICATOR, THE
PREDICTED N1, I.E. THE WHITE CIRCLE, REPLACES THE
PREDICTED EPR AND AN AMBER INDEX APPEARS TO
INDICATE THE TAKE-OFF N1 LIMIT.
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BACK-UP N1 MODE - RATED N1 MODE
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UNRATED N1 MODE
T2 is no longer available; the EEC reverts from the EPR mode to the
unrated N1 mode. In this case the N1 limitation is no longer
computed. The N1 command is directly related to the TLA.
NOTE: IN UNRATED N1 MODE, THERE IS NO LONGER AN
ENGINE PROTECTION AGAINST OVER BOOST, E.G. GOAROUND.
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BACK-UP N1 MODE - UNRATED N1 MODE
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IDLE CONTROL
ƒ
ƒ
ƒ
ƒ
ƒ
Minimum idle (56 % - 60% N2) is corrected for ambient temp
>30°C, then N2 will increase.
Approach idle (approx. 70% N2)
It varies as a function of Total Air Temperature (TAT) and
altitude. This idle speed is selected to ensure sufficiently short
acceleration time to go around thrust and is set when the
aircraft is in an approach configuration. (Flap Lever Position“NOT UP”)
Reverse Idle (approx. 70% N2) = Approach Idle + 1000 RPM
FADEC sets the engine speed at reverse idle when the throttle
is set in the reverse idle detent position
Bleed Idle = Bleed demand.
Bleed Idle command will set the fuel flow requested for
ensuring correct aircraft ECS system pressurization wing ant ice
and engine ant ice pressurization (P/B-“ON” or valves not
closed)
HMS Idle (Min Idle - Approach Idle)
For conditions where the compensated fuel temperature is
greater than 140 deg. C, the heat management control logic
calculates raised idle speed. (in flight and on ground!)
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IDLE SETTINGS
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ƒ a reverse latching lever.
The thrust lever is linked to a mechanical rod. This rod drives the input
lever of the throttle control artificial feel unit (Mechanical Box).
ENGINE CONTROLS (3)
THROTTLE CONTROL SYSTEM
REVERSE THRUST LATCHING LEVER
GENERAL
To obtain reverse thrust settings, the revers thrust latching lever must
be lifted.
The throttle control system consist of
ƒ the throttle control lever
ƒ the throttle control artificial feet unit (Mechanical Box)
ƒ the thrust control unit
ƒ the electrical harness.
The design of the throttle control is based upon a fixed throttle
concept this means that the throttle control levers are not servo
motorized.
A mechanical cam design is provided to allow reverse thrust selection
when thrust lever is at forward idle position.
The thrust lever has 3 stops at the pedestal and 3 detents in the
artificial feel unit:
ƒ
ƒ
ƒ
0° STOP = FWD IDLE THRUST
20° STOP = FULL REVERSE THRUST
45 STOP = MAX.TAKE OFF THRUST
ƒ
ƒ
ƒ
DETENT 1 = (REVERSE) IDLE THRUST
DETENT 2 = MAX.CLIMB (ALSO CRUISE SELECTION)
DETENT 3 = MAX. CONTINOUS (FLEX TAKE OFF THRUST)
THRUST CONTROL UNIT
The Thrust Control Unit contains two resolvers, each of whom sends
the thrust lever position to the Electronic, Engine Control The
extraction current for the resolvers is provided by the EEC.
AUTTHRUST DISCONNECT PUSBUTTON
The autothrust instinctive disconnect pushbutton can be used to
disengage the autothrust function.
THRUSTLEVERS
GENERAL
The thrust levers comprises
ƒ
ƒ
a thrust lever which incorporates stop devices and autothrust
instinctive disconnect pushbutton switch
a graduated fixed sector
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THROTTLE CONTROL SYSTEM
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BUMP RATING PUSH BUTTON
THIS PUSH BUTTONS ARE OPTIONAL EQUIPMENT.
In some cases the throttle control levers are provided with "BUMP"
rating push buttons, one per engine. This enables the EEC to be rerated to provide additional thrust capability for use during specific
aircraft operations.
BUMP RATING DESCRIPTION
The takeoff bump ratings can be selected, regardless of the thrust
lever angle, only in the EPR mode when the airplane is on the ground.
The bump ratings, if available, are selected by a push button located
on the thrust lever. Actuation of the switch will generate a digital
signal to both EEC’s via the EIU. The maximum take-off rating will then
be increased by the pre-programmed delta EPR provided the airplane
is on the ground. The bump ratings can be de-selected at anytime by
actuating the bump rating push button as long as the airplane is on
the ground and the thrust lever is not in the maximum takeoff (TO)
detent. In flight, the bump ratings are fully removed when the thrust
lever is moved from the TO detent to, or below, the MCT detent. The
bump rating is available in flight (EPR or rated N1 mode) under the
following conditions.
ƒ Bump rating initially selected on the ground.
ƒ TO/GA thrust lever position set.
ƒ Airplane is within the takeoff envelope.
The bump rating is a non-standard rating and is only available on
certain designated operator missions. Use of the bump rating must be
recorded. This information is for tracking by maintenance personnel.
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PUMP RATING PUSHBUTTON
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ARTIFICIAL FEEL UNIT (MECANICAL BOX)
The Throttle control artificial feet unit is located below the cockpit
center pedestal, this artificial feel unit is connected to engine 1(2)
throttle control lever and to the engine 1(2) throttle control unit by
means of rods. The artificial feel unit is a friction system, which
provides a load feedback to the throttle control lever. This artificial feet
unit comprises two symmetrical casings, one left and one right. Each
casing contains an identical and independent mechanism. Each
mechanism is composed of:
• a friction brake assembly
• a gear assembly
• a lever assembly
• a bellcrank assembly
Throttle lever travel is transmitted to the to the artificial feel unit and
to the throttle control unit. The linear movement of the throttle levers
is transformed into a rotary movement at the bellcrank, which turns
about the friction brake assembly shaft. This movement rotates a
toothed quadrant integral with the shaft. This toothed quadrant causes
inverse rotation of a gear equipped with a disk, which has four detent
notches. Each notch corresponds to a throttle lever setting and is felt
as a friction point at the throttle levers.
THROTTLE CONTROL UNIT
a pin device for rigging the resolvers and potentiometers
a safety device which leads the resolvers outside the normal
operating range in case of failure of the driving device two
output electrical connectors.
The input lever drives two gear sectors assembled face to face. Each
sector drives itself a set of one resolver and three potentiometers.
ƒ
ƒ
RELATION BETWEEN TRA AND TLA:
The relationship between the throttle lever angle and throttle resolver
angle (TRA) is linear and: 1 deg. TLA = 1.9 TRA. The accuracy of the
throttle control unit (error between the input lever position and the
resolver angle) is 0.5 deg. TRA.
The maximum discrepancy between the signals generated by the two
resolvers is 0.25 deg. TRA. The TLA resolver operates in two
quadrants the first quadrant serves for positive angles and the fourth
quadrant for negative angles. Each resolver is dedicated to one
channel of the EEC and receives its electrical excitation from the EEC.
The EEC considers a throttle resolver angle value:
ƒ less than -47.5 deg. TRA or
ƒ greater than 98.8 deg. TRA as resolver position signal failure.
The EEC incorporates a resolver fault accommodation logic. This logic
allows engine operation after a failure or a complete loss of the
throttle resolver position signal.
THE THROTTLE CONTROL UNIT COMPRISES:
ƒ an input lever
ƒ mechanical stops which limit the angular range
ƒ 2 resolvers whose signals are dedicated to the EEC (one
resolver per channel of the EEC)
ƒ 6 potentiometers fitted three by three. Their signals are used
by the flight control system
ƒ a device that drives the resolver and the potentiometer
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ARTIFICIAL FEEL UNIT/ THROTTLE CONTROL UNIT
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points as the oil pressure sensor. This allows for cockpit monitoring of
low oil pressure. The engine oil temperature is measured in the
combined scavenge line to the oil tank.
OIL SYSTEM D/O (3)
GENERAL
The lubrication system is self-contained and thus requires no airframe
supplied components other than certain instrumentation and remote
fill and drain port disconnectors on the oil tank. These ports are used
to refill the oil tank promptly and, precisely by allowing the airlines to
quick-connect a pressurized oil line and a drain line. It is a hot tank
system that is not pressure regulated. Oil from the oil tank enters the
one stage pressure pump and the discharge flow, is sent directly to the
oil filter. A coarse cleanable filter is employed. The oil then is piped
through the air cooled oil cooler and the fuel cooled oil cooler which
are part of the Heat Management System (HMS) which ensures that
engine oiI, IDG oil and fuel temperatures are maintained at acceptable
levels, to the bearings. Except for the No.3 bearing damper and the
No.4 bearing compartment, the pressure supplied to each location is
controlled by a restrictor. There is a "last chance" strainer at the entry
of each compartment to prevent blockage by any debris/ carbon flakes
in the oil. The scavenge oil is then piped, either directly or through the
de-oiler to the 5 stage scavenge pumps. There is a disposable
cartridge type scavenge filter at the outlet of the scavenge pumps
before returning to the oil tank. A valve allows oil to bypass the
scavenge filter when the filter differential pressure exceeds 20 psi. A
differential pressure warning switch set at 12 psi gives cockpit
indication of impending scavenge filter bypass.
The No.4 bearing two position scavenge valve is operated
pneumatically by tenth stage air and controls vented air flow from the
bearing compartment in response to specific levels of engine thrust
setting. At engine idle power, the valve opens to provide the maximum
area for scavenge flow. At higher power, the valve closes to a reduced
area, which provides, adequate pressure in the Nobel bearing
compartment to protect the seals by maintaining low pressure
differentials across compartment walls and minimizes air leakage into
the bearing chamber.
The scavenge valve pressure transducer senses the pressure present
in the scavenge line upstream of the scavenge valve and supplies a
signal to the EIU.
A pressure relief valve at the filter housing limits pump discharge
pressure to approximately 450 psi to protect downstream components.
The oil pressure is measured as a differential between the main supply
line pressure, upstream of any restrictors, and the pressure in the
Nobel bearing compartment scavenge line, upstream of the two
position scavenge valve. A low pressure warning switch, which is set
for 60 psi, is provided in the main oil line before the bearing
compartments and after the ACOC and FCOC at the same tapping
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OIL SUPPLY CIRCUIT
The oil from the tank passes through the pressure pump and the
pressure filter to lubricate the bearing compartments and also the
gearboxes. Between the supply line and the No.4 bearing scavenge
line, an oil Low Pressure (LP) switch and an oil pressure transmitter
are provided for indication and monitoring. The oil tank content is
measured through an oil quantity transmitter.
NOTE: THE INSTALLATION OF AN OIL TEMPERATURE SENSOR
FOR THE HEAT MANAGEMENT SYSTEM. THERE IS NO
PRESSURE REGULATOR, SO THE OIL PRESSURE VARIES WITH
N2. A PRESSURE RELIEF VALVE IS PROVIDED TO LIMIT THE
PRESSURE DURING COLD STARTS. THE COLD START
PRESSURE RELIEF VALVE OPENS AT 450 PSI.
NOTE: THAT THE ANTI-SIPHON SYSTEM PREVENTS THE
SIPHONING OF THE OIL FROM THE TANK TO THE GEARBOXES
WHEN THE ENGINE IS STATIC.
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OIL SCAVENGE CIRCUIT
Six scavenge pumps suck the oil from bearing compartments, and
gearboxes. The pumps then return this oil to the tank through the
scavenge filter. The main scavenge line is provided with:
•
•
•
a master chip detector for inspection,
an oil temperature sensor for indication,
an oil differential pressure switch for monitoring and warning to
the cockpit when the scavenge filter is clogged.
If the scavenge filter becomes clogged a differential pressure switch
sends a signal to the ECAM and eventually the by-pass valve will open.
The ECAM is activated when the differential pressure is greater than
12 PSI. The by-pass valve opens at 20 PSI.
• engine oil temperature
o MINIMUM STARTING: ..................... -40°C
o MIN.PRIOR EXCEEDING IDLE: ......... -10°C
o MIN. PRIOR TAKE OFF: .................. 50°C
o MAX CONTINIOUS: ........................ 155°C
o MAX TRANSIENT. ........................... 165°C
• oil tank contents 25 US quarts
In addition warnings may be given for the following non normal
conditions:
•
•
•
low oil pressure
o RED LINE LIMIT: 60 PSI
o AMBER LINE LIMIT: 80 PSI
scavenge filter clogged.
No. 4 compartment scavenge valve inoperative.
NOTE: THE SCAVENGE LINE OF THE NO. 4 BEARING
COMPARTMENT IS CONTROLLED BY THE NO. 4 BEARING
SCAVENGE VALVE. EACH SCAVENGE LINE IS EQUIPPED WITH
A STRAINER AND A MAGNETIC CHIP DETECTOR TO PROTECT
THE PUMPS.
LUBRICATION SYSTEM COMPONENTS
The lubrication system consists of four subsystems:
•
•
•
•
the
the
the
the
lubrication supply system
lubrication scavenge system
oil seal pressurization system
sump venting system.
SYSTEM MONITORING AND LIMITATION
The operation of the engine oil system may be monitored by the
following flight deck indications
• engine oil pressure
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GENERAL ... NO.4 BEARING SCAVENGE VALVE
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OIL INDICATING SYSTEM
GENERAL
The oil system monitoring is performed by:
•
indications:
ƒ oil quantity (quarts)
ƒ oil temperature (degree celsius)
ƒ oil pressure (psi)
• audio and visual warnings:
ƒ oil low pressure (LO PRESS)
ƒ oil filter clogging (OIL FILTER CLOG)
ECAM OIL INDICATIONS
1. Oil quantity indication flashes green (Advisory):
ƒ when QTY<4quarts.
2. Oil pressure indication color turns red (Warning)
ƒ when press <60 PSI.
3. Oiltemperature indication flashes green (Advisory)
ƒ when TEMP >156 deq.C
ƒ turns amber when oil TEMP < 10 deg C or > 165 deg C.
4. Oil fitter clog (White & amber) warning appears on the screen when
the engine scavenge filter is clogged.
Oil HI TEMP is displayed:
ƒ
when oil TEMP >165 deg C or 156 deg C more than 15 min.
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OIL PARAMETERS INDICATION
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VENT CIRCUIT
Air drawn in with the scavenge oil is separated in the tank by a deaerator and is vented to a de-oiler. The No.4 bearing scavenge line is
connected to the de-oiler through the scavenge valve. The de-oiler
separates the oil still in the air and discharges this air overboard. Oil
and air from the No.4 bearing compartment is scavenged through a
common line. A pressure transducer, installed on this line, monitors
the No.4 bearing scavenge pressure. The No.4 bearing scavenge valve
maintains the compartment differential pressure to ensure adequate
sealing.
