IAE-V2500 Engine for Airbus Family 319/320

Airbus A319/A320/A321 (IAE V2500A) vs A318/A319/A320/A321 (CFM56) Training Manual (EASA Part. 66 Cat. B1)
Issue 2 / September 2008 / Technical Training

Contents - I
Sep08/Technical Training
Copyright by SR Technics
Table of Contents EASA Part 66 Cat. B1
Training Manual
A319/A320/A321
for training purposes only
71 Power Plant - V2500A 71-00 Introduction Engine Mark Numbers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
IAE V2530-A5 Data. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Safety Zones. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Engine Inlet Hazard Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Jet Wake Hazard Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Noise Danger Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 71-00 Nacelle Access Doors & Openings Nacelle General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Access Doors & Openings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Fan Cowls Opening / Closing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Nacelle D/O. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Nacelle D/O. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Fan Cowl Latch Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
Thrust Reverser Cowl Doors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
T/R Cowling ("C-Duct") Opening / Closing . . . . . . . . . . . . . . . . . . . . . . . . . 20
Thrust Reverser Half Latches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Latch Access Panel & Take Up Device . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Front Latch and Open Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
C - Duct Opening / Closing System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
C - Duct Hold Open Struts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30 71-00 Engine Mounts
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
Forward Engine Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
AFT Engine Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
Engine Change. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
Engine Removal / Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
Nacelle D/O. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
71-00 Power Plant Drains
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
Drain System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46
Pylon Drains . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
72 Engine - V2500A 72-00 Engine Presentation Engine Main Bearings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Front Bearing Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
NO 4 Bearing Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Rear Bearing Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Engine Modules. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Module 31 (Fan Module) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Inlet Cone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Front Blade Retaining Ring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Fan Blade Removal / Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
Annulus Fillers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Reposition of the Annulus Filler Seals. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
72-31-11 Fan Blade Repair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Fan Blade Inspection / Repair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Repair of the Fan Disk Rear Ramp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35
TAP Transient Acoustic Propagation Test . . . . . . . . . . . . . . . . . . . . . . . . . 37
Fan Trim Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
One Shot Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
Module 32 Intermediate Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Module 40 HP Compressor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
Combustion Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47
Common Nozzle Assembly (CNA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49
Angle and Main Gearbox. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51
Drive Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
Borescoping . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
Borescope Inspection of the HP Comp. . . . . . . . . . . . . . . . . . . . . . . . . . . . 57
Borescope Access. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58 73 Engine Fuel and Control - V2500A

Sep08/Technical Training
Copyright by SR Technics
for training purposes only
Training Manual
A319/A320/A321
Table of Contents EASA Part 66 Cat. B1
Contents - II
73-00 Fuel System Presentation General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
73-10 Fuel Distribution Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Fuel Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Fuel Filter Diff. Press. Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Fuel Temperature Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Fuel Diverter & Return Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Fuel Distribution Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Fuel Manifold and Tubes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Fuel Nozzle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Fuel Pump. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Fuel Metering Unit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Fuel Metering Unit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Overspeed Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Low Pressure Fuel Shut Off Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
HP & LP Fuel SOV Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
73-20 Heat Management System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Presentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Fuel Temp. Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
IDG Oil Cooler Temp. Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
ACOC Oil Temp. Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
ACOC Modulating Air Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Fuel Diverter & Return Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Return to Tank Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
HMS Mode 1 (Normal Mode) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
HMS Mode 4 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
No Return to Tank Modes 3 and 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
HMS Mode 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
HMS Mode 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Air Modulating Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
IDG Fuel Cooled Oil Cooler . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
IDG Oil Cooler Temp. Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
FADEC Presentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
FADEC System Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
FADEC Functions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
Engine Control Pushbuttons and Switches . . . . . . . . . . . . . . . . . . . . . . . . . 38
Failures and Redundancy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Engine Limits Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Power Management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
Autothrust Activation / Deactivation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
EPR Setting Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46
Rated N1 Setting Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
The processing of the N1 error signal is the same as for EPR error signal. 48
Unrated N1 Setting Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
The processing of the N1 error signal is the same as for the rated N1 error sig-
nal.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
FADEC Fault Strategy. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52
Component Fail Safe States . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54
Loss of Inputs from Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
Idle Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56
N1 Speed Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58
FADEC Power Supply. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60
FADEC LRU‘S. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
Data Entry Plug Modification. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
Electronic Engine Control (EEC). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64
FADEC Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66
FADEC LRU‘S Sensors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66
FADEC LRU‘S Sensors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68
P12.5 Sensor. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70
P2.5 / T2.5 Sensors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70
FADEC Test. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72
FADEC Previous Legs Report. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72
FADEC Troubleshooting Report . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
FADEC Failure Types Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
FADEC System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78
FADEC Ground Scanning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80
FADEC Class 3 Fault Report . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81
Scheduled Maintenance Report . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
Engine Interface Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84
EIU Presentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84