As the engine power increases, the P10 pressure rises. When this
pressure exceeds 150 psi, the valve moves away from the max flow
stop. This is due to the pressure acting on the differential areas of the
valve and overcoming the spring load. The valve moves towards the
min flow or high power setting. As the valve moves towards the
peripheral ports in the seat, totally closing these ports, the flow
through the valve is now restricted to one central port in the valve
seat. Full travel is achieved at P10 pressure of approximately 210 psi.
As the valve moves away from the max flow stop, the influence of the
magnets on the reed switch decreases and the reed switch opens. The
circuit is broken, indicating that the valve has moved.
The No.4 bearing scavenge valve maintains No.4 bearing compartment
seal differential pressure by controlling the venting of the de-oiler
compartment air/oil mixture. It is a pneumatically operated twoposition valve which uses stage 10 compressor air pressure as servo
air. The valve is fully open at low engine speeds and closed to
minimum-flow position at high engine speeds.
As the engine power decreases, the spring load overcomes the
decreasing P10 pressure. The valve moves towards the max flow or
low power position, uncovering the ports in the valve seat and
restoring maximum flow through the valve. As the valve approaches
the maximum flow stop, the influence of the magnets on the reed
switch increases.
NO.4 BEAR. SCAV. VALVE DESCRIPTION
The reed switch closes, completing the circuit and indicating the valve
position.
OPERATION
There are two basic operating positions, low power and high power. In
where the compressor 10th stage pressure (P10) is less, than 150 PSI,
the valve is held spring loaded in the fully open position the low-power
position,. The bearing compartment scavenge flow passes through the
valve, restricted only by the porting in the valve seat.
The valve operates in the following manner. There are two basic
operating positions, low power and high power. In the low-power
position, where the compressor 10th stage pressure (P10) is less than
150 PSI, the valve is held spring loaded in the fully open position. The
bearing compartment scavenge flow passes through the valve,
restricted only by the porting in the valve seat.
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NO.4 BEARING SCAVENGE VALVE
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NO.4 BEARING SCAVENGE VALVE INDICATING
The EIU incorporates three logics allowing the monitoring of the
scavenge valve operation as well as a No.4 bearing carbon - seal
failure
differential pressure of the oil pressure feed line, and the scavenge
line). The No. 4 compartment scavenge oil pressure range is 0 to 160
PSI. Normal operating pressure is 0-145 PSI after three minutes of
stabilization at idle speed.
LOW POWER SETTING:
At engine low power, the bearing scavenge valve is open and the reed
switch on the valve closes providing a ground signal for the EIU logic.
HIGH POWER SETTING:
At engine high power, the bearing scavenge valve closes (to maintain
the No.4 bearing pressure ratio in the bearing compartment) and the
reed switch on the valve opens.
The No.4 bearing internal pressure is measured by the No.4 bearing
pressure XMTR in the oil return line to the de-oiler. The transducer
supplies a pressure signal to one of the three EIU logics.
Two EIU logics provide a warning message to the ECAM: ENG 1 (2)
BEARING 4 OIL SYS. (class 2) and a CFDS message, when the valve is
not in the correct position according to the sensed burner pressure.
One EIU logic provides a message on the lower ECAM: ENG.1 (2)
BEARING (class 2) and a fault message is set on the CFDS (EIU menu)
when the No. 4 bearing compartment pressure is to high according to
the valve position and a high burner press. (possible Carbon seal
failure)
ENGINE OIL PRESSURE
The Oil pressure is directly linked to the opening and closing of the
No.4 Bearing, Scavenge Valve.
A closing of the valve (at approx. 85% N2) will restrict the return
scavenge flow to the de-oiler. This will result in a pressure drop,
because the ratio of the pressures will change. (the oil pressure is the
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OIL PRESSURE TABLE
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EEC ELECTRICAL PWR SPLY CONTROL (3)
AUTO DEPOWERING
GENERAL
The Electronic Engine Control (EEC) is electrically supplied by:
•
•
the A/C network when High Pressure Rotor Speed (N2) is
below 10% or when the dedicated generator has failed,
the dedicated generator when N2 is above 10%.
POWERING N2 < 10%
The EEC is supplied by the A/C electrical power network when N2 is
below 10%. Each channel is independently supplied by the A/C 28V DC
through the Engine Interface Unit (EIU).
•
•
The FADEC is automatically depowered on the ground, through the
EIU, after engine shutdown. EEC automatic depowering occurs on the
ground:
•
•
5 min after A/C power-up,
5 min after engine shutdown.
NOTE: AN ACTION ON THE ENGINE FIRE P/B PROVIDES EEC
POWER CUT-OFF FROM THE A/C NETWORK.
MANUAL REPOWERING
The aircraft 28V DC permits:
•
28V DC supply and the dedicated generator power supplies is done
automatically by the EEC.
automatic ground check of the Full Authority Digital Engine
Control (FADEC) system before the engine is running, that is to
say FADEC GrouND PoWeR ON,
engine starting: MASTER lever ON or mode selector on IGNition
or CRANK,
powering the EEC while the engine reaches 10% of N2.
NOTE: THE EIU TAKES ITS POWER FROM THE SAME BUS BAR
AS THE EEC.
For maintenance purposes and MCDU engine tests, the ENGine FADEC
GrouND PoWeR panel permits FADEC power supply to be restored on
the ground with engines shut down. When the corresponding ENGine
FADEC GrouND PoWeR P/B is pressed ON the EEC recovers its power
supply.
NOTE: THE FADEC IS ALSO REPOWERED AS SOON AS THE
ENGINE START SELECTOR IS IN IGNITION/START OR CRANK
POSITION, OR THE MASTER LEVER IS SELECTED ON.
POWERING N2 > 10%
As soon as the engine is running above 10% of N2, the dedicated
generator directly supplies the EEC. The dedicated generator supplies
each channel with three-phase AC. Two transformer rectifiers provide
28V DC power supply to channels A and B. Switching between the A/C
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GENERAL ... MANUAL REPOWERING
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IGNITION & STARTING SYSTEM
PRESENTATION (3)
GENERAL
The ignition system provides the electrical spark needed to start or
continue engine combustion. The ignition system is made up of two
independent subsystems energized by a relay box. Each subsystem
includes an ignition exciter, a coaxial shield ignition lead and an igniter
plug. The pneumatic starting system drives the engine High Pressure
(HP) rotor at a speed high enough for a ground or in flight start to be
initiated. The start system is made up of the start valve and the
starter.
MANUAL START
During a manual start, the start valve opens when the engine MANual
START P/B is pressed in, then the ignition system is energized when
the MASTER control lever is set to the ON position.
NOTE: THERE IS NO AUTOMATIC SHUTDOWN FUNCTION IN
MANUAL MODE.
CRANKING
Engine motoring could be performed for dry cranking or wet cranking
sequences.
NOTE: DURING CRANKING IGNITION IS INHIBITED.
CONTROL AND INDICATING
The Electronic Engine Control (EEC) controls the ignition through the
relay box and starting through the start valve, either in automatic or
manual mode. The operation of the start valve and of the ignition
system is displayed on the ENGINE ECAM page.
CONTINUOUS IGNITION
AUTOMATIC START
SAFETY PRECAUTIONS
During an automatic start, the EEC opens the start valve, then the
ignition exciter is energized when the HP rotor speed is nominal. The
EEC provides full protection during the start sequence. When the
automatic start is completed, the EEC closes the start valve and cuts
off the ignition. In case of an incident during the automatic start the
EEC aborts the start procedure.
Safety precautions have to be taken prior to working in this area.
TDTI / HAT / ATA 71-80
Issue: 06/08
With engine running, continuous ignition can be selected via the EEC
either manually using the rotary selector or automatically by the Full
Authority Digital Engine Control (FADEC).
WARNING: THE EEC AND THE RELAY BOX SEND 115 VOLTS TO
THE IGNITION BOXES, WHICH CONVERT IT AND SEND HIGH
ENERGY PULSES THROUGH THE IGNITION LEADS TO THE
IGNITERS PLUGS.
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GENERAL ... MAINTENANCE PRACTICES
TDTI / HAT / ATA 71-80
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STARTING COMPONENTS
STARTER MOTOR
The pneumatic starter motor is mounted on the forward face of the
external gearbox and provides the drive to rotate the H.P. compressor
to a speed at which light up can occur.
A V-clamp adaptor does attachment to the gearbox.
Ducting to the aircraft pneumatic system connects the starter motor.
The starter motor gears and bearings are lubricated by an integral
lubrication system.
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 H.P. 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 r.p.m., centrifugal force overcomes the tension of the
clutch leaf springs, allowing the pawls to be pulled clear of the
MAINTENANCE PRACTICES
To increase A/C dispatch, the start valve is equipped with a manual
override. For this manual operation, the mechanic has to be aware of
the engine safety zones.
Servicing features include:
ƒ
ƒ
ƒ
oil level sight glass
oil fill plug
oil drain plug with magnetic chip detector
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, shown below, 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.R compressor rotor assembly.
TDTI / HAT / ATA 71-80
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STARTER SYSTEM COMPONENTS
TDTI / HAT / ATA 71-80
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IGNITION & STARTING SYSTEM D/O (3)
GENERAL
The Electronic Engine Control (EEC) controls and monitors the start
sequence either in automatic or in manual mode.
The start sequence is aborted below 50% N2 in case of:
• starter valve failure,
• ignition failure,
• pressure raising Shut-Off Valve (SOV) failure,
• hot start,
• hung start,
The system consists of a starter valve, a pneumatic starter, a relay
box, two ignition exciters and igniters A and B. The starter valve is
fitted with a manual override for mechanic operation on ground.
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GENERAL
TDTI / HAT / ATA 71-80
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AUTO START
In aircraft configuration, the APU is running and APU bleed air is
available. When the MODE selector is set to IGNition START, the EEC
is armed for the start sequence. After checking indications and
obtaining Ground Clearance, set the ENG/MASTER control switch to
ON.
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AUTO START
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MASTER SWITCH ON
As soon as the MASTER switch is set to ON, the Low Pressure (LP) fuel
SOV opens and the EEC opens the starter valve:
•
•
N2 increases,
the pressure raising and SOV solenoid is de-energized because
the MASTER switch is ON.
TDTI / HAT / ATA 71-80
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AUTO START - MASTER SWITCH ON
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DURING ENGINE START
When the MASTER switch is set to ON, the starter air valve will open,
the engine will dry crank for 30 seconds. Then the EEC provides
ignition, which is automatically selected by the EEC and displayed on
the ECAM ENGINE page.
TDTI / HAT / ATA 71-80
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AUTO START - DURING ENGINE START
TDTI / HAT / ATA 71-80
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AFTER 30 SECONDS
After 30 seconds and together with the ignition, the EEC opens the
Fuel Metering Valve (FMV) and the resulting fuel flow opens the
pressure raising and SOV. In case of malfunction, the EEC
automatically shuts down the engine and performs a dry motoring
sequence.
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AUTO START - AFTER 30 SECONDS
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N2 AT 43%
When N2 reaches 43%, the EEC closes the starter valve and cuts off
the ignition.
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AUTO START - N2 AT 43%
TDTI / HAT / ATA 71-80
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MODE SELECTOR SET TO NORM
Engine 2 is now stabilized at minimum idle.
NOTE: TO START THE SECOND ENGINE, YOU LEAVE THE MODE
SELECTOR IN THE IGN START POSITION, AND SET THE
MASTER LEVER 1 TO ON.
After engine start the MODE selector is set to NORMal with the engine
running.
WARNING: IF IGN START IS RESELECTED, THE CONTINUOUS
RELIGHT FUNCTION IS INITIATED ON THE RUNNING
ENGINE(S).
TDTI / HAT / ATA 71-80
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AUTO START - MODE SELECTOR SET TO NORM
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MASTER LEVER OFF
At any time, if the MASTER lever is set to OFF, the start sequence or
engine operation is stopped because the MASTER lever directly
energizes the pressure raising and SOV solenoid. With the MASTER
lever to OFF, the LP and pressure raising SOVs close.
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AUTO START - MASTER LEVER OFF
TDTI / HAT / ATA 71-80
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MANUAL START
In aircraft configuration, the APU is running and APU bleed air is
available. When the MODE selector is set to IGN START, the EEC is
armed for the start sequence. Action on the ENG MANual START P/B
opens the starter valve, via the EEC. After 30 seconds dry crank, set
the ENG/MASTER control switch to ON.
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MANUAL START
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AFTER 30 SECONDS START MAN P/B PRESSED
As soon as the MASTER lever is in the ON position, both ignition
systems are energized, LP and pressure raising SOVs open and the
fuel flow increases. A dual ignition and a fuel flow start. The MAN
START P/B stays latched and is normally left alone during the start
sequence.
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MANUAL START - N2 AT 18%
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MAN START P/B RELEASED
Action on the MAN START P/B has no effect on the start sequence as
long as the MASTER lever is set in the ON position. The MAN START
P/B is normally released at the end when the engine parameters are
stabilized. In case of malfunction, set the MASTER lever in the OFF
position to abort the start. Start for up to 2 minutes maximum.
NOTE: THERE IS NO AUTOMATIC SHUTDOWN FUNCTION IN
MANUAL START.
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MANUAL START - MAN START P/B RELEASED
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N2 AT 43%
When N2 reaches 43%, the EEC closes the starter valve and cuts off
the ignition.
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MANUAL START - N2 AT 43%
TDTI / HAT / ATA 71-80
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MODE SELECTOR SET TO NORM
Engine 2 is now stabilized at minimum idle.
NOTE: TO START THE SECOND ENGINE, YOU LEAVE THE MODE
SELECTOR IN THE IGN START POSITION, AND SET THE MAN
START P/B TO ON, AND THEN WHEN N2 REACHES 18% SET
THE MASTER LEVER TO ON.
After engine start the MODE selector is set to NORM with the engine
running.
WARNING: IF IGN START IS RESELECTED, THE CONTINUOUS
RELIGHT FUNCTION IS INITIATED ON THE RUNNING
ENGINE(S).
TDTI / HAT / ATA 71-80
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MANUAL START - MODE SELECTOR SET TO NORM
TDTI / HAT / ATA 71-80
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IGNITION STARTING- OPERATION
DESCRIPTION
The ignition circuit is supplied with 115VAC - 400Hz. The electrical
power is supplied via the EEC and EIU, which controls the ignition of
the igniter plugs. A dormant failure of an ignition exciter is not possible
for more than one flight because:
ƒ
ƒ
the two ignition systems are independent
the EEC selects alternately ignition system A or B.
CONTINUOUS IGNITION SELECTION
MANUAL SELECTION
When the engines are running on the ground or in flight the
continuous ignition is obtained by positioning the ENG/MODE selector
switch in IGN/START position.
FAIL SAFE POSITION: “IGN RELAYS, IGN ON”
IGNITION DURING AUTOMATIC START SEQUENCE
When an automatic start sequence has been activated by the EEC
(ENG/ MODE selector switch in IGN/START position and MASTER
control switch to ON), the EEC energizes automatically the appropriate
ignition exciter when N2 reaches between 10%-16% depending on
TAT and keeps it energized until N2 reaches 43%. For inflight restart
the EEC selects simultaneously both ignition exciters, on the ground,
after engine start, the selector must be placed in NORM position, then
back to IGN/START to select continuous ignition (both igniters). In
flight after engine restart, if the selector is maintained in IGN/START
position, the EEC selects the continuous ignition on the corresponding
engine. In case of a fault during an automatic starting on the ground,
the EEC aborts automatically the sequence by closing the starter shutoff valve and the HP fuel shut-off valve and deenergizing the igniters.
IGNITION DURING ALTERNATE START SEQUENCE (MANUAL
START PROCEDURE)
When a manual start sequence has been activated by the EEC
(ENG/MODE selector switch in IGN/START position and the ENG/MAN
TDTI / HAT / ATA 71-80
START pushbutton switch selected to ON) the EEC energizes both
ignition exciters. The deenergization of the ignition exciters is
automatically commanded by the EEC when engine N2 speed teaches
43% (Starter cut-out). Positioning of the MASTER control switch to
”OFF”, during that starting sequence, results in ignition exciter
deenergization.
Issue: 06/08
AUTOMATIC SELECTION
The EEC selects automatically the continuous ignition in some specific
conditions:
ƒ
ƒ
ƒ
ƒ
ƒ
ƒ
engine running and air intake cowl anti-icing is selected to ON
EIU failed.
take-off or during flexible take off
approach idle selected.
In flight, when there is an engine flameout or stall
Reverse
IGNITER PLUG TEST
The operation of the igniter plugs can be checked on the ground,
engine not running, through the maintenance MENU mode of the
FADEC or manually (Manual Start without air)
IGNITION SYSTEM CIRCUIT BREAKERS
There are 5 ignition CB's installed in the cockpit. 49VU and 121 VU
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CONTINUOUS RELIGHT
TDTI / HAT / ATA 71-80
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IGN START SELECTION
IF IGNition START is reselected with an engine running, the
corresponding EEC supplies the 2 igniters together, to provide
continuous ignition.
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CONTINUOUS RELIGHT - IGN START SELECTION
TDTI / HAT / ATA 71-80
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NORM SELECTION
When NORM is restored, the continuous relight is cut off.
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CONTINUOUS RELIGHT - NORM SELECTION
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MASTER LEVER OFF
When the MASTER lever is set to OFF, the LP and pressure raising
SOVs close and the EEC functions are reset. Engine 2 is shut down.
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CONTINUOUS RELIGHT - MASTER LEVER OFF
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ENGINE CRANK
In aircraft configuration:
• the APU is running and APU bleed air is available,
• both engines are shut down.
When CRANK is selected on the ground, the ignition is inhibited. Action
on the ENG MAN START P/B provides opening of the starter valve via
the EEC. During the crank sequence, the starter limitations should be
observed. If the starter operation time is exceeded, a warning
message is displayed on the ECAM, but there is no automatic abort.
The starter limitations when performing a dry crank are:
•
•
a maximum of 3 consecutive cycles; 2 minutes on, 15 seconds
off up 2 times and one minute on, then 30 minutes off for
cooling,
or 4 continuous minutes on, then 30 minutes off for cooling.
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ENGINE CRANK
TDTI / HAT / ATA 71-80
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MASTER LEVER ON
To perform a wet crank, allow N2 to increase to 20% RPM before
setting the MASTER lever to ON. When the MASTER lever is set in the
ON position, the LP and pressure raising SOVs open.
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ENGINE CRANK - MASTER LEVER ON
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DRY CRANK
After wet cranking, the MASTER lever should be set to OFF, and dry
cranking will continue to eliminate fuel vapors from the engine. The
pressure raising and SOV is closed. Continue to dry crank for at least
60 seconds.
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ENGINE CRANK - DRY CRANK
TDTI / HAT / ATA 71-80
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MODE SELECTOR SET TO NORM
When the MAN START P/B is released out, the starter valve closes and
the engine shuts down. To complete the crank sequence the MODE
selector is set in the NORM position.
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ENGINE CRANK - MODE SELECTOR SET TO NORM
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ENGINE INDICATION/ MONITORING D/O
(3)
SECONDARY ENGINE DISPLAY
The lower display shows the secondary engine parameters listed
below. The engine page is available for display by command, manually
or automatically during engine start or in case of system fault:
INDICATION GENERAL
ƒ
PRIMARY ENGINE DISPLAY
ƒ
The, primary engine parameters listed below are permanently
displayed or, the Engine and Warning display (E/ WD)
ƒ
ƒ Engine Pressure Ratio (EPR)
ƒ Exhaust Gas Temperature (EGT)
ƒ N1 (low rotor speed)
ƒ N2 (high rotor speed)
ƒ FF (fuel flow)
After 5 min of the power up test the indication is displayed in amber
and figures are crossed (XX). Normal indication can be achieved by
using the FADEC GRD power switches, one for each engine at the
maintenance panel or by the MODE selector switch on the Engine
panel at the pedestal in CRANK or IGN/ START position for both
engine.
If a failure occurs on any indication displayed, amber crosses replace
the indication, the analog indicator and the marks on the circle
disappear, the circle becomes amber.
ƒ
ƒ
ƒ
ƒ
ƒ
ƒ
Some
Total FUEL USED
For further info see ATA 73
OIL quantity
For further info see ATA 79
OIL pressure
For further info see ATA 79
OIL temperature
For further info see ATA 79
Starter valve positions, the starter duct pressure and during
eng start up, that operating Ignition system (ONLY ON ENGINE
START PAGE)
In case of high nacelle temperature a indication is provided
below the engine oil temp. indication.
Engine Vibration - of N1 and N2
As warnings by system problems only:
o OIL FILTER CLOG
o Fuel FILTER CLOG
No. 4 BRG SCAV VALVE with valve position
engine parameters also displayed on the CRUISE page
Only in case of certain system faults and flight phases a warning
message appears on the Engine Warning Display.
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ENGINE PARAMETERS
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POWER INDICATING
Thrust limit mode is displayed in digital form, it indicates the mode
which the EPR limit value will be computed.
EPR INDICATION
•
EPR - ENGINE PRESSURE RATIO
The Engine Pressure Ratio indicating system consists of one combined
P2/ T2 sensor and eight ports located in each of the three LPT exhaust
case struts, P4.9. The pressure from these sensors are routed to the
EEC pressure transducer. The EEC converts the signal to a digital
format and process the pressure to form actual ERP (P 4.9/ P 2 ) and
transmits the EPR value to the ECAM. Each of the two channels
performs this operation independently.
1) Actual EPR
Actual EPR is green.
2) Cyan EPR command arc (transient)
from current EPR pointer to EPR command value is only
displayed with A/ THR engaged.
3) EPR TLA (white circle)
Predicted EPR corresponding to the thrust lever position.
4) EPR max (thicker amber mark)
It is the limit value of EPR corresponding to the full forward
thrust lever position.
5) REV indication
Appears in amber when one reverser is unstowed or unlocked
or, inadvertently deployed. In flight, the indication first flashes
for 9 sec. and then remains steady. It changes to green when
the reverser is fully deployed.
6) Thrust limit mode, EPR rating limit
TO GA, FLX, MCT, CL, MREV selected mode is displayed in
green, the associated EPR rating is displayed in blue. In MREV
no EPR value is displayed.
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In flight (or on ground with ENG stopped):
o The selected mode corresponds to the detent of the
most advanced thrust lever position
o Rating limit is computed by the EEC receiving the
highest actual EPR value (except on ground with ENG
stopped where it is computed by the EEC receiving the
most advanced thrust lever position).
NOTES: 1 - WHEN A THRUST LEVER IS SET BETWEEN TWO
POSITIONS THE EEC SELECTS THE RATING LIMIT
CORRESPONDING
TO
TH E
HIGHEST
MODE.
2 - WHEN IDLE IS SELECTED THE EEC SELECTS CL
3 - WHEN MREV IS SELECTED, THE EPR RATING
LIMIT VALUE IS RE PLACED BY AMBER CROSSES (
MREV MODE IS LIMITED BY N1)
•
On ground (with engines running)
o With engines running, on ground, whatever the lever
position is, this limit corresponds to: TO GA thrust limit.
o With engine running, on ground, if FLX mode is
selected, FLX EPR is displayed whatever the thrust lever
position between IDLE and FLX/ MCT.
If FLX mode is selected, the flexible take off temperature in °C,
selected through the FMS MCDU's, is displayed. For FLX mode
indication the ADIRU’s must be switched on.
The temperature value is displayed in green and the °C is displayed in
blue.
If a failure occurs on any indication displayed, amber crosses replace
the analog indication, the analog indicator and the marks on the circle
disappear, and the circle becomes amber.
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EPR INDICATION SYMBOLS
EPR INDICATION
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EPR SYSTEM COMPONENTS
P2/ T2 SENSOR
The P2/ T2 sensor is located near the 12 o'clock position of the inlet
cowl. It measures total pressure and temperature in the inlet air
stream of the engine forward of the engine front flange. The dual
output total temperature measurement is accomplished by two
resistance-sensing elements housed in the P2/ T2 sensor body. Each
channel of the Electronic Engine Control (EEC) monitors one of these
resistance elements and converts the resistance measurement to a
temperature equivalent. The total air pressure is carried via pressure
tubing to the pressure sensor located in channel A of the EEC. The P2/
T2 sensor has an anti-icing function accomplished by a single heating
element internally bonded to the sensor. The heater is a thermetically
seated, coaxial resistance element brazed internally to the sensor
casting. Aircraft power, which is used for the heater, is switched on
and off by the EEC depending on TAT (< 7,2°C heater “ON”) via the
relay box.
manifold is connected to the Electronic Engine Control (EEC channel
A). A pressure transducer located within the EEC converts the average
pressure at station 4.9 into a useable electronic signal (proportional to
pressure) that can be processed and used by the EEC to control the
engine.
NOTE: IN CASE OF LOSS OF P2/ T2 HEATING, AN AUTOMATIC
REVERSION FROM EPR MODE TO UNRATED N1 MODE
OCCURS.
P4.9 SENSORS
The P4.9 sensor and manifold has three probes, which measure the
total pressure of the exhaust gas stream.
Struts 4, 7 and 10 contain the pressure sensing ports. Each sensing
point contains eight radial pressure sensing ports which are combined
to yield an average pressure. The resulting average radial pressure
value from each strut is then plumbed into a manifold which provides
an overall turbine exhaust pressure average (P4.9). A tube from this
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EPR INDICATION COMPONENTS
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P2/ T2 HEATER
Aircraft Power which is used for the heater, is switched on and off by
the EEC, via the relay box. The heater and the heating Circuit can be
tested using the FADEC WDS Test menu.
NOTE: THE RELAY BOX ALSO CONTAINS THE 115V IGNITION
RELAYS.
FAIL SAFE POSITION: PROBE HEATER ON
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P2/T2 HEATING – RELAY BOX
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TEMPERATURE
EGT INDICATION
EGT INDICATOR
ACTUAL EGT
Normally displayed in green. Pulses amber up to MCT when EGT ≥
610°C. Pulses red when EGT ≥ 650°C.
NOTE: EGT INDEX PULSING AMBER MUST BE DISREGARDED
WHEN USING TO OR FLX THRUST.
MAX EGT
Thicker amber mark is set at ≥ 610°C, it is the max EGT value up to
MCT thrust. It is not displayed during:
ƒ
ƒ
Engine start up, instead a amber mark is placed at 635°C
Take Off sequence.
MAX PERMISSIBLE EGT
Goes up to 650°C. A red band begins at the point of over temperature
and a red cross line appears at the max value achieved.
RED CROSS LINE
is set at the max EGT over temperature achieved during the last leg.
The red cross line will disappear through corresponding DMC's - MCDU
action or by the next T/ O.
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EGT INDICATION SYMBOLS
EGT INDICATION
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The EGT probes are located at engine station 4.95 (LPT exhaust case
strut), at 9.5, 7.5, 4.5 and 2 o'Clock.
EGT PROBES
The measurement channel for the exhaust gas temperature consist of:
Four probe assemblies, each comprizing 2 thermocouples.
four thermocouples (one from each probe assembly) are used
to form an averaged signal send to the channel "A" of the EEC.
ƒ the remaining four thermocouples (one from each probe
assembly) are used to form an averaged signal, send to
channel "B" of the EEC.
The EEC uses the Exhaust Gas Temperature in the engine start control
logic and also transmits the EGT signal to the ECAM.
ƒ
The thermocouples are connected, in parallel, to the junction box for
each channel, from where two independent signals are send to the
EEC. Each signal is an average of the four probes.
EGT PROBES LOCATION
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EGT INDICATION COMPONENTS
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•
•
POWER
N1, AND N2 INDICATION
N1 INDICATION
•
N1 MODE switches
ON: - Thrust control reverts from EPR mode to N1 rated mode.
Following an automatic reversion to N1 rated or unrated mode,
pressing the P/B switch to confirm the mode.
ON, it illuminates blue
OFF: - If available, EPR mode is selected
The low pressure rotor speed signal is used in the EEC for engine
control computation and for ECAM visual display.
•
Actual N1
Displayed normally in green.
Pulses red if N1 exceeds 100%.
Pulses amber when N1 exceeds the N1 rating limit, in N1 MODE.
•
Max permissible N1
is 100 %. At 100 % a red band begins.
If the RPM exceeds 100 % index and numeric value pulses red.
•
Red cross line
is set at the max N1 over speed value achieved during the last
leg.
•
White circle
N1 command corresponding to the thrust lever (angle) position
(predict N1) appears when in rated N1 mode. N1 rated MODE
can activated automatically or by switching the N1 MODE switch
at the overhead panel (close to the ENG MAN START switches).
Both engine just-be in the same MODE, rated or unrated. Not
displayed in unrated N1 MODE. Auto thrust is not active in rated
N1 mode.
General: A failure title will be displayed on E/ WD in the MEMO display.
•
CHECK
appears for EPR, EGT, N1, N2 and FF, if the displayed value
compared by the DMC' s with the actual value from the EEC
differs and the last digit from the value shown will be XX ed.
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N1 CONTROLS AND INDICATION
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N2 INDICATION
The signal fore the HP rotor speed is originated from the dedicated
alternator to the EEC for use in engine control computation and to the
ECAM for visual display on ECAM. A separate signal goes to the engine
vibration monitoring unit (EVMU) for use in processing engine vibration
data.
•
Actual N2
Digital indication normally green.
It is overbrightness and grey boxed during engine start sequence
up to 43% (starter cut out).
Turns red if N2 exceeds 100 % and a red “X” appears. The red
“X” will disappear through corresponding DMC's - MCDU action
or by the next
T/ O.
General: A failure title will be displayed on E/ WD on the MEMO
display.