Contents - III
Sep08/Technical Training
Copyright by SR Technics
Table of Contents EASA Part 66 Cat. B1
Training Manual
A319/A320/A321
for training purposes only
EIU Input Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84
EIU Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86
CFDS System Report/Test EIU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87
LAST Leg Report . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88
LRU Indentification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88
Ground Scanning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
EIU CFDS Discrete Outputs Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . 90
EIU CFDS Discrete Outputs Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . 92
EIU Discrete Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94
74 Ignition - V2500A 74-00 Ignition System Presentation General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Ignition System Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Ignition Starting - Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Ignition System Circuit Breakers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Ignition System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Ignitor Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Ignition Test without CFDS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 74-00 Starting 80-00 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Starting Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Starting Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Starter Air Control Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Start Air Control Valve Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Cranking-Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Wet Cranking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Automatic Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
EEC Auto Start Abort . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
Manual Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Continuous Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30 75 Engine Air - V2500A
75-00 System Presentation
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
FADEC Compressor and Clearance Control. . . . . . . . . . . . . . . . . . . . . . . . . 2
Compressor Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
75-31 LP Comp. Air Flow Sys.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Booster Bleed System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
BSBV Actuating Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
75-32 HP Comp. Air Flow Sys. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
VSV System Components. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
VSV Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Handling Bleed Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Handling Bleed Valves Function . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Bleed Valve Locations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
Handling Bleed Valve Malfunctions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
HP Turbine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Turbine Cooling Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Operating Schedule. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
HPT / LPT Active Clearance Cont. Sys. . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
HPT / LPT Cooling Manifolds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Nacelle Ventilation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
75-41 Nacelle Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
Nacelle Temperature General. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
76 Engine Controls - V2500A 76-00 Engine Controls
Throttle Control System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Thrust Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Bump Rating Push Button(A1 Engined Aircraft only) . . . . . . . . . . . . . . . . . . 4
Artificial Feel Unit (Mechanical Box) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Throttle Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
AIDS Alpha Call Up of TRA. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
77 Indicating - V2500A

Sep08/Technical Training
Copyright by SR Technics
for training purposes only
Training Manual
A319/A320/A321
Table of Contents EASA Part 66 Cat. B1
Contents - IV
77-00 Engine Indicating Presentation
Indication General. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
77-10 Power Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
EPR Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
EPR System Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
P2 / T2 Heater. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
FADEC P2/T2 Heater Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
77-20 Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
EGT Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
EGT Probes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
77-10 Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
N1 and N2 Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
31 Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Max Pointer Reset (N1, N2 & EGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
77-10 Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
N1 Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Interchange of N1 Speed Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Dedicated Alternator (PMA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
77-30 Analyzers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Vibration Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Engine Vibration Monitoring Unit (EVMU). . . . . . . . . . . . . . . . . . . . . . . . . . 24
Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
CFDS System Report / Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
CFDS System Report /Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
CFDS System Report /Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
CFDS System Report /Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
CFDS Accelerometer Reconfig. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
78 Exhaust - V2500A 78-00 Reverser System
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Thrust Reverser System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Thrust Reverser System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Thrust Reverser Hydraulic Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Thrust Reverser Manual Deployment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
Thrust Reverser Independent Locking System . . . . . . . . . . . . . . . . . . . . . . 10
Component Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Reverser Hydraulic Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
HCU in Forward Thrust Position . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
HCU Deploy Sequence Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
HCU Stow Sequence Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Command Limitation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Flexshaft Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
Hydraulic Actuators Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Upper Nonlocking Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
Lower Locking Actuators. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
Thrust Reverser Manual Deploy / Stow. . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Thrust Reverser Deactivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
FADEC CFDS Reverser Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
FADEC T/R Test (Fault Detected). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
FADEC T/R Test (NOT O.K.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35
79 Oil - V2500A 79-00 Oil System Oil System Presentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Oil System Bearings and Gears Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . 6
Front Bearing Compartment (Bearings no. 1, 2, 3) . . . . . . . . . . . . . . . . . . . . 6
Centre Bearing Compartment (Bearing no.4) . . . . . . . . . . . . . . . . . . . . . . . . 8
Rear Bearing Compartment (Bearing no.5). . . . . . . . . . . . . . . . . . . . . . . . . 10
Oil System Components Presentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Oil Tank. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Oil Quantity Transmitter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Oil Pressure Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Air Cooled Oil Cooler (ACOC). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
ACOC Oil Temperature Thermocouple. . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
Fuel Cooled Oil Cooler (FCOC) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Scavenge System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Scavenge Pumps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Scavenge Oil Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
De-oiler . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
No 4 Bearing Scavenge Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

Contents - V
Sep08/Technical Training
Copyright by SR Technics
Table of Contents EASA Part 66 Cat. B1
Training Manual
A319/A320/A321
for training purposes only
No 4 Bearing Pressure Transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
No 4 Bearing Scavenge Valve Description . . . . . . . . . . . . . . . . . . . . . . . . . 28
No 4 Bearing Scavenge Valve Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Engine Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Oil System Pressure Sensing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
Low Oil Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
Magnetic Chip Detectors (M.C.D.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
Master Chip Detector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36
IDG Oil Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
79-30 Oil Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
ECAM Oil Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
Oil Quantity Indicating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Oil Temperature Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Oil Pressure Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Low Oil Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
Scav. Filt. Diff. Pressure Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
24 Electrical Power - V2500A 24-22 AC Main Generation
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator Drive . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Speed Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Control and Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator Control Unit Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator Operation Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator 1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Generator Reset . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Integrated Drive Generator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Servicing of IDG . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
AC Main System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
26 Fire Protection - V2500A
26-12 Engine Fire and Overheat Detection
Fire Detectors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Fire Detection Unit (FDU) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Warnings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Test P/B. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Engine Fire Detection Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Fire Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Loop Fault Warning. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Detection Fault Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Electrical Circuits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Fire Detection Circuit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Fire Extinguishing Circuit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Engine Fire Pushbutton Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15
26-99 CFDS System Report / Test
FDU - Bite . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
30 Ice and Rain Protection - V2500A 30-00 Eng. Air Intake Ice Protection
System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
System Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
System Control Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Engine Anti Ice Duct and Valve. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
Anti-Ice Valve Deactivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
36 Pneumatics - V2500A 36-10 General Distribution - Description and Operation. . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
HP Bleed Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Pressure Regulator Valve (PRV). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
Bleed Temperature Control. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
System Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