If a failure occurs on any indication displayed amber crosses replace
the analog indication, the analog indicator and the marks on the circle
disappear, the circle becomes amber.
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N2 INDICATION
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•
N1 SENSOR
The fan speed (N1) indication system has four sensors:
•
•
•
•
•
•
Two of them are used to provide EEC channels “A” and “B” with
N1 rotational speed signal.
One sensor acts as a spare fore either EEC channel (it can be
activated by changeover connectors at the junction box).
This sensor cannot be used in place of the N1 sensor dedicated
to the Engine Vibration Monitoring Unit with N1 analog signals
(trim balance sensor), see below.
One sensor provides the Engine Vibration Monitoring Unit with
N1 analog signals (trim balance sensor).
The N1 electrical harness tube goes through the inner strut of
the No. 3 strut of the intermediate structure and to the terminal
block.
The electrical leads from each sensor goes through the N1 tube
and is connected to the terminal block.
For the fan speed sensors, one turn on the LP shaft causes 60
teeth on the phonic wheel to pass its sensor.
For the trimbalance sensor, one slot in the phonic wheel passes
the sensor one time for one turn.
The EEC speed sensors have two pole pieces compared to the
trimbalance sensor that has only one pole piece.
If, the fan speed sensor No. 3 is unserviceable, disconnect the
harness leads No. 5 and No. 6 from their terminals No. 5 and No.
6 and reconnect the harness leads to the spare speed sensor as
described above.
INTERCHANGE OF N1 SPEED SENSORS
Task 77-11-00-860-010
•
If the fan speed sensor No. 1 is unserviceable, disconnect the
harness leads No. 1 and No. 2 from their terminals No. 1 and No.
2.
Reconnect the harness lead No. 1 to the terminal No. 3 and the
harness lead No. 2 to the terminal No. 4 of the spare speed
sensor.
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N1 INDICATION COMPONENTS
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DEDICATED ALTERNATOR (PMA)
The alternator function are:
•
•
the primary power source for the Electronic Engine Control (EEC)
N2 signal source for the EEC and Engine Vibration Monitoring
Unit (EVMU) and the cockpit
DESCRIPTION
The unit is designed for maximum reliability by the elimination of
splines, bearings or similar parts which can deteriorate or fail. The
rotor is mounted directly on the gearbox output shaft and the stator is
bolted to the gearbox housing.
The alternator provides two identical and independent power outputs,
one for each channel of the EEC.
•
•
•
It comprises two stators (one power and one speed) and a rotor.
Is driven from the main accessory gearbox
Consists of a magnetic rotor running in a stator the stator has
four independing windings, two of which provide three phase
frequency AC electric power to respectively channel “A” and “B”.
The third winding provides a single phase AC analog signal
proportional to N2 for the Engine Vibration Monitoring System.
The forth winging provides a dedicated N2 signal to Channel “A”
of the EEC.
•
The N2 windings gives an analog signal through the cockpit for
ECAM indication.
A shaft seal seals the stator and rotor from the gearbox. If a shaft seal
failure occurs and the alternator fills with engine oil, the alternator will
continue to function normally. To maintain the temperature of the
dedicated alternator at an acceptable level the alternator incorporate
an integral cooling air manifold using fan air.
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N2 INDICATION COMPONENTS
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INTERFACES
ANALYZERS
DUAL ACCELEROMETER
A dual accelerometer is installed on each engine. It provides the
analog signals of N1 and N2 vibration frequencies. Only one sensor, A
or B, of the dual accelerometer is used at a time and during one flight.
It is automatically selected by the Engine Vibration Monitoring Unit
(EVMU) at each flight, the second one is in back-up mode in case of
failure. These accelerometers are also used for fan trim balance.
The EVMU interfaces with the ECAM, with the Centralized Fault Display
System (CFDS) and the Aircraft Integrated Data System (AIDS). CFDS
interface: maintenance fault messages and vibration data analysis.
AIDS interface: performance data reports.
NOTE: THERE IS NO INTERFACE WITH THE ELECTRONIC
ENGINE CONTROL (EEC).
NOTE: IN CASE OF SENSOR FAILURE, THE TRANSFER TO THE
SECOND SENSOR IS CARRIED OUT ON THE GROUND
THROUGH THE MCDU.
EVMU
An EVMU monitors the N1 and N2 vibration levels of both engines. The
EVMU determines for each engine, the N1 and N2 vibration levels by
analyzing:
• N1 and N2 speeds,
• dual accelerometer frequency signals.
The EVMU also computes the position and amplitude of the unbalance
and is capable of on-board fan trim balancing.
VIBRATION INDICATION
The N1 and N2 vibrations of the left and right engines are displayed on
the ENGINE and CRUISE pages. The maximum value that can be
displayed is of 10 units. 1 unit for N1 or N2 rotor corresponds to
0.3IPS (Inch Per Second).
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GENERAL ... INTERFACES
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ƒ
ƒ
ƒ
ENGINE VIBRATION MONITORING UNIT (EVMU)
DESCRIPTION
the CFDIU
the DMU
and printer (if installed) for maintenance purposes.
The signal conditioner is composed of:
POWER SUPPLY MODULE
ƒ 2 channel modules
ƒ 1 balancing module
ƒ 1 data processing module
ƒ 1 power supply module.
These modules are removable parts from the signal conditioner and
are repairable subassemblies.
The power supply module receives the 115 VAC/400Hz power. It
provides the other modules with the necessary voltages.
CHANNEL MODULES
Each channel module processes the signals from the two engine
accelerometers and from the two speed signals N1 and N2: this
enables the extraction from the overall vibration signal of a component
due to rotor first order unbalance.
The N1 and N2 signals are used to:
ƒ drive the tracking filters, and
ƒ slave their center frequencies at the shaft rotational speed.
The accelerometer signals pass through these tracking filters, which
extract the N1 and N2, related fundamental vibration. The acceleration
signal is then integrated in order to express the vibration in velocity
terms.
The EVMU receives analog signals from:
ƒ the 2 engine accelerometers (1 per engine)
ƒ and the N1 and N2 speed sensors of each engine.
It also receives digital input from CFDS through ARINC 429 data bus.
The EVMU sends signals through the digital ARINC 429 data bus to:
ƒ
SDAC1 and 2 for cockpit indication
TDTI / HAT / ATA 71-80
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POWER SUPPLY
The EVMU is supplied with 115W400Hz by the busbar 101 XPA,
through the circuit breaker 1EV.
BUILT IN TEST EQUIPMENT (BITE) MAINTENANCE AND FAULT
INFORMATION
The equipment contains a BITE system to detect internal and external
failure.
During the execution of the cyclic BITE sequence, the following parts
of the EVMU are checked:
ƒ the non-volatile memory
ƒ the timers
ƒ the analog-to-digital converter
ƒ the ARINC 429 transmitter and receivers
ƒ the tacho generators.
During the power-up sequence of the BITE, the following parts of the
EVMU system are checked:
ƒ N1 and N2 NB velocity
ƒ unbalance data
ƒ N1 and N2 tacho frequencies
ƒ accelerometer signals.
Any detected failure is stored in the non-volatile memory with GMT,
the date and other reference parameters.
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EVMU
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COMPONENTS
The vibration transducer including two indipendent channels is
installed on the fan case at the top left side of the engine.
The EVMU is located in the Avionics compartment 86VU.
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ANALYZER SYSTEM COMPONENTS
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FUEL FILTER CLOG
ENGINE WARNINGS (3)
In case of FUEL FILTER CLOG, the failure is shown amber on the EWD.
This warning appears when the pressure loss across the fuel filter is
excessive, i.e. the differential pressure is greater than 5 PSI.
OIL LO PR
In case of OIL LOw Pressure (OIL LO PR), the MASTER WARNing
flashes and the aural warning sounds, i.e. a Continuous Repetitive
Chime (CRC). The failure is shown red on the EWD. This warning
appears when the oil pressure is lower than 60 PSI.
OIL HI TEMP
In case of OIL HIgh TEMPerature (OIL HI TEMP), the MASTER
CAUTion comes on and the aural warning sounds, i.e. Single Chime
(SC). The failure is shown amber on the EWD. Firstly the oil
temperature indication flashes green when the temperature is higher
or equal to 155ºC and the warning appears amber when the engine oil
temperature is between 155°C and 165°C for more than 15 minutes,
or if the oil temperature is greater than 165°C.
OIL LO TEMP
In case of OIL LO TEMP, the MASTER CAUT comes on and the aural
warning sounds, i.e. an SC. The value of the corresponding parameter
is displayed amber on the ECAM page. This warning appears when the
oil temperature is below -10ºC.
OIL FILTER CLOG
In case of OIL FILTER CLOG, the failure is shown amber on the ECAM
displays. This warning appears when the pressure loss across the main
scavenge oil filter is excessive, i.e. the differential pressure is greater
than 12 PSI.
TDTI / HAT / ATA 71-80
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FUEL VALVE FAULT
In case of FUEL VALVE FAULT, the MASTER CAUT
corresponding engine panel FAULT light come on and
warning sounds i.e. an SC. This warning appears when the
(pressure regulating and shut-off valve) fails open or closed
of fuel valve switch position failure.
and the
the aural
fuel valve
or in case
EGT OVER LIMIT
In case of N1, N2 or EGT OVER LIMIT, the MASTER CAUT comes on
and the aural warning sounds, i.e. an SC. The failure message appears
amber on the EWD. The failure indications appear in the following
cases:
•
•
•
•
•
N1 100%: flashes red,
N2 100%: steady red,
EGT 610ºC: flashes amber,
EGT 635ºC: flashes red for A319/A320,
EGT 650ºC: flashes red for A321.
THR LEVER DISAGREE
In case of THRust LEVER DISAGREE, the MASTER CAUT comes on and
the aural warning sounds, i.e. an SC. The failure is shown amber on
the EWD. This warning appears when there is a disagreement between
both resolvers of a throttle lever.
In case of THR LEVER ANGLE FAULT, the MASTER CAUT comes on and
the aural warning sounds, i.e. an SC. The failure is shown amber on
the EWD. This warning appears when both resolvers on one throttle
lever are faulty.
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OIL LO PR ... FADEC FAULT
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EPR MODE FAULT
In case of EPR MODE FAULT, the MASTER CAUT comes on and the
aural warning sounds, i.e. an SC. The EPR indication is no longer
available. This warning appears when the Full Authority Digital Engine
Control (FADEC) is unable to sense or calculate EPR. The engine
automatically reverts to N1 MODE. Selection of the N1 MODE P/Bs on
the overhead panel will display:
•
•
the N1 MODE window,
the throttle lever reference, i.e. the white circle, and N1 max
limit, i.e. the amber index, on the N1 indicator.
FADEC FAULT
In case of FADEC FAULT, the MASTER CAUT comes on and the aural
warning sounds, i.e. an SC. The failure is shown amber on the EWD.
This warning appears when both Electronic Engine Control (EEC)
channels are faulty.
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OIL LO PR ... FADEC FAULT
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THRUST REVERSER SYSTEM
PRESENTATION (3)
REVERSER DESIGN
The thrust reverser system is of the aerodynamic blockage type. It
consists of two translating sleeves, blocker doors and cascade vanes to
redirect fan discharge airflow. The thrust reverser system is designed
for use on the ground only to reduce aircraft landing roll.
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REVERSER DESIGN
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(EPR) dial. The signals come from the lock sensor and the Linear
Variable Differential Transducer (LVDT).
HYDRAULIC SUPPLY
The thrust reverser system is hydraulically actuated utilizing the
aircraft hydraulic pressure from the corresponding engine. The thrust
reverser system is isolated from the hydraulic supply by a Shut-Off
Valve (SOV).
ACTUATION
Each translating sleeve is operated by two hydraulic actuators. The
actuators receive fluid from the Hydraulic Control Unit (HCU), which is
controlled by the Electronic Engine Control (EEC). When the deploy
sequence is commanded the pressure in the lower actuators releases
the locks as the four actuator pistons move rearward to deploy the
reverser. The actuators are linked together by a synchronizing system.
REVERSER CONTROL
Basically the thrust reverser system is controlled through the EEC from
the two reverser latching levers located on the throttle control levers.
The HCU has an isolation valve and a directional valve to select deploy
or stow mode. The directional valve is operated to deploy only. For
third defense line purposes, the Spoiler Elevator Computers (SECs)
have previously opened the SOV and the hydraulic pressure is supplied
to the HCU. Then, the Engine Interface Unit (EIU) permits reverser
deployment by energizing of the inhibition relay, so the directional
valve can be opened by the EEC. To command the thrust reverser, the
EEC needs an "aircraft on ground" signal supplied by the Landing Gear
Control and Interface Units (LGCIUs).
REVERSER INDICATING
The actual state of the thrust reverser is shown on the upper ECAM,
REV indication appears in the middle of the Engine Pressure Ratio
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HYDRAULIC SUPPLY ... REVERSER INDICATING
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MAINTENANCE PRACTICES
To help trouble shooting, a reverser test can be performed through the
MCDU. For maintenance purposes or to increase aircraft dispatch, the
HCU is fitted with a manual deactivation lever to inhibit the thrust
reverser system. In case of an inoperative thrust reverser, lockout pins
stowed on the translating sleeves have to be installed to complete the
deactivation procedure.
WARNING: THE THRUST REVERSER SYSTEM SHOULD BE
INHIBITED USING THE HCU DEACTIVATION LEVER BEFORE
WORKING ON THE SYSTEM OR ON THE ENGINE. THE SYSTEM
MUST BE DEACTIVATED IN ORDER TO PREVENT THE THRUST
REVERSER FROM OPERATING ACCIDENTALLY AND CAUSE
SERIOUS INJURIES TO PERSONNEL AND/OR DAMAGE TO THE
REVERSER.
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MAINTENANCE PRACTICES
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THRUST REVERSER MANAGEMENT (3)
GENERAL
The thrust reverser system is controlled independently for each engine
by the associated Full Authority Digital Engine Control (FADEC) system.
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GENERAL
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THRUST REVERSER ACTUATION
The hydraulic power required for the actuators is supplied by the
normal aircraft hydraulic system:
the EEC controls the thrust reverser operation through the
HCU. The lock sensors and the Linear Variable Differential
Transducers (LVDTs) are used to monitor the thrust reverser
position and for EEC control.
• green system for engine 1,
• yellow system for engine 2.
A Shut-Off Valve (SOV) located upstream of the Hydraulic Control Unit
(HCU) make up an independent locking system. Each channel of the
Electronic Engine Control (EEC) controls and monitors solenoid valves
included in the HCU, which provides the deployment and stowage of
two translating sleeves. Internal locks in the lower actuators are
hydraulically operated. The operation of the actuators is synchronized
by flexible drive shafts inside the deploy hydraulic lines, i.e. the
synchronizing system. The HCU includes an isolation valve, a pressure
switch and a directional valve, which is controlled through the
inhibition relay.