Sep08/Technical Training
Copyright by SR Technics
for training purposes only
Training Manual
A319/A320/A321
Table of Contents EASA Part 66 Cat. B1
Contents - VI
BMC Bleed Monitoring Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
High Pressure Bleed Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
Bleed Pressure Regulator Valve (PRV) . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
Overpressure Valve (OPV) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
Fan Air Valve (FAV) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
Fan Air Valve Control Thermostat CT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
Temperature Limitation CTS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Bleed Transfer Regulated Pressure Transducers Pt . . . . . . . . . . . . . . . . . 15
Temperature Control Description and Operation . . . . . . . . . . . . . . . . . . . . 16
CFDS MCDU Pages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
IAE V2500-Study Questions

71-00-1
Power Plant V2500A
71-00 Introduction
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
71 Power Plant - V2500A 71-00 Introduction It is produced by International Aero Engines (IAE) corporation.
On March 11, 1983 five of the world’s leading aerospace manufacturers signed a
collaboration agreement to create, for the firs t time in history, a new family of aero
engines developed form the best proven technology that each of the five could
provide.
Headquarters for IAE were established in Connecticut, USA, and from there the
V2500 turbofan engine, designed to power the world’s 120-180 seat aircraft, was
launched on January 1, 1984.
Shared Technology, shared Strenght
Each shareholder is responsible for the development and production of discrete
modules reflecting their best proven technology.
The senior partners Rolls Royce and Pratt & Whitney assemble the engines at
their respective plants in Derby, UK and Middletown Connecticut, USA
Fiat Aviazone have since withdrawn as a risksharing partner, but still remains as
a Primary Supplier. Rolls Royce now has responsibility for all external gearbox re-
lated activity.
IAE is responsible for the coordination of manufacture and assembly of the en-
gines, sales, marketing, contractin g and in-service support of V2500.
The engine entered revenue service on May 22, 1989.
This corporation consits of the following companys:
• JAEC (Japanese Aero Engines Corporation)
• Rolls Royce
• Pratt & Whittney
• MTU (Motoren & Turbinen Union)
Pratt & Whitney
32.5%

Diffuser-Combustor, High Pressure Turbine,
Turbine Exhaust Case
Rolls-Royce
32.5%

High Pressure Compressor, Gear Box

Japanese Aero Engines Corporation
23%

Fan Case, Low Pressure Compressor
MTU Aero Engines
12%

Low Pressure Turbine Module

71-00-2
Training Manual
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Power Plant V2500A
71-00 Introduction
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Engine Mark Numbers The V2500 engine has been designated the “V” because IAE was originally a five-
nation consortium. The “V” is the Roman numeral for five.
For easy identification of the present and all future variants of the V2500, Interna-
tional 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 de signation are V25, identifying each en-
gine in the family
• The next two figures indicate the engine’s rated sea - level takeoff thrust. The
following letter shows the aircraft manufacturer.
• The following letter shows the aircraft manufacturer.
• The last figure represents the mechanical standard of the engine.
This system will provide a clear designation of a particular engine as well as a sim-
ple way of grouping by name, engines with similar characteristics.
The designation V2500 - D collectively describes, irrespective of thrust, all en-
gines for McDonnell Douglas applications and V2500 - A all engines for Airbus In-
dustrie.
Similarly, V2500 - 5 describes all engines built to the -5 mechanical standard, ir-
respective of airframe application.
The only engine exempt from this idents is the current service engine, which is al-
ready certified to the designated V2500-A1.
For example:
The V2500 - A1 engine is used on A320 and has only a 3 stage booster.
The D5 variant is now no longer in production, however the engine is still
extensively overhauled and re-furbished

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71-00 Introduction
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Figure 1: Engine Mark Numbers
V2530-A5
Generic to all
V2500 engines
Takeoff thrust in
thousands of
pounds
Mechanical Standarts
of engine
Airframe manufacturer
- A for Airbus Industrie
- D for McDonnell Douglas
MARK NUMBER
TAKEOFF THRUST (LB)
AIRCRAFT
V2500 - A1
V2522 - A5
25.000
22.000
A320 - 200
A319
V2530 - A5
30.000
A321 - 100
V2525 - A5
25.000
A320 - 200
V2527 - A5
26.500
A320 - 200
V2528 - D5
28.000
MD - 90 - 40
V2525 - D5
25.000
MD - 90 - 30
V2522 - D5
22.000
MD - 90 - 10