THRUST REVERSER CONTROL
When the reverse thrust is selected in the cockpit, the following
sequence occurs:
•
when the potentiometers detect a Throttle Lever Angle (TLA)
lower than -3°, the SOV opens if the altitude is less than 10 ft
and if high forward thrust, i.e. TLA < 30°, is not selected on
the opposite engine.
Then the HCU is supplied hydraulically. The SOV is controlled
energized open by the Spoiler Elevator Computers (SECs) through the
static and power relays. When the aircraft on ground signals from
LGCIU and TLA reverse signal from TCU are sent to EIU, the EIU
energize the inhibition relay.
•
when the aircraft is on ground with the engine running, i.e. in
N2 condition, and the resolvers detect a TLA lower than -4.3°,
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THRUST REVERSER ACTUATION & THRUST REVERSER CONTROL
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THRUST REVERSER ACTUATION & THRUST REVERSER CONTROL
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THIS PAGE INTENTIONALLY LEFT BLANK
TDTI / HAT / ATA 71-80
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THRUST REVERSER INDICATION
The thrust reverser operating sequences are displayed in the cockpit
on the EWD. An amber REVerse indication appears when the
translating sleeves are in transit (10% deployed) and then becomes
green (95% deployed)when the sleeves are deployed.
If this occurs in flight, REV will flash first for 9 sec then it will remain
steady. This indication will change to green color when the translating
sleeves are fully deployed and the reverse thrust can be applied.
In stowage, the indication changes to amber when one sleeve at least
is less than 95 % deployed and disappears when both sleeves are
stowed.
CFDS INTERFACE
The Centralized Fault Display System (CFDS) interfaces with the EIU to
provide thrust reverser fault diagnostics. For maintenance purposes, a
thrust reverser test can be performed through the MCDU menus. In
this case the Centralized Fault Display Unit (CFDIU) simulates an
engine running, i.e. the N2 condition, to permit the thrust reverser
deployment.
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THRUST REVERSER INDICATION & CFDS INTERFACE
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INITIAL CONDITIONS
Hydraulic pressure is available upstream of the Shut-Off Valve (SOV)
to isolate the Hydraulic Control Unit (HCU). Inside the HCU, the
isolation valve is in the closed position. The control solenoids are deenergized. The directional control valve is in the stow position. The
control solenoids are de-energized. The thrust reverser is maintained
in the FWD thrust position by mechanical locks, which are an integral
part of the lower actuators. The actuators are not yet pressurized.
Hydraulic supply:
•
•
green system for engine 1,
yellow system for engine 2.
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INITIAL CONDITIONS
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DEPLOY SEQUENCE
The selection of reverse thrust will provide signals to open the SOV
independently following the third defense line logic and through the
Electronic Engine Control (EEC), to energize the isolation valve so that
it moves to the open position. The EEC then energizes the directional
valve through the inhibition relay so that it moves to the deploy
position. The pressure switch provides signals to the EEC to indicate
that the hydraulic pressure downstream of the isolation valve is
sufficient. Hydraulic pressure is then applied on both sides of the
actuators but, due to differential piston areas, the actuators will extend
to move the translating sleeves to the deploy position. Tine locks
within the lower actuators are hydraulically released before translating
sleeve movement occurs.
NOTE: THE SIGNAL FROM THE EEC TO THE DIRECTIONAL
CONTROL VALVE IS ROUTED VIA AN INHIBITION RELAY
WHICH IS CLOSED BY THE ENGINE INTERFACE UNIT (EIU).
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DEPLOY SEQUENCE
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STOW SEQUENCE
Selection of FWD thrust will de-energize the directional control valve,
through the EEC, to allow the valve to move to the stow position. The
isolation valve remains energized providing hydraulic pressure to the
stow side of the actuators. The extend side of the actuators is opened
via the directional control valve to the hydraulic return.
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STOW SEQUENCE
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The EEC will de-energize the isolation valve 5 seconds after the
translating sleeves reach the fully stowed position to ensure full lock
engagement. Then the SOV is independently closed following the third
defense line logic.
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STOW SEQUENCE
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COMMAND LIMITATION
If the Linear Variable Differential Transducers (LVDTs) sense an
uncommanded movement of the thrust reverser:
From the stowed position, the EEC commands an automatic stowage
From the full deployed position, the EEC commands an automatic
deployment.
AUTO-RESTOW
In FWD thrust, if the EEC detects any uncommanded movement
greater than 10% from stow, it commands an auto-re stow of the
thrust reverser. Following auto-re stow, the isolation valve in the HCU
remains energized for the rest of the flight. In FWD thrust, if the EEC
detects any uncommanded movement greater than 15% from stow, it
commands engine idle power.
AUTO-REDEPLOY
In reverse thrust, if the EEC detects any un commanded movement
greater than 10% from full deploy, it commands an auto-re deploy of
the thrust reverser. When auto-re deploy is initiated to counteract
inadvertent stow, the EEC will command the isolation valve to close
and maintain it closed until FWD thrust has been reselected. The air
aerodynamic load on the translating sleeves will normally be sufficient
to redeploy the thrust reverser. In reverse thrust, if the EEC detects
any un commanded movement greater than 22% from full deploy, it
commands engine idle power.
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COMMAND LIMITATION - AUTO-RESTOW & AUTO-REDEPLOY
TDTI / HAT / ATA 71-80
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SERVICING, MAINTENANCE AND MEL
ITEMS
•
CHECK OIL LEVEL ON THE SIGHT GAGE ON THE OIL
TANK,
•
RAISE FILLER CAP HANDLE TO VERTICAL (UNLOCKED
POSITION),
•
TURN THE OIL FILLER CAP TO REMOVE,
•
ADD OIL AS NECESSARY UP TO THE FULL MARK ON THE
SIGHT GAGE,
•
INSTALL OIL FILLER CAP - MAKE SURE TO LOCK THE
CAP.
ENGINE OIL SERVICING
CAUTION: THE ENGINE SHOULD BE SHUT DOWN FOR AT
LEAST 5 MINUTES PRIOR TO OIL SERVICING. THIS ALLOWS
THE RESIDUAL PRESSURE IN THE OIL TANK TO DECREASE. IF
YOU OPEN THE FILLER CAP WHEN THERE IS PRESSURE IN
THE TANK THE HOT OIL CAN SPRAY OUT AND BURN YOU.
NOTE: IF POSSIBLE, THE ENGINE OIL SHOULD BE CHECKED
AND SERVICED WITHIN 5 TO 60 MINUTES AFTER
SHUTDOWN.
NOTE: IF THE ENGINE HAS BEEN SHUTDOWN MORE 1 HOUR
BUT LESS THAN 10 HOURS, START THE ENGINE AND RUN AT
IDLE FOR 3 MINUTES PRIOR TO SERVICING.
NOTE: IT IS ALSO POSSIBLE TO PRESSURE FILL THE ENGINE
OIL. TWO PORTS ARE INSTALLED ON THE OIL TANK, ONE FOR
PRESSURE AND ONE FOR OVERFLOW. SEE AMM FOR
PROCEDURE.
NOTE: IF THE ENGINE HAS BEEN SHUT DOWN FOR 10 HOURS
OR MORE, YOU MUST DRY CRANK THE ENGINE FOLLOWED BY
AN ENGINE START AND IDLE RUN OF AT LEAST 3 MINUTES
DURATION. THIS IS TO ENSURE THAT THE OIL LEVEL SHOWN
IN THE TANK IS CORRECT BEFORE OIL IS ADDED.
•
OPEN ENGINE OIL SERVICE DOOR ON LEFT FAN COWL,
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ENGINE OIL SERVICING
TDTI / HAT / ATA 71-80
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MASTER CHIP DETECTOR CHECK
The Master Magnetic Chip Detector (MCD) is located on the oil
scavenge filter housing attached to the oil tank. The probe will collect
any magnetic particles in the oil system. To check for contamination,
remove the Master MCD first:
•
•
•
open the left fan cowl,
push in and turn the MCD plug counterclockwise,
check the AMM for examples of NORMAL and ABNORMAL
contamination
NOTE: NO CONTAMINATION ON MASTER MCD
•
NO MAINTENANCE REQUIRED.
NOTE: CONTAMINATION ON MASTER MCD
•
INSPECT ALL OTHER MCD'S.
•
CLEAN THE MCD,
•
REPLACE SEAL RING AND RE-INSTALL
•
CHECK THAT THE RED MARKS ARE ALIGNED.
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MASTER CHIP DETECTOR CHECK
TDTI / HAT / ATA 71-80
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ADDITIONAL CHIP DETECTORS
Additional magnetic chip detectors are installed in the oil system to
isolate the source of metallic debris.
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MASTER CHIP DETECTOR CHECK - ADDITIONAL CHIP DETECTORS
TDTI / HAT / ATA 71-80
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MEL / DEACTIVATION
FUEL FILTER CLOGGING
In case of a failure of the FUEL CLOG warning on ECAM, the aircraft
may be dispatched per MEL as long as the fuel filter is changed once
each day. The filter housing is part of the fuel cooled oil cooler on the
fan case LH side.
Procedure:
•
•
•
•
•
•
•
FADEC GND PWR selected OFF,
open LH fan cowl,
drain residual fuel using drain plug,
open filter cover to remove and replace fuel filter element and
o-rings,
replace filter cover. Check AMM for correct torque value for
filter cover bolts,
perform minimum idle check for leaks,
close fan cowl.
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MEL / DEACTIVATION - FUEL FILTER CLOGGING
TDTI / HAT / ATA 71-80
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T/R DEACTIVATION AND LOCKOUT
Per the MEL, one or both Thrust Reversers may be deactivated in the
STOWED position for dispatch. The deactivation procedure has two
parts. First, the Hydraulic Control Unit (HCU) is deactivated. Moving
the deactivation lever to the inhibit position prevents the pressurizing
valve from supplying hydraulic pressure to the reverser actuators. In
the second part of the deactivation procedure each translating sleeve
is secured (bolted) to the reverser structure preventing any
movement.
TDTI / HAT / ATA 71-80
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MEL / DEACTIVATION - T/R DEACTIVATION AND LOCKOUT
TDTI / HAT / ATA 71-80
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OIL FILTER CLOGGING
In case of a failure of the OIL CLOG warning on ECAM, the aircraft
may be dispatched per MEL as long as the scavenge filter is changed
once each day. The filter housing is attached to the oil tank on the fan
case LH side.
Procedure:
•
•
•
•
•
•
•
•
FADEC GND PWR selected OFF,
open LH fan cowl,
drain residual oil using drain plug,
open filter cover to remove and replace the oil scavenge filter
element and o-rings,
replace filter cover. Check AMM/MEL for correct torque value
for filter cover bolts,
check Master MCD for contamination,
perform minimum idle check for leaks,
close fan cowl.
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MEL / DEACTIVATION - OIL FILTER CLOGGING
START VALVE MANUAL OPERATION
In case of an electrical failure of the start valve, the valve may be
operated manually to start the engine. The aircraft may be dispatched
per the MEL with the valve INOP closed.
•
MAKE SURE THAT THE START VALVE IS FULLY CLOSED.
NOTE: DO NOT OPERATE THE VALVE UNLESS THE STARTER
SYSTEM IS PRESSURIZED. DAMAGE TO THE VALVE CAN
OCCUR.
•
OPEN THE START VALVE ACCESS DOOR ON THE RH
COWL,
•
ESTABLISH COMMUNICATIONS WITH THE COCKPIT
(INTERPHONE JACK ON ENGINE INLET COWL),
•
ON COMMAND FROM THE COCKPIT, USE A 3/8"
SQUARE DRIVE TO MOVE THE START VALVE MANUAL
HANDLE TO THE OPEN POSITION.
NOTE: MAKE SURE YOU MAINTAIN PRESSURE AGAINST THE
SPRING TENSION TO KEEP THE VALVE OPEN.
•
AFTER ENGINE START, ON COMMAND FROM THE
COCKPIT, MOVE START VALVE MANUAL HANDLE TO
CLOSED.
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MEL / DEACTIVATION - START VALVE MANUAL OPERATION
TDTI / HAT / ATA 71-80
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MAINTENANCE TIPS
The engine and pylon drain system is designed to collect fuel, oil,
water and hydraulic fluid from engine systems and accessories and
discharge them overboard through the engine drain mast and the
pylon drain tubes. For troubleshooting and leak isolation the drain
mast body has separate drains identified and visible with the cowls
closed. The pylon drain tubes collect fluids from individual pylon
chambers, also for leak isolation. If fluid leaks are found during transit
operations, run the engine at idle for 5 minutes. If the leak stops, the
aircraft may be dispatched without maintenance action. If leaks
continue after 5 minutes, consult the AMM (ATA 71-70) for maximum
permitted leakage limits for all of the drains. There are 2 limits for
each drain. If the first limit is exceeded, the aircraft may be dispatched
and can continue to operate for a maximum of 25 hours or 10 flights
as long as the second limit is not exceeded. Here are some examples
of engine drains with both leakage limits. See the AMM for complete
list.
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MAINTENANCE TIPS
TDTI / HAT / ATA 71-80
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ENVIRONMENTAL PRECAUTIONS
Do not discharge products such as oil, fuel, solvent, lubricant either in
trash bins, soil or into the water network (drains, gutters, rain water,
waste water, etc...). Sort waste fluids and use specific waste disposal
containers. Each product must be stored in an appropriate and specific
cabinet or room such as a fire-resistant and sealed cupboard.
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ENVIRONMENTAL PRECAUTIONS
TDTI / HAT / ATA 71-80
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OPENING & CLOSING OF ENGINE COWL
DOORS (3)
FAN AND
OPENING
THRUST
REVERSER
COWL
DOORS
WARNING: DO NOT OPEN THE FAN COWL DOORS IF THE
WIND SPEED IS MORE THAN 60 MPH (96 KM/H). BE CAREFUL
IF YOU OPEN THE FAN COWL DOORS IF THE WIND SPEED IS
MORE THAN 30 MPH (48 KM/H), INJURY OR DAMAGE TO THE
ENGINE CAN OCCUR IF THE WIND MOVES THE FAN COWL
DOORS. MAKE SURE THAT THE HOLD OPEN RODS OF THE FAN
COWL DOORS ARE IN THE EXTENDED POSITION AND ARE
ATTACHED CORRECTLY. IF NOT, THE COWL DOORS CAN
CLOSE ACCIDENTALLY.
WARNING: MAKE SURE THAT THE TAKE-UP DEVICE IS
CORRECTLY ENGAGED BEFORE YOU RELEASE THE LATCHES.
FAILURE TO DO SO CAN CAUSE THE LATCHES TO OPEN FAST
AND CAUSE INJURY TO PERSONS.