71-00-4
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Power Plant V2500A
71-00 Introduction
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Copyright by SRTechnics
Introduction The V2530 - A5 engine is a two spool, axial flow, high bypass ratio turbofan en-
gine.
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 fivestage
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.
IAE V2530-A5 Data The IAE V2530-A5 engine is flat rated.
The rated thrust can be obtained for a limi ted time up to an ambient temperature
of 30C otherwise engine operating limits can be exceeded.
To have a constant thrust at variable ambient conditions the engine RPM has to
be adjusted (regulated) to compensate the variying air density.
The Thrust parameter is EPR. In case this parameter is not available the N1 is
used as the Thrust parameter.
Fan tip diameter: 63.5 in (161cm)
Bare engine length: 126 in (320 cm)
Weight: 4942 lbs (2242 KG)
Take - off thrust: 30,000 lb, flat rated to +30 deg. C
Bypass ratio: 5.44 : 1
Overall Pressure Ratio: 31.9 : 1
Mass Flow lbs/s: 856 lbs
N1: 100% (5650 RPM)
N2: 100% (14950 RPM)
EGT (Takeoff) 650 deg. C
EGT (Starting) 635 deg. C
EGT (Max Continous/Climb) 610 deg. C

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Figure 2: V2500 Propulsion Unit

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71-00 Introduction
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Safety Zones Engine Inlet Hazard Areas Warning: During run up operations, extreme care should be exercised when operating the
engines. Refer to the diagram showing the inlet suction hazard areas for the con-
ditions at idle and take-off thrust.
Figure 3: Engine Inlet Hazard Areas
Jet Wake Hazard Areas Warning: During run up operations, extreme care should be exercised when operating the
engines.
Refer to the diagram showing the jet wake hazard areas for the conditions at idle
and take-off thrust. Noise Danger Areas Warning: Ear protection must be worn by all persons working near the engine while it oper-
ates.
Loud noise from the engine can cause temporary or permanent damage to the
ears.

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Figure 4: Jet Wake Hazard Areas

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Power Plant V2500A
71-00 Nacelle Access Doors & Openings
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Copyright by SRTechnics
71-00 Nacelle Access Doors & Openings Nacelle General The nacelle ensures airflow around the engine during its operation and also pro-
vides protection for the engine and accessories.
The major components which comprise the nacelle are:
• the air inlet cowl
• the fan cowls (left and right hand)
• The "C" ducts which incorporate the hydraulically operated thrust reverser unit.
• the Combined Nozzle Assembly (CNA) Access Doors & Openings Access to units mounted on the low pressure compressor (fan) case and external
gearbox is gained by opening the hinged fan cowls.
Access to the core engine, and the units mounted on it, is gained by opening the
hinged "C" ducts.
Pressure relief Doors:
Two access doors also operate as pressure relief doors. They are installed on
each nacelle.
• The air starter valve and pressure relief door in the right fan cowl
• and the oil fill and sight glass pressu re relief door in the left fan cowl.
The two pressure relief doors protect the core compartment against a differential
overpressure of 0.2 bar (2.9007 psi) and more.
Spring-loaded latches hold the doors in place. If overpressure causes one or the
two doors in a nacelle to o pen during flight, they will no t latch close again automat-
ically. The door (doors) will be found open during ground inspections.
Figure 5: Access Doors & Openings
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71-00-9
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71-00 Nacelle Access Doors & Openings
Training Manual
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Figure 6: Nacelle Access Doors
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71-00-10
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71-00 Nacelle Access Doors & Openings
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Fan Cowls Opening / Closing The fan cowl doors extend rearwards from the inlet cowl to overlap leading edge
of the "C" ducts. When in the open position the fan cowls are supported by two
telescopic hold - open struts, using support points provided on the fan case (rear)
and inlet cowl (front). Storage brackets are provided to securely locate the struts
when they are not in use.
The fan cowl hold open struts must be in the extended position and both
struts must always be used to hold the doors open.
Be careful when opening the doors in winds of more than 26 knots (30mph)
The fan cowl doors must not be opened in winds of more than 52 knots
(60mph)

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Figure 7: Power Plant Installation Presentation - General

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Power Plant V2500A
71-00 Nacelle Access Doors & Openings
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Figure 8: Nacelle D/O - Air Intake Cowl

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Figure 9: Nacelle D/O - Fan Cowl Doors (LH & RH)

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71-00 Nacelle Access Doors & Openings
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Nacelle D/O Thrust Reverser "C" Ducts The thrust reverser "C" ducts are in two halves fitted with cascades, blocker doors
and translating sleeves.
Each half is supported by four hinges at the pylon.
The halves assembly is latched along the bottom centerline with six latches.
LH door weight: 580 lbs (263 kg).
RH door weight: 574 lbs (260 kg).
Each half is provided with:
• 3 attachment points for handling,
• 1 opening actuator operated with a hand pump,
• 2 hold open rods for opening.
The latch assembly consists of:
• 1 forward bumper latch,
• 3 center latches, accessible through a hinged access panel,
• 1 aft twin latch.

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Training Manual
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Figure 10: Nacelle D/O - Thrust Reverser "C" Ducts

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71-00 Nacelle Access Doors & Openings
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Nacelle D/O Common Nozzle Assembly The Common Nozzle Assembly (CNA) mixes the exhaust gases from the second-
ary and primary airflows.
It is bolted to the rear flange of the turbine exhaust case.
The Common Nozzle Assembly is attached to the LP turbine frame by means of
56 bolts.
Weight: 181 lbs (82 kg). Exhaust Cone The exhaust cone provides the inner contour of the common exhaust stream flow.
It is attached to the inner flange of the turbine exhaust case.
The exhaust cone is bolted to the inner LP turbine frame by means of 13 bolts.
Weight: 10 lbs (4.5 kg).

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Figure 11: Nacelle D/O - Mixed Exhaust System

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Power Plant V2500A
71-00 Nacelle Access Doors & Openings
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Figure 12: Fan Cowls Opening / Closing

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71-00 Nacelle Access Doors & Openings
Training Manual
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Fan Cowl Latch Adjustment The mismatch between the two cowl doors can be adjusted by fitting / removing
shims, as shown below.
Latch tension is adjusted by use of the ad justing nut at the back of the latch keeper
Figure 13: Fan Cowl Latch Adjustment

71-00-20
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Power Plant V2500A
71-00 Nacelle Access Doors & Openings
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Copyright by SRTechnics
Thrust Reverser Cowl Doors T/R Cowling ("C-Duct") Opening / Closing Before opening:
1. All 6 latches & take - up devices must be released in sequence.
2. If reverser is deployed, pylon fairing must be removed.
3. Deactivate Thrust Reverser Hydraulic Control Unit (HCU)
4. FADEC power "OFF"
5. Put Warning Notices in the Cockpit
Figure 14: Thrust Reverser Hydraulic Control Unit (HCU)
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,/#+/54 0).
,/#+/54 ,%6%2 ,/#+ 0/3)4)/.

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Power Plant V2500A
71-00 Nacelle Access Doors & Openings
Training Manual
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Figure 15: C-Duct Opening/Closing
THE FAIRING MUST BE REMOVED
BEFORE THE REVERSER IS DEPLOYED
AND THE CDUCT OPENED

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71-00 Nacelle Access Doors & Openings
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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)

71-00-23
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Training Manual
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Figure 16: Thrust Reverser Half Latches
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$

71-00-24
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Power Plant V2500A
71-00 Nacelle Access Doors & Openings
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Latch Access Panel & Take Up Device An access panel, as shown below, is provided to gain access to the three BIFUR-
CATION "C" duct latches and the "C" duct take up device (also called, Auxiliary
Latch Assembly).
The take up device is a "turnbuckle" arrangement which is used to draw the two
"C" ducts together. This is necessary to compress the "C" duct seals far enough
to enable the latch hooks to engage with the latch keepers.
The take up device is used both when closing and opening the "C" ducts.
The take up device must be disengaged and returned to its stowage bracket, in-
side the L/H "C" duct, when not in use.
Red Open Flags, installed on the C-Duct indicate that the Bifurcation latches
are open.

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Training Manual
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Figure 17: Latch Panel & Take Up Device

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71-00 Nacelle Access Doors & Openings
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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 propper closed reverser cowl.
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.(ta ke up device) If you do not do this, the front latch
can get caught between the thrust reverser halves and the auxiliary latch as-
sembly and the hook can get damaged.

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Figure 18: Front Latch with Open 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 hydrau-
lic connection for opening.
The hydraulic fluid used in the system is engine lubricating oil.

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Figure 19: "C" Duct Opening/Closing

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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. ’C’ duct is similar.
Both struts must always be used to support the ’C’ ducts in the open posi-
tion. The ’C’ ducts weigh approx 578 lbs each. Serious injury to personnel
working under the ’C’ ducts can occur if the ’C’ duct is suddenly released.

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Figure 20: „C“ Duct Hold Open Struts
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71-00-32
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71-00 Engine Mounts
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71-00 Engine Mounts General The engine is attached to the aircraft pylon by two mount assemblies, one at the
front and one at the rear of the engine. The mount assemblies transmit loads from
the engine to the aircraft structure.
Spherical bearings in each mount permit thermal expansion and some movement
between the engine and the pylon.
Both mounts are made to be fail-safe and have a tolerance to damage.
• the forward mount:
it is attached to the engine via the intermediate casing. It takes the X loads
(thrust), Y loads (lateral) and Z loads (vertical).
• the aft mount:
it is attached to the engine via the exhau st casing. It takes the loads in a plane
normal to the engine centerline i.e.: Y loads (lateral), Z loads (vertical) and Mx
(engine rotational inertia moment + Y load transfer moment). Component Location The front mount is installed at the top center of the low pressure compressor case.
The rear mount is installed at the top center of the low pressure turbine case.
The engine mount system has these components:
• A front mount
• A rear mount. Forward Engine Mount The front mount has these parts:
• Two thrust links.
• A beam assembly.
• A cross beam assembly.
• A support bearing assembly.
The thrust links attach to lugs on the cross beam and to the engine mount lugs on
the low pressure compressor using solid pins. A spherical bearing is installed at
each end of the links. Vertical and side loads are transmitted through the support
bearing to the beam assembly and then to the aircraft pylon.
The beam assembly is aligned on the aircraft pylon by two shear pins and at-
tached with five bolts.
The thrust of the engine is transmitted th rough the thrust links, the cross beam as-
sembly and the beam assembly to the aircraft pylon.
The support bearing permits the engine to turn so that torsional loads are not
transmitted to the aircraft structure.
The front mount is made to be fail-safe. If one of the two thrust links or the cross
beam should fail, then thrust loads are transmitted through the ball stop and into
the beam assembly. The thrust is then transmitted to the pylon structure.