FAN COWL DOORS OPENING
Before working on the engine, initial precautions have to be taken in
the cockpit. On the engine START panel, check that the MASTER lever
is at OFF position and install a warning notice stating not to start the
engine. On the maintenance panel, make sure that the ON legend of
the engine Full Authority Digital Engine Control (FADEC) GrouND
PoWeR switch is extinguished and install a warning notice. Then,
install a warning notice on the SLATS control lever, prohibiting use of
the slats. For maintenance operations, the fan and thrust reverser cowl
doors can be opened, but also, access panels are provided on the
nacelle. On the left fan cowl door, there are accesses to the oil tank
filler cap for servicing, and to the oil scavenge master chip detector for
inspection. On the right fan cowl door, an access is provided for the
start valve manual operation, to an override square drive. Let us now
see the opening of the fan cowl doors. First, unlock the four tension
latches on the engine centerline. For each latch, push the latch snap to
release the latch handle. Manually lift and support the door at the
lower edge. Release the front hold-open rod from its storage bracket
and attach it to its support on the fan case. Open the door sufficiently
to engage the rod, the rear hold-open rod is then extended and
attached to its support. Make sure that it is correctly engaged. The
second fan cowl door is opened in the same way.
CAUTION: DO NOT OPEN THE INBOARD THRUST REVERSER
"C" DUCT IF THE WING LEADING EDGE SLATS ARE EXTENDED.
DAMAGE TO THE THRUST REVERSER, WING LEADING EDGE
SLATS AND WING CAN OCCUR.
TDTI / HAT / ATA 71-80
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FAN COWL/ THRUST REVERSER DOORS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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FANCOWL DOORS LOCKS
TDTI / HAT / ATA 71-80
HOLD OPEN RODS
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Revision: 18.06.2008
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Opening of Fan Cowl Doors.wmv
FAN AND THRUST REVERSER COWL DOORS OPENING - FAN COWL DOORS OPENING
TDTI / HAT / ATA 71-80
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THRUST REVERSER COWL DOORS
T/R COWLING ("C-DUCT") OPENING/ CLOSING
CAUTION:
BEFORE OPENING:
1. WING SLATS MUST BE RETRACTED AND DEACTIVATED.
2. ALL 6 LATCHES & TAKE - UP DEVICES MUST BE RELEASED.
3. IF REVERSER IS DEPLOYED, PYLON FAIRING MUST BE
REMOVED.
4. DEACTIVATE THRUST REVERSER HYDRAULIC CONTROL
UNIT (HCU)
5. FADEC POWER "OFF"
6. PUT WARNING NOTICES IN THE COCKPIT
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C-DUCT DOORS OPENING WITH DEPLOYED THRUST REVERSER
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THRUST REVERSER HALF LATCHES
6 Latches are provided to keep the Thrust Reverser Halfs in the closed
position. They are located:
• 1 Front latch (access through the left fan cowl)
• 3, Bifurcation latches (access through a panel under the C-Duct
halves)
• 2 latches on the reverser translating sleeve (Double Latch)
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THRUST REVERSER DOORS LATCHES
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LATCH ACCESS PANEL & TAKE UP DEVICE
An access panel as shown below, is provided to gain access to the
three BIFURCATION "C" duct latches and the "C" duct take up device
(also called, Auxiliary Latch Assembly ).
The take up device is "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 both when closing and opening the "C"
ducts.
The take up device must be disengaged and returned to its stowage
bracket, inside the L/H "C" duct, when not in use.
NOTE: RED OPEN FLAGS INSTALLED ON THE C-DUCT
INDICATE THAT THE BIFURCATION LATCHES ARE OPEN.
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C-DUCT LATCHES/ TAKE UP DEVICE
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FRONT LATCH AND OPEN INDICATOR
Access to the front latch is gained through the left hand fan cowl. The
latch is equipped with a red open indicator. The open-indicator gets in
view through a gap in the cowling (also when the thrust reverser
halfs, are closed ) to indicate a not proper closed reverser cowl.
CAUTION:
MAKE SURE THAT YOU POSITION THE FRONT LATCH
CORRECTLY AGAINST THE FRONT LATCH OPEN INDICATOR
WHILE YOU PULL THE THRUST REVERSER HALVES TOGETHER
WITH THE AUXILIARY LATCH ASSEMBLY (TAKE UP DEVICE).
IF YOU DO NOT DO THIS THE FRONT LATCH CAN GET CAUGHT
BETWEEN THE THRUST REVERSER HALVES AND THE
AUXILIARY LATCH ASSEMBLY AND THE HOOK CAN GET
DAMAGED.
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FRONT LATCH INDICATOR
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C - DUCT OPENING/ CLOSING SYSTEM
On each “C” duct a single acting hydraulic actuator is provided for
opening. A hydraulic hand pump must be connected to a self sealing/
quick release hydraulic connection for opening.
NOTE: THE HYDRAULIC FLUID USED IN THE SYSTEM IS
ENGINE LUBRICATING OIL.
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C-DUCT OPENING ACTUATORS
TDTI / HAT / ATA 71-80
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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 L.H. “C”
duct is shown below, the R.H. V duct is similar.
WARNING
BOTH STRUTS MUST ALWAYS BE USED TO SUPPORT THE “C”
DUCTS IN THE OPEN POSITION. THE V DUCTS WEIGH
APPROX 578 LBS EACH. SERIOUS INJURY TO PERSONNEL
WORKING UNDER THE “C” DUCTS CAN OCCUR IF THE V DUCT
IS SUDDENLY RELEASED.
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C-DUCT HOLD OPEN RODS
TDTI / HAT / ATA 71-80
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THRUST REVERSER COWL DOORS OPENING
First, gain access to the hydraulic control unit lever, and deactivate the
thrust reverser system. Before unlocking the latches, open the latch
access panel. Engage the thrust reverser take-up device on its bracket,
and turn the adjustment nut to push the two halves together and to
relieve the tension on the latches. Once the tension is relieved, release
the five latches at all positions, aft, center and forward. Now the
adjustment nut of the take-up device is loosened to disengage the
take-up device which is then stored on its storage bracket. Before
opening the thrust reverser cowl doors, make certain that the slats are
retracted and the translating sleeves stowed. The thrust reverser cowl
doors are opened using a hydraulic hand pump. Remove the dust
cover from the quick disconnect and connect the hand pump. Open
the hand pump valve, and pump to pressurize the opening actuator
until the reverser half reaches the fully open position. When the door
is opened, unstow the hold open rods from their storage brackets and
attach them on their supports on the engine core. Unload the hand
pump until the rods hold the weight of that half. Disconnect the hand
pump manifold and put the cap on the quick disconnect. The second
half is opened in the same way. Once all doors are opened, the engine
is accessible for maintenance operations.
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Opening of Thrust Reverser Doors.wmv
FAN AND THRUST REVERSER COWL DOORS OPENING - THRUST REVERSER COWL DOORS OPENING
TDTI / HAT / ATA 71-80
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FAN AND
CLOSING
THRUST
REVERSER
Austrian Technical Training
School Notes - For Training Purposes Only
COWL
DOORS
CAUTION: MAKE SURE THAT THE ENGINE AREA IS CLEAR OF
TOOLS AND EQUIPMENT BEFORE CLOSING THE FAN COWL
DOORS.
THRUST REVERSER COWL DOORS CLOSING
Let us now see the cowls closing sequence. The thrust reverser doors
are closed first. Pressurize the opening actuator to release the load
from the hold-open rods. Disengage the rods from their supports and
store them. Slowly open the hand pump relief valve, the actual door
closing rate should be controlled by the hand pump. However, as a
safety device, the actuator ensures a minimum door closing time.
When the door is fully closed, disconnect the hand pump and replace
the cap on the quick disconnect. The left thrust reverser cowl door is
then closed. On the engine centerline, engage the take-up device, and
turn the adjustment nut to push the halves together. Then, engage
and lock all the latches. Do not forget to release and restore the takeup device before closing the latch access panel. On the Hydraulic
Control Unit (HCU), remove the inhibition pin to reactivate the thrust
reverser system.
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Closing of Thrust Reverser Doors.wmv
FAN AND THRUST REVERSER COWL DOORS CLOSING - THRUST REVERSER COWL DOORS CLOSING
TDTI / HAT / ATA 71-80
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FAN COWL DOORS CLOSING
Now close the fan cowl doors. Remove the aft and forward hold-open
rods from their support on the fan case and store them on the fan
cowl door. The second fan cowl door is closed in the same way. Push
the doors together to lock the latches. Check that the hooks are
correctly engaged and close the latches. Finally, in the cockpit, remove
the warning notices from the panels, and the SLATS control lever.
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Closing of Fan Doors.wmv
FAN AND THRUST REVERSER COWL DOORS CLOSING - FAN COWL DOORS CLOSING
TDTI / HAT / ATA 71-80
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THRUST REVERSER DEACTIVATION AND LOCKOUT
INDICATING THAT THE CORRESPONDING REVERSER IS
INOPERATIVE AND NOTE IT IN THE LOGBOOK.
WARNING: THE THRUST REVERSER SYSTEM SHOULD BE
DEACTIVATED USING THE HYDRAULIC CONTROL UNIT (HCU)
LEVER, BEFORE WORKING ON THE SYSTEM OR ON THE
ENGINE. IF NOT THE THRUST REVERSER CAN ACCIDENTALLY
OPERATE AND CAUSE SERIOUS INJURIES TO PERSONNEL
AND/OR DAMAGE TO THE REVERSER.
This procedure is carried out when a fault occurs on a thrust reverser,
which cannot be repaired for the next flight. Deactivation and lockout
are therefore provided to secure the translating sleeves in the stowed
position when the aircraft has to be dispatched with an inoperative
thrust reverser. First, on the ENGine START panel, check that the
MASTER lever is at OFF position and install a warning notice stating
not to start the engine. On the maintenance panel, make sure the
engine Full Authority Digital Engine Control (FADEC) GrouND PoWeR
switch is OFF, and install a warning notice. On the engine, open the
left fan cowl door. Gain access to the hydraulic control unit, move the
lever to deactivate the thrust reverser and install the safety pin to hold
the lever in the OFF position. Make certain that the translating sleeves
are in the retracted position. Lock out each translating sleeve by using
lock pins. For each translating sleeve, remove the thrust reverser lock
pin from the lower lock pin receptacle and the dummy pin from the
upper lock pin receptacle then, interchange them.
NOTE: THE RED HEAD OF THE LOCK PIN WHICH EXTENDS
ABOVE THE SURFACE OF THE SLEEVE TO INDICATE THE
LOCKOUT. CLOSE THE FAN COWL DOOR AND MAKE SURE
THAT THE WORKING AREA IS CLEAN AND CLEAR OF THE
TOOLS AND OTHER ITEMS. FINALLY, IN THE COCKPIT,
REMOVE THE WARNING NOTICES, AND INSTALL A PLACARD
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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School Notes - For Training Purposes Only
Thrust Reverser deactivation and Lockout.wmv
THRUST REVERSER DEACTIVATION AND LOCKOUT
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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PRECAUTIONS
CAUTION: DO NOT EXTEND AND STOW THE TRANSLATING
SLEEVES WITH THE THRUST REVERSER "C" DUCTS OPEN.
DAMAGE TO THE TRANSLATING SLEEVES, SYNCHRONIZING
SYSTEM AND THE HINGE ACCESS PANELS COULD OCCUR. DO
NOT USE POWER TOOLS IN THE MANUAL DRIVE SOCKETS.
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PRECAUTIONS
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MANUAL DEPLOYMENT AND STOWAGE PROCEDURE
WARNING: THE THRUST REVERSER SYSTEM SHOULD BE
DEACTIVATED USING THE HYDRAULIC CONTROL UNIT (HCU)
LEVER, BEFORE WORKING ON THE SYSTEM OR ON THE
ENGINE. IF NOT THE THRUST REVERSER CAN ACCIDENTLY
OPERATE AND CAUSE SERIOUS INJURIES TO PERSONNEL
AND/OR DAMAGE TO THE REVERSER.
NOTE: BOTH TRANSLATING SLEEVES MOVE TOGETHER DUE
TO THE SYNCHRONIZING SYSTEM.
Make sure that the A/C is in the same configuration as for the manual
deploying task.
Before the manual deployment of the translating sleeves some
precautions have to be taken in the cockpit. On the ENGine START
panel, check that the MASTER lever is at OFF position and install a
warning notice. On the maintenance panel, make sure that the ENGine
FADEC GrouND PoWeR SWitch is OFF and install a warning notice.
Gain access to the HCU, move the lever to the OFF position to
deactivate the thrust reverser system and install the safety pin. Gain
access to the manually by-passable non-return valve by removing the
corresponding pylon access panel. Now, move the non-return valve
lever to the bypass position. After that, on the lower actuator move
the lock lever to the unlocked position, this releases the actuator and
permits manual deployment. Before operating, check that the flexible
shaft is correctly engaged inside the manual drive, then insert a speed
wrench into the manual drive shaft and deploy the translating sleeves.
When looking FWD, you can see the blocker doors moving. With the
translating sleeves deployed the cascades are accessible. To stow the
translating sleeves, reverse the procedure. When the translating
sleeves reach the fully stowed position the lock lever on the lower
actuator automatically engages. Then return the manually by-passable
non-return valve lever to the normal position and reactivate the thrust
reverser system through the HCU lever. Finally, in the cockpit remove
the warning notices.
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Manual Deploying and Stowing of Thrust Reverser Translating Sleeves.wmv
MANUAL DEPLOYMENT AND STOWAGE PROCEDURE
TDTI / HAT / ATA 71-80
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PRECAUTIONS
WARNING: MAKE SURE THAT YOU HAVE THE CORRECT FIRE
FIGHTING EQUIPMENT AVAILABLE BEFORE YOU START ANY
TASK ON THE FUEL SYSTEM. MAKE SURE THAT THE L/G
SAFETY-LOCKS AND THE WHEEL CHOCKS ARE IN POSITION.
PUT THE SAFETY DEVICES AND THE WARNING NOTICES IN
POSITION BEFORE YOU START ANY TASK ON OR NEAR:
•
THE FLIGHT CONTROLS,
•
THE FLIGHT CONTROL SURFACES,
•
THE L/G AND THE ASSOCIATED DOORS,
•
ANY COMPONENT THAT MOVES.
MAKE SURE THAT ALL THE CIRCUITS IN MAINTENANCE ARE
ISOLATED BEFORE YOU SUPPLY ELECTRICAL POWER TO THE
A/C.
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PRECAUTIONS
TDTI / HAT / ATA 71-80
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Revision: 18.06.2008
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ENGINE REMOVAL AND INSTALLATION (3)
BOOTSTRAP SYSTEM INSTALLATION
The removal and installation of the engine requires the installation of a
bootstrap system on the aircraft pylon. The bootstrap system is
composed of two elements, to be installed at the front and at the rear
of the pylon. Each element permits to attach at its ends the chain
pulley blocks assembly and dynamometers that are used to lower or to
lift the transportation stand attached to the engine.