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Figure 21: Forward Engine Mount

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AFT Engine Mount The aft mount has these parts:
• Two side links.
• A center link.
• A beam assembly.
The two side links attach to the beam assembly at one end and the engine aft
mount ring on the low pressure turbine case at the other end.
The aft mount is aligned on the pylon by two shearpins and is attached to the pylon
by four bolts and washers.
Vertical and side loads are transmitted through the side links and beam assembly
and into the pylon.
Torsional loads are transmitted by the center link to the beam assembly and in to
the pylon.
The mount is made to be fail-safe. The side links are each made up of two parts
which are attached together to make one unit. If one part of the link should fail, the
remaining part will transmit the loads to the beam assembly.

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Figure 22: AFT Engine Mount

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Engine Change Engine Removal / Installation The arrangements for slinging / hoisting the engine are shown below (Bootstrap).
During this operation the "C" ducts are supported by rods which are posi-
tioned between the "C" duct and the engine pylon.
After a new engine was installed different Test Tasks have to be performed:
• Check of engine datas via CFDS (ESN, EEC P/N, Engine Rating, Bump level)
to make sure that they are the same as written on the EEC, data entry plug and
engine identification plates.
• Operational Test of EEC via CFDS.
• If A/C is operated in actual CAT III co nditions, a Land Test must be performed.
• Functional check of IDG disconnect system.
• Functional check of engine ice protection system.
• TEST NO. 1 (Dry motor leak check)
• TEST NO. 2 (Wet motor leak check)
• TEST NO. 3 (Idle leak check)
• TEST NO. 6 (EEC system idle test)
• TEST NO. 13 (Prestested engine replacement test)
For further information refer to AMM ATA 71-00-00.

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Figure 23: Hold Open Braces and Adjustable Struts
ADJUSTABLE
STRUT
T/R OPENING
ACTUATOR
THRUST REVERSER
COWL DOOR
PYLON

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Figure 24: Engine Removal

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Figure 25: Bootstrap Equipment

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Nacelle D/O Fluid Disconnect Panel The fluid disconnect panel provides the fluid connection between engine and py-
lon. It is located on the left hand side of the fan case upper part.
Fluid connection lines:
FUEL SYSTEM
• fuel supply,
• fuel return to tank,
HYDRAULIC SYSTEM
• hydraulic pump suction,
• hydraulic pump pressure delivery,
• hydraulic pump case drain.
Figure 26: Fluid Disconnect Panel

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Fan Electrical Connector Panel The fan electrical connector panel provid es the interface between the fan electri-
cal harnesses and the pylon.
It is located on the right hand side of the fan case upper part.
Figure 27: Fan Electrical Connector Panel

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Core Electrical Junction Box The core electrical junction box provides the interface between the core electrical
harnesses and the pylon.
It is located in the forward mount zone.
Figure 28: Core Electrical Junction Box

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71-00 Power Plant Drains

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Power Plant V2500A
71-00 Power Plant Drains
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
General The powerplant drain system collects fluids that may leak from some of the engine
accessories and drives. The fluids collected from the power plant are discharged
overboard through the drain mast installed below the engine accessory gearbox.
The drain system comprises two sub-systems:
• fuel drains
• oil, hydraulic and water drains
The two sub-systems come together at the same drain mast.

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Power Plant V2500A
71-00 Power Plant Drains
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for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 29: Drain Mast

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Training Manual
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Power Plant V2500A
71-00 Power Plant Drains
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Drain System Description Fuel Drain The fuel drain lines come from engine ac cessories on the engine core, the engine
fan case and gearbox. The engine core drains go through the bifurcation panel.
The fuel drain system is connected to these engine accessories:
• Booster bleed master actuator)
• Booster bleed slave actuator) Engine- Variable Stator Vane Actuator) Core
• Active Clearance Control Actuator)
• Fuel diverter valve) Engine fan Case
• Fuel metering unit) Gearbox
• LP/HP fuel pumps) Oil, Hydraulic and Water Drains The oil, hydraulic and water drains system comes from engine accessories on the
engine fan case and gearbox.
The drain system is connected to these engine accessories:
• Air Cooled Oil Cooler actuator) Engine fan case
• Integrated Drive Generator)
• Air starter) Gearbox
• Hydraulic Pump)
• Oil tank scupper) Oil tank
The only hydraulic fluid drain is from th e hydraulic pump. The other drains are for
engine oil or accessory lubricant.

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71-00 Power Plant Drains
Training Manual
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Copyright by SRTechnics
Figure 30: Drain System
,%&4 3)$%
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05-03
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#/2% $2!).

71-00-48
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Power Plant V2500A
71-00 Power Plant Drains
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Copyright by SRTechnics
Pylon Drains The engine pylon is divided into 7 compartments. Various systems are routed
through these areas.
Any leckage from fluid lines is drained overboard through seperate lines in the rear
of the pylon.