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BOOTSTRAP SYSTEM INSTALLATION
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BOOTSTRAP SYSTEM INSTALLATION
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THIS PAGE INTENTIONALLY LEFT BLANK
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ENGINE TRANSPORTATION STAND ATTACHMENT
POINTS
The engine transportation stand, which is used for engine removal and
installation, can be fixed to the engine by means of four trunnions:
•
•
two front trunnions fixed on the LP compressor case, LH side
and RH side,
two rear trunnions fixed on the LP turbine case, LH side and
RH side.
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ENGINE TRANSPORTATION STAND ATTACHMENT POINTS
TDTI / HAT / ATA 71-80
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ENGINE REMOVAL
Before the removal operation the engine systems must be deactivated
and isolated from the aircraft, and all the connections disconnected
from the pylon interfaces. To support the cowl doors during the
operation hold-open braces are installed. Then the engine trolley and
cradle are placed under the engine. To support the engine during
removal and transportation fixtures are installed at handling points on
the fan case and also on the turbine rear frame. On the FWD engine
mount, each bolt is loosened and retorqued to the maximum force an
operator using a standard wrench can apply. The center hinge clamp
of the FWD bootstrap is then fixed on the pyramid where hinge arms
will be installed. The rear beam of the bootstrap equipment is fixed on
the pylon lower part, then the dynamometers and chain pulley blocks
are installed. Disconnect the engine cradle from the trolley and raise
the cradle to the engine using the four chain pulley blocks. When the
cradle reaches the engine handling points secure the supports on the
front and on the rear parts. Loosen and remove the bolts of the FWD
mount, then the bolts of the aft mount through an access provided
when the hinge access panel is removed. Make sure that no lines or
unions remain connected to the pylon and lower the engine using the
four chain pulley blocks, also check that the engine mounts disengage
correctly from their attaches. Slowly continue the descent until the
cradle touches the trolley. Then secure and lock it. Remove the chain
hoists and the FWD hinge arms, then the chain pulley blocks and
dynamometers from the rear beam. Now push forward the engine
from under the pylon, making sure that nothing catches. With the
engine removed it is now possible to see the aft and FWD mounts in
more detail and also the front hinge clamp and the rear beam
installation. Notice the hold-open brace which holds the thrust reverser
cowl door in the open position.
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Engine Removal.wmv
ENGINE REMOVAL
TDTI / HAT / ATA 71-80
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ENGINE INSTALLATION
All the equipment is re-installed to provide the engine lifting. Start the
lifting using the pulley levers and check that the dynamometers show
the correct load. Raise the engine cradle gradually. As the engine
reaches the pylon take care that the FWD and rear mounts engage
correctly. Then install the bolts to both mounts and tighten them. Note
that you never use the same bolts. Always change them for new or
inspected ones. When the engine cradle is on the trolley remove the
four chain pulley blocks and dynamometers, then the bootstrap
equipment. Reconnect all lines and electrical harnesses to the pylon
interfaces. Finally lock the cradle to the trolley. Push the trolley out
from under the engine and remove the hold-open braces from the
cowls. The engine can now be put back in serviceable condition.
NOTE: AVOID ANY DYNAMIC EFFECT DURING DESCENT.
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Engine Installation.wmv
ENGINE INSTALLATION
TDTI / HAT / ATA 71-80
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FAN COWL DOOR REMOVAL/INSTALLATION
WARNING: DO NOT OPEN THE FAN COWL DOOR(S) WHEN
THE WIND SPEED IS 60 MPH (96 KPH) OR MORE. INJURY
AND/OR DAMAGE TO THE ENGINE CAN OCCUR IF THE WIND
MOVES THE FAN COWL DOOR(S).
REMOVAL
Make sure that the A/C is prepared for the removal task. First open the
fan cowl door and install the hoisting sling. Disengage and store the
hold-open rods. Lift the door with the hoist and gain access to the
hinges. Remove the four hinge bolts and nuts. Make a note of the
position of the bolts because they must be put at the same place
during the installation. Finally remove and lift the door over to the
transit stand.
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Fan Cowl Door Removal.wmv
FAN COWL DOOR REMOVAL/INSTALLATION - REMOVAL
TDTI / HAT / ATA 71-80
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School Notes - For Training Purposes Only
INSTALLATION
We will now see a fan cowl door installation. First of all, prepare the
A/C for maintenance tasks on the engine. On the transit stand attach
the hoisting sling to the lower part of the fan cowl door and secure it.
Then on the opposite side attach the sling to the upper hoisting points
and secure it. Remove the fan cowl door from the fan cowl transit
stand with the sling and the hoist and carefully install the door.
Examine the fan cowl hinges and the pylon hinges and make sure that
they are not damaged before aligning them. Install the bolts and nuts.
Make sure that the bolts are installed in the same positions noted
during removal. Deploy the hold-open rods and engage them on their
attach brackets on the fan case. Finally remove the sling and return to
the initial A/C configuration as defined in the close-up subtasks.
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Fan Cowl Door Installation.wmv
FAN COWL DOOR REMOVAL/INSTALLATION - INSTALLATION
TDTI / HAT / ATA 71-80
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THRUST
REVERSER
REMOVAL/INSTALLATION
Austrian Technical Training
School Notes - For Training Purposes Only
COWL
DOOR
WARNING: DO NOT GET HYDRAULIC FLUID ON YOUR SKIN,
IN YOUR MOUTH OR IN YOUR EYES. HYDRAULIC FLUID IS
POISONOUS AND CAN GO THROUGH YOUR SKIN AND INTO
YOUR BODY. FLUSH HYDRAULIC FLUID FROM YOUR MOUTH
OR YOUR EYES AND GET MEDICAL AID. BE CAREFUL DURING
THE REMOVAL OR INSTALLATION OF THE THRUST REVERSER
"C" DUCT, IT WEIGHS 580 LBS (263 KG).
the weight of the thrust reverser cowl door, and after having removed
the hinge access panel from the pylon, attach the last hoisting fixture
to the thrust reverser hinge upper beam. Next, disconnect the
hydraulic manifold from the thrust reverser cowl door opening
actuator. Install the protective caps, and remove the actuator. On the
thrust reverser system disconnect the electrical connector from the
electrical receptacle, and disconnect the deploy and retract hoses from
the upper actuator, then gain access to the pylon hinges and remove
the four hinge bolts and nuts. Carefully lift and disengage the thrust
reverser cowl door from the pylon hinges. Remove the cowl door
taking care that it does not hit against the slat surfaces. Finally lower it
onto its dolly.
CAUTION: DO NOT OPEN OR REMOVE THE THRUST REVERSER
"C" DUCTS IF THE WING LEADING EDGE SLATS ARE
EXTENDED. THIS WILL CAUSE DAMAGE TO THE THRUST
REVERSER, THE WING LEADING EDGE SLATS AND THE WING.
WARNING: MAKE SURE THAT THE HYDRAULIC HAND PUMP
HOSE IS CORRECTLY CONNECTED TO THE THRUST REVERSER
"C" DUCT OPENING ACTUATOR, IF NOT THE THRUST
REVERSER "C" DUCT CAN CLOSE QUICKLY AND CAUSE
INJURY TO PERSONS BETWEEN THE THRUST REVERSER "C"
DUCT AND THE ENGINE. MAKE SURE THAT ALL THE TRAPPED
AIR IS BLED FROM THE SYSTEM, IF NOT TRAPPED AIR WILL
CAUSE AN IRREGULAR CLOSE RATE OF THE THRUST
REVERSER "C" DUCT.
REMOVAL
First of all prepare the A/C for the removal task and deactivate the
thrust reverser system. Install the thrust reverser sling to the sling
points. First at the two lower hoisting points. Using the hoist hold up
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Thrust Reverser Door Removal.wmv
THRUST REVERSER COWL DOOR REMOVAL/INSTALLATION - REMOVAL
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INSTALLATION
Make sure that the A/C is in same configuration as for the removal
task with the thrust reverser system deactivated. On the dolly, attach
the thrust reverser hoisting sling to the three sling points on the thrust
reverser half. First, attach the sling to the hoisting fixture provided on
the thrust reverser hinge upper beam. Then attach the sling to the
lower hoisting points of the cowl door. Lift the cowl door from the
dolly, and carefully position the cowl door onto the pylon. Align the
thrust reverser cowl door hinges with the hinges on the crossover
beams of the pylon, and install the four hinge bolts, first at position
four, then at positions one, two and three. Next, remove the hoisting
fixture from the hinge upper beam and keep the lower two sling points
attached. Using the hoist, lift the door to the maximum open position.
Extend the rear and FWD hold-open rods and make sure that they are
correctly engaged. Remove the thrust reverser sling from the lower
hoisting points. After that, remove the dust cover from the quick
disconnect and connect the hydraulic hose from a hand pump or
service unit. Inside the thrust reverser half gain access to the
mounting bracket of the thrust reverser cowl door opening actuator.
Position the cowl door opening actuator, and attach it to its mounting
bracket on the thrust reverser cowl door. Then remove the protective
cap from the opening actuator hydraulic connector, and connect the
hydraulic manifolds. Using the hydraulic service units or hand pump
pressurize the actuator to bleed the air. When finished, tighten and
torque the hydraulic manifold connector. Then reinstall the protective
shroud. On the opposite side of the engine, gain access to the opening
actuator hinge beam on the pylon, and using the hydraulic service
units, extend the opening actuator rod to the hinge beam. Align and
attach it. Next close the thrust reverser cowl door by operating the
hydraulic service unit. On the engine bottom centerline, engage and
latch correctly both thrust reverser halves together. On the pylon area,
finish tightening the hinge bolts and torque them at all positions. On
the thrust reverser system, remove the dust cover of the hydraulic
TDTI / HAT / ATA 71-80
Issue: 06/08
deploy hose. Use a container to collect the drained hydraulic fluid, then
insert the reverser flexible shaft into the deploy hose, and connect the
deploy hose to the thrust reverser upper actuator. In the same way,
connect the thrust reverser retract hose. On the thrust reverser system
electrical receptacle, reconnect the electrical connector. On the pylon
area, reinstall the hinge access panel, and the pylon fairings. Finally,
return the A/C to the initial configuration as defined in the close-up
subtasks.
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Thrust Reverser Cowl Door Installation.wmv
THRUST REVERSER COWL DOOR REMOVAL/INSTALLATION - INSTALLATION
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COMMON
NOZZLE
REMOVAL/INSTALLATION
Austrian Technical Training
School Notes - For Training Purposes Only
ASSEMBLY
The Common Nozzle Assembly or CNA can be removed easily. Place
the special dolly in position, and raise the support up to the CNA,
loosen and remove the 56 bolts which fix the CNA to the engine
flange. Disconnect it. Then lower the assembly and move it away from
the engine. For transportation, it is recommended to attach the CNA to
its support. With the CNA removed it is now possible to see the
exhaust plug in detail. The installation is the reverse procedure.
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Common Nozzle Assembly Removal and Installation.wmv
COMMON NOZZLE ASSEMBLY - REMOVAL/INSTALLATION
TDTI / HAT / ATA 71-80
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AIR INTAKE COWL - REMOVAL/INSTALLATION
WARNING: BE CAREFUL DURING THE REMOVAL AND
INSTALLATION OF THE AIR INTAKE COWL, IT WEIGHS 238
LBS (108 KG). DO NOT GET THE CLEANING FLUID ON YOUR
SKIN, IN YOUR EYES OR IN YOUR MOUTH, THE CLEANING
FLUID IS POISONOUS. FLUSH CLEANING FLUID FROM YOUR
EYES, MOUTH OR SKIN WITH WATER. USE ONLY IN AN AREA
OPEN TO THE AIR.
To remove the air inlet cowl, the fan cowl doors must be opened.
Connect the special sling to a hoist, and attach the sling to the air inlet
cowl hoisting points. Using the hoist, hold up the weight of the cowl
and remove the bolts, then disconnect the pressure tube and electrical
connectors of the P2/T2 sensor and remove the coupling at the antiice duct connection. Move the air inlet cowl FWD to disengage it. Then
carry up and lower it on to its support. The air inlet cowl installation is
the reverse procedure. During the installation take care that the inlet
cowl locating pins engage correctly to the front engine flange. Then
install and tighten the bolts. Finally, reconnect everything up and
remove the sling equipment.
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Air Inlet Cowl Removal and Installation.wmv
AIR INTAKE COWL - REMOVAL/INSTALLATION
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NOTE: THAT THE MOMENT WEIGHT OF THE FAN BLADE IS
IDENTIFIED ON
FAN BLADE - REMOVAL/INSTALLATION
CAUTION: DO NOT LET TOOLS, PARTS OR UNWANTED
MATERIALS FALL ON THE AIR INTAKE COWL SURFACE WHEN
YOU USE THE WORK MAT. TO PREVENT THIS, SEAL THE
EDGES OF THE WORK MAT WITH TAPE BEFORE YOU USE IT.
BE CAREFUL DURING THE REMOVAL OF THE FRONT BLADE
RETAINING RING AND THE STAGE 1 FAN BLADE, THEY EACH
WEIGH APPROXIMATELY 10 LBS (4.5 KG). DO NOT PUT ANY
TOOLS OR PARTS THAT ARE REMOVED IN THE STAGE 1 FAN
DISK INNER SURFACE. BE CAREFUL DURING THE
INSTALLATION OF THE FRONT BLADE RETAINING RING AND
THE
STAGE
1
FAN
B LA D E ,
THEY
EACH
WEIGH
APPROXIMATELY 10 LBS (4.5 KG).
the bottom face of the blade root. For the installation, align the
replacement blade with the disk dovetail and install the two annulus
fillers to their initial position. Reinstall the front blade retaining ring
and secure it. Next install the 22 bolts which fix the annulus filler.
Finally, reinstall the inlet cone and leave the work area making certain
that it is clean and clear of tools and other items.
WARNING: DO NOT LET ENGINE OIL STAY ON YOUR SKIN A
LONG TIME, IT IS POISONOUS AND CAN GO THROUGH YOUR
SKIN AND INTO YOUR BODY.
An individual fan blade can be changed easily if it is damaged, in this
case, the fan blade must be changed with the same moment weight.
First, remove the inlet cone. With the inlet cone removed, the front
blade retaining ring is now accessible to remove it. Remove the 22
bolts which secure the annulus fillers between the fan blades and the
36 bolts which maintain the front blade retaining ring. Next, install
puller bolts at six hole positions, and tighten them. Using these puller
bolts, release the front blade retaining ring from the stage one fan
disk. Turn the fan so that the blade to be removed is at twelve o'clock
position. Then remove the two annulus fillers of the corresponding
blade. Pull the blade forward along the dovetail slot and remove it.