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71-00 Power Plant Drains
Training Manual
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for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 31: Pylon Drains

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Power Plant V2500A
71-00 Power Plant Drains
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Copyright by SRTechnics

72-00-1
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Copyright by SRTechnics
72 Engine - V2500A

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Power Plant V2500A
72-00 Engine Presentation
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Copyright by SRTechnics
72-00 Engine Presentation Gas Path A simplified view of the engine is shown below.
All the air entering the engine passes trough the inlet cowl to the fan.
At the fan exit the air stream divides into two flows:
• the core engine flow
• the by-pass flow Core Engine Flow The core engine flow passes trough the fixed inlet guide vanes to the L.P.
Compressor which consits of 4 stages on the V2500 - A5 engine, then to the H.P.
Compressor, the combustion section and the H.P. and L.P. turbines and finally ex-
hausts into the Common Nozzle Assembly (C.N.A.) By-pass Flow The fan exhaust air (cold stream) enteri ng the by-pass duct passes through the
fan outlet guide vanes and flows along the by-pass duct to exhaust into the C.N.A. Nacelle The nacelle ensures airflow around the engine during its operation and also pro-
vides protection for the engine and accessories.
The major components which comprise the nacelle are:
• the air inlet cowl
• the fan cowls (left and right hand)
• The "C" ducts which incorporate the hydraulically operated thrust reverser unit.
• the Combined Nozzle Assembly (CNA) Common Nozzle Assembly (CNA) The core engine "hot" exhaust and the "cool" by-pass flow are mixed in the C.N.A.
before passing through the single propelling nozzle to atmosphere.

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72-00 Engine Presentation
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Figure 1: Engine Components Location (L/H Side)
REAR
ENGINE MOUNT
HP COMPRESSOR
SECTION
FUEL COOLED
OIL COOLER
FAN
CASE
NOSE CONE
FUEL
FILTER
OIL
TANK
HYDRAULIC
PUMP
GEARBOX
OIL
PUMP
FUEL PUMP
STAGE 7C
BLEED VALVE
COMBUSTION
SECTION
COMMON
NOZZLE

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Power Plant V2500A
72-00 Engine Presentation
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Figure 2: Engine Components Location (R/H Side)
STARTER
LP COMPRESSOR
(FAN)
TURBINE SECTION
STAGE 7
BLEED VALVES
RELAY
BOX
ELECTRONIC
ENGINE CONTROL
DE-OILER
NO.4 BEARING
COMPARTMENT
AIR COOLER
BLEED VALVE
CONTROL VALVES
INTEGRATED
DRIVE
GENERATOR
AIR COOLED OIL COOLER

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Figure 3: Propulsion Unit Outline
V2500-A5
V2500-D5
V2500-D5
V2500-A1
V2500-A5V2500-A1
COWL
FAN BLADES
SPLITTER FAIRING
GUIDE VANES
1 AIR INLET
2 LP COMPRESSOR
3 LP COMPRESSOR
PYLON
WING
COLD STREAM
COLD STREAM
OUTLET
4 LP COMPRESSOR
5 HP COMPRESSOR
10 INLET CONE
9 INLET CONE FAIRING
8 LP COMPRESSOR CASE
7 LP COMPRESSOR STAGE 1.5
AND 2 VANES
6 LP COMPRESSOR STAGE 1.5,
2, 2.3 AND 2.5 BLADES
HOT STREAM
63” 63.5”

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72-00 Engine Presentation
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STAGE NUMBERING V2530-A5
STAGES :
COMPONENT :
STAGE NUMBER :
NOTES :
1
FA N
1
ACOC,ACC,ACAC
1
2
3
4
LOW PRESSURE
COMPRESSOR
( BOOSTER )
1,5
2
2,3
2.5
B.S.B.V.
1
2
3
5
6
7
8
9
10
HIGH PRESSURE
COMPRESSOR
3
4
5
7
8
9
10
11
12
VSV ( & IGV )
VSV
VSV
CUST. BLEED, A / I, Hdlg. Bleed,
Internal Cooling
CUST. BLEED Hdlg. Bleed,
Buffer Air, 1. HPT & NGV, Muscl Air
COMBUSTION CHAMBER
20 Fuel Nozzles, 2 Ignitor Plugs
1
2
HIGH PRESSURE
TURBINE
1
2

ACTIVE CLEARANCE CONTROL
1
2
3
4
5
LOW PRESSURE
TURBINE
3
4
5
6
7
ACTIVE CLEARANCE CONTROL
COMMON NOZZLE
(Booster Stage
Bleed Valve 2.5 Bleed Ring) =

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72-00 Engine Presentation
Training Manual
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Figure 4: Stage Numbering

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A319/A320/A321
Power Plant V2500A
72-00 Engine Presentation
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Flowpath aerodynamic stations have been established to facilitate engine per-
formance assessment and monitoring.
The manufacture uses numerical station designations. The station numbers are
used as subscripts when designating different temperatures and pressures,
throughout the engine.
STA Designation Designation
often used
Abbreviations
Engine Trend
Monitorin
g

1
)

0 Ambient P0 Pamb
1 Intake Lip of Air Intake – – – –
P2 FIP
T2 FIT
P2.5 CIP
T2.5 CIT
P3 Pb, CDP
T3 CDT
4 Combustion Section Exit – – –
4.5 HP Turbine Exit – – –
P4.9 (P5)
T4.9 EGT
5 Turbine Exhaust Case Exit – – –
12.5 Fan Exit P12.5 FEP
Le
g
end
1)
– Same Sensor used as for En
g
ine Control is also used for Trend Monitorin
g
;
– Special Sensor used for Engine Trend Monitoring
2) EPR / Engine Pressure Ratio =
Engine
Control
Remarks
Airflow Stations Measured Parameters Parameters used for
–
–
4.9 LP Turbine Exit
for EPR Calculation
2
)
–
–
3 HP Compressor Exit
2.5
LP Compressor Exit, (HP)
Compressor Inlet
–
2 Fan Inlet
for EPR Calculation
2
)
P4.9
P2