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Fan Blade Removal and Installation.wmv
FAN BLADE - REMOVAL/INSTALLATION
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FAN MODULE - REMOVAL/INSTALLATION
WARNING: BE CAREFUL DURING THE REMOVAL AND
INSTALLATION OF THE FAN MODULE, IT WEIGHS
APPROXIMATELY 395 LBS (179 KG). PUT APPROVED GLOVES
ON YOUR HANDS BEFORE YOU HOLD THE FAN BLADES. THE
LEADING EDGES OF THE FAN BLADES CAN CAUSE INJURY TO
YOUR HANDS. BE CAREFUL DURING THE ASSEMBLY OF THE
CANTILEVER SLING AND THE SLING ADAPTER, THE SLING
WEIGHS 165 LBS (75 KG) AND THE ADAPTER WEIGHS 74 LBS
(34 KG). DO NOT GET ENGINE OIL ON YOUR SKIN FOR A
LONG TIME, THE OIL IS POISONOUS AND CAN GO THROUGH
YOUR SKIN AND INTO YOUR BODY.
CAUTION: DO NOT TOUCH THE STAGE 1 FAN DISK WITH
TOOLS WHEN YOU LOOSEN OR REMOVE THE CURVIC
COUPLING BOLTS. THERE ARE TWENTY-FOUR CURVIC
COUPLING BOLTS WHICH ATTACH THE FAN MODULE TO THE
LP COMPRESSOR/INTERMEDIATE CASE MODULE. DURING
REMOVAL OF THE FAN MODULE, KEEP THREE OF THE BOLTS
IN POSITION UNTIL AFTER THE INSTALLATION OF THE
SUPPORT AND SLINGING TOOLS. OBSERVE THE CORRECT
UNTORQUE PROCEDURE AND PLACE THE HUB PROTECTOR.
MAKE SURE THAT THE QUICK-RELEASE PIN IS INSTALLED IN
THE INNER HOLES IN THE CANTILEVER SLING BRACKET
BEFORE MOVING THE FAN MODULE/SLINGING ASSEMBLY.
PUT TWO PERSONS ON EACH SIDE OF THE FAN MODULE
BEFORE MOVING THE FAN MODULE/SLING TOOL ASSEMBLY.
MAKE SURE THAT THE QUICK-RELEASE PIN IS INSTALLED IN
THE INNER HOLES IN THE CANTILEVER SLING BRACKET
BEFORE MOVING THE FAN MODULE/SLINGING TOOL
TDTI / HAT / ATA 71-80
Issue: 06/08
ASSEMBLY. DO NOT TOUCH THE STAGE 1 FAN DISK WITH
TOOLS WHEN YOU INSTALL OR TIGHTEN THE CURVIC
COUPLING BOLTS. DO NOT USE A POWER WRENCH TO
TIGHTEN THE BOLTS. THE USE OF A POWER WRENCH CAN
CAUSE DAMAGE TO THE BOLTS. WHEN YOU TORQUE THE
CURVIC COUPLING BOLTS, IT IS IMPORTANT THAT THE
TORQUE VALUE IS ON THE INCREASE. A NEW BOLT MUST BE
INSTALLED IF THE TORQUE VALUE IS CONSTANT OR
DECREASES.
For this operation, the air inlet cowl must be removed first, then
remove the inlet cone and its fairing. There are 25 curvic bolts which
attach the fan module. Put hub protectors and remove 21 out of 24
bolts observing the unscrew procedure. The three bolts adjacent to the
guide pins of the stub shaft remain in position. Next, the hub
protectors are removed to install the sling adapter. Then the sling
adapter is fixed. Using the hoist, connect the cantilever sling to the
sling adapter. Secure the sling adapter with a quick-release pin. Using
the adjustment handle, align the hoisting plate to the witness mark.
Remove the remaining three curvic bolts and carefully move FWD the
fan assembly making sure that nothing hits against the fan case. The
fan module is reinstalled with the same precautions. Use the
correlation mark and the U slot of the stub shaft to align the fan disk.
Make sure that the fan module is correctly engaged and secure it with
the curvic bolts at two, four, eight and ten o'clock position. Then
remove the sling adapter with the cantilever sling. Install and tighten
the other curvic bolts. Finally, reinstall the air inlet cowl and the inlet
cone. When leaving the work area, make sure that the air intake is
clear and clean of tools.
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Fan Module Removal and Installation.wmv
FAN MODULE - REMOVAL/INSTALLATION
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ENGINE BORESCOPE - INSPECTION/CHECK
WARNING: IF YOU GET FLUID ON YOUR SKIN OR IN YOUR
EYES, FLUSH IT AWAY WITH CLEAN WATER AND GET
MEDICAL AID. MAKE SURE THAT THE BLANKING PLUG IS
SUFFICIENTLY COOL BEFORE REMOVAL, THE TEMPERATURE
STAYS HIGH FOR A SHORT TIME AFTER ENGINE SHUTDOWN.
IF YOU GET FLUID ON YOUR SKIN OR IN YOUR EYES, FLUSH
IT AWAY WITH CLEAN WATER AND GET MEDICAL AID.
damages. Here the operator is inspecting the HP turbine first stage
vanes through an igniter port, and here we can see the leading edge
of the HP turbine first stage blades. In fact, it is through this unit and
an optic fiber that the pictures can be picked up and transmitted.
CAUTION: BE CAREFUL DURING INSTALLATION AND
REMOVAL OF THE GUIDE TUBE NOT TO CAUSE DAMAGE TO
THE OUTER OR INNER OUTLET GUIDE VANES. THE LIMITS
WHICH FOLLOW ARE APPLICABLE ON A CONTINUING
SERVICE BASIS ONLY. EACH LIMIT IS BASED ON WHAT THE
STRUCTURE OF THE VANE MUST BE AND DOES NOT INDICATE
THAT THERE WILL NOT BE DETERIORATION OF THE ENGINE
PERFORMANCE, STABILITY, OPERATING LIMITS OR PART
REPAIRABILITY. MAKE SURE THAT THE BORESCOPE TIP IS
NOT IN THE PATH OF THE STAGE 1 HIGH PRESSURE TURBINE
(HPT) BLADES BEFORE ROTATING THE ENGINE.
We will now see a presentation of the borescope equipment. First of all
a rotator kit is installed on the cranking pad near the starter for
manual drive of the High Pressure (HP) rotor, then to inspect the
affected sections, the corresponding borescope port covers have to be
removed. Here are the borescope port covers removed from the HP
compressor stage 3, stage 4, stage 7 and 8, stage 9 and 10, stage 11
and 12, and the HP turbine stage 1 and 2, stage 2 and LP turbine, the
borescope inspection requires skill and competence. When inserting
the borescope probe, we can examine the blades, and by turning the
compressor we can detect if there are some cracks, nicks or tip
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Engine Borescope Inspection.wmv
ENGINE BORESCOPE - INSPECTION/CHECK
TDTI / HAT / ATA 71-80
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AIR STARTER - REMOVAL/INSTALLATION
WARNING: DO NOT LET ENGINE OIL STAY ON YOUR SKIN FOR A
LONG TIME, THE OIL IS POISONOUS AND CAN GO THROUGH YOUR
SKIN AND INTO YOUR BODY. BE CAREFUL DURING REMOVAL OR
INSTALLATION OF THE STARTER, IT WEIGHS 33 LBS (15 KG).
CAUTION: DURING INSTALLATION OF THE AIR DUCT, DO NOT
PUT MORE STRESS THAN IS NECESSARY ON THE DUCT. IN
CASE OF MECHANICAL RUPTURE OR FAILURE, THE AIR
STARTER CAN BE EASILY REMOVED ON LINE. FIRST RELEASE
THE CLAMP OF THE LOWER AIR DUCT AT THE AIR STARTER
INLET JUNCTION AND DISCARD THE GASKET. HOLD THE AIR
DUCT AND RELEASE THE CLAMP AT THE START VALVE OUTLET
JUNCTION THEN REMOVE THE AIR DUCT AND THE CLAMPS.
RELEASE THE QUICK-DETACH CLAMP AND DISCONNECT THE
STARTER FROM ITS ADAPTER HOUSING. WITH THE STARTER
REMOVED, IT IS POSSIBLE TO SEE THE DRIVE PAD AND THE
SPLINES OF THE ADAPTER WHERE THE STARTER IS
ENGAGED. ALSO ON THE STARTER YOU CAN SEE THE TURBINE
ROTOR SECTION, THE ENGAGEMENT SPLINES AND DRIVE
SHAFT. THE INSTALLATION OF THE STARTER IS THE REVERSE
PROCEDURE.
NOTE: DURING
INSTALLED.
THIS
TASK,
NEW
TDTI / HAT / ATA 71-80
GASKETS
MUST
BE
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Air Starter Removal and Installation.wmv
AIR STARTER - REMOVAL/INSTALLATION
TDTI / HAT / ATA 71-80
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IDG - REMOVAL/INSTALLATION
WARNING: BE CAREFUL DURING THE REMOVAL OR THE
INSTALLATION OF THE INTEGRATED DRIVE GENERATOR
(IDG), IT WEIGHS 124 LBS (60 KG). DO NOT LET ENGINE OIL
ON YOUR SKIN FOR A LONG TIME, THE OIL IS POISONOUS
AND CAN GO THROUGH YOUR SKIN AND INTO YOUR BODY.
DO NOT TOUCH THE FUEL OR OIL SYSTEM COMPONENTS FOR
A SHORT TIME AFTER THE ENGINE SHUTDOWN, THE ENGINE
COMPONENTS STAY HOT FOR SOME TIME AND CAN CAUSE
INJURY.
Before removing the IDG, the engine must be shut down for at least 5
minutes to prevent burns because the oil stays hot for a few minutes.
First of all, the IDG must be drained and the oil tubes disconnected
and removed. Also remove the banjo couplings and note their initial
positions. Then disconnect the three electrical connectors from the
IDG. Remove the cover plate of the IDG terminal box and disconnect
the output cables. Next place the special dolly in position and raise it
up to the IDG. When the IDG is on the dolly, release the quickattach/detach coupling then pull the IDG FWD to disengage it. Let's
now see the IDG installation. Using the dolly, lift the IDG up to its
position and align the open marks on the quick-attach/detach
coupling. When the IDG is engaged, tighten the coupling sufficiently to
secure it, then remove the dolly, reconnect the electrical connectors
and reinstall the oil tubes.
CAUTION: DO NOT LET ENGINE OIL FALL ON THE ENGINE,
UNWANTED OIL MUST BE REMOVED IMMEDIATELY. THE OIL
CAN CAUSE DAMAGE TO THE SURFACE PROTECTION AND TO
SOME PARTS. HELP THE IDG OUT IN A STRAIGHT LINE, THE
INPUT SHAFT CAN CATCH AND CAUSE DAMAGE TO THE INPUT
SEAL. DO NOT USE EXTERNAL PARTS OR TOOLS TO PULL THE
IDG FROM THE GEARBOX, THEY CAN CAUSE DAMAGE. DO NOT
APPLY SIDE LOADS WHICH COULD BEND THE SPLINED DRIVE
SHAFT DURING THE INSTALLATION OF THE IDG. A BENT
SPLINED
DRIVE
SHAFT
WILL
PREVENT
CORRECT
INSTALLATION OF THE IDG AND DAMAGE IT IN OPERATION.
MAKE SURE THAT YOU DO NOT DAMAGE THE INPUT SEAL
DURING THE INSTALLATION OF THE IDG. MAKE SURE THAT
THE TUBES ARE NOT STRESSED, THIS CAN CAUSE DAMAGE.
THE SQUARE WASHERS MUST BE INSTALLED BETWEEN THE
PHASE CABLES AND THE TERMINAL BLOCK ON THE IDG TO
PREVENT THE TERMINAL BLOCK BECOMING HOT AND
CAUSING DAMAGE TO THE IDG.
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IDG Removal and Installation.wmv
IDG - REMOVAL/INSTALLATION
TDTI / HAT / ATA 71-80
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DECREASE TO 120°F (48,8°C) OR LESS BEFORE YOU DRAIN
THE OIL.
IDG - SERVICING
WARNING: DO NOT LET ENGINE OIL STAY ON YOUR SKIN
FOR A LONG TIME, THE OIL IS POISONOUS AND CAN GO
THROUGH YOUR SKIN AND INTO YOUR BODY. YOU MUST PUT
THE DRAIN HOSE INTO A CONTAINER BEFORE YOU CONNECT
THE DRAIN HOSE TO THE OVERFLOW DRAIN VALVE. MAKE
SURE THAT THE DRAIN HOSE HANGS DOWN VERTICALLY AND
THE END OF THE HOSE IS ABOVE THE OIL LEVEL IN THE
CONTAINER. WHEN YOU CONNECT THE DRAIN HOSE TO THE
IDG, YOU RELEASE THE OIL PRESSURE AND OIL WILL SPRAY
FROM THE DRAIN HOSE. YOU MUST PUT THE DRAIN HOSE ON
THE OVERFLOW DRAIN VALVE BEFORE YOU REMOVE THE
CASE DRAIN PLUG. THE INSTALLATION OF THE DRAIN HOSE
RELEASES THE PRESSURE IN THE IDG OIL SYSTEM. YOU
MUST PUT THE DRAIN HOSE ONTO THE OVERFLOW DRAIN
VALVE BEFORE YOU REMOVE THE CASE DRAIN PLUG. THE
INSTALLATION OF THE DRAIN HOSE RELEASES THE
PRESSURE IN THE IDG OIL SYSTEM.
Do not touch the IDG system components for a short time after the
engine is shut down. The engine components stay hot for some time
and can cause injury. If the oil is below the minimum level, the green
band, add oil to the IDG. Now prepare the IDG for servicing. Remove
the caps from the overflow drain and pressure fill valves. First connect
a drain hose to the overflow drain valve then connect the pressure fill
hose from a service unit to the pressure fill valve. Using the service
units, add oil, until the level reaches the full mark and the oil overflows
through the drain hose. Drain oil until the oil level moves down to the
green band. Remove the drain and pressure fill hoses from connectors
and reinstall the caps. Finally perform a leak test. Do another check of
the oil level 5 minutes later.
NOTE: THIS PROCEDURE IS TO LET THE AIR ESCAPE FROM
THE IDG OIL SYSTEM AND TO LET THE OIL FLOW AROUND
THE SYSTEM.
CAUTION: DO NOT LET ENGINE OIL FALL ON THE ENGINE,
UNWANTED OIL MUST BE REMOVED IMMEDIATELY. THE OIL
CAN CAUSE DAMAGE TO THE SURFACE PROTECTION AND TO
SOME PARTS. THE DRAIN HOSE MUST BE CONNECTED TO LET
THE OIL IN THE IDG DRAIN TO THE CORRECT LEVEL. THE IDG
CAN BECOME TOO HOT IF IT CONTAINS TOO MUCH OIL. DO
NOT OPERATE THE IDG WITHOUT SUFFICIENT OR WITH TOO
MUCH OIL, THIS WILL CAUSE DAMAGE TO THE IDG. WHEN
YOU SERVICE THE IDG OIL SYSTEM, DO NOT MIX TYPES OR
BRANDS OF OIL. LET THE TEMPERATURE OF THE IDG
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IDG Servicing.wmv
IDG - SERVICING
TDTI / HAT / ATA 71-80
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