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Figure 5: Engine Stations
AERODYNAMIC STATIONS
1 2 12.5 2.5 3 4 4.5 4.9
STATION 2 2.5 12.5 3 4.5 4.9
PT PSIA 14.7 26.2 24 438.9 82.9 19.9
TT °C 15 74.1 64.6 540 791.4 496.6
TT °F 59 164.4 148.3 1003.9 1456.5 925.9
V2500-A1

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Power Plant V2500A
72-00 Engine Presentation
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Copyright by SRTechnics
Engine Main Bearings The 5 bearings are located in 3 bearing compartments. Front Bearing Compartment The front bearing compartment is located at the centre of the intermediate case,
and houses bearing No. 1, 2 & 3. Center Bearing Compartment The center bearing compartment is located in the diffuser/combustor case and
houses bearing No. 4 Rear Bearing Compartment The rear bearing compartment is located in the turbine exhaust case No.5 Bearings The Low Pressure or N1 rotor, is supported by three bearings:
• Bearing 1 (Single track thrust ball bearing).
• Bearing 2 (Single track roller bearing utilising "squeeze film" oil damping).
• Bearing 5 (Single track roller bearing utilising "squeeze film" oil damping).
The High Pressure or N2 rotor is supported by two bearings:
• Bearing 3 (thrust ball bearing mounted in an hydraulic damper which is cen-
tered by a series of rod springs ("Squirrel Cage")).
• Bearing 4 (Single track roller bearing).

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Figure 6: Engine Bearings & Compartments

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72-00 Engine Presentation
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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 (hy-
draulic) seal between the two shafts. This seal is supported by 8th stage air.
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 cou-
pling 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 stubs haft, this interacts with speed probes to pro-
vide LP shaft speed signals (N1) to th e 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 pressu re. This allows larger jet di ameters to facilitate flow tol-
erance 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 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.

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Figure 7: Front Bearing Compartment

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NO 4 Bearing Compartment The No 4 bearing compartment is situated in an inherently hostile, high tempera-
ture and pressure environment at the centre of the combustion section.
The bearing compartment is shielded from radiated heat by a heat shield and an
insulating supply of relatively cool air.
This supply of cooled 12th stage air (ca lled "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 co oler air from which the seal leakage is ob-
tained.
This results in an acceptable temperature of the air leaking into the bearing com-
partment.
Buffer air flow rates are controlled by restrictors at the outlet from the cooling pas-
sages.
The bearing compartment internal pressure level is determined by the area
of the variable scavenge valve. (called No 4 bearing scavenge valve and de-
scribed in the oil system). This valve ac ts as a variable restrictor in the com-
partment vent / scavenge line.
A drain hole is provided to indicate a possible leckage at the No 4 bearing
compartment. It is located in the exhaus t at 5 o clock position (aft looking for-
ward)
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.

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Figure 8: No.4 Bearing Compartment
P
COMPRESSOR
12 TH STAG E
AIR
HEATSHIELD
COOLING DUCT
FRONT WALL
ASSEMBLY
BEARING SUPPORT
INNER FLANGE
DIFFUSER CASE REAR
REAR WALL
FRONT SEAL
FRONT SEAL SEAT
No. 4 BEARING
REAR SEAL SEAT
REAR SEAL
RING LOCK AND NU
T

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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 0- ring and a thermally insulated heat
shield, both secured by the same twelve bol ts. Inside the LP shaft there is a small
disc type plug with an 0-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.
Figure 9: Bearing No.5 Compartment

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Figure 10: Rear Bearing Compartment
REAR THERMAL BLANKET COMPARTMENT
COVER
TEC
BLIND CAP
BLIND CAP
LPT SHAFT
STAGE 8 AIR
PACKING

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72-00 Engine Presentation
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Engine Modules Modular construction has the following advantages:
• Lower overall maintenance costs
• Maximum life achieved from each module
• Reduced turn-around time for engine repair
• Reduced spare engine holdings
• Ease of transportation and storage
• Rapid module change with minimum ground running
• Easy hot section inspection
• Vertical/horizontal build strip
• Split engine transportation
• Compressors/turbines independently balanced
The engine modules are:
31 the fan module,
32 the intermediate case module,
40 / 41 the high pressure compressor, & diffuser/combustor module,
45 the high pressure turbine,
50 the LP turbine
60 the accessory drive gearbox.
The module numbers refer to the ATA chapter reference for that module. Fan Module It consists of a single stage, wide-chord, shroudless fan and hub. Intercase Module It consists of the fan containment case, fan exit guide vanes (EGV), intermediate
case, booster, low spool stubshaft, the accessory gearbox towershaft drive as-
sembly, high spool stubshaft and the stat ion 2.5 bleed valve (BSBV). The booster
consists of inlet stators, rotor assembly, and outlet stators. The No. 1, 2 and 3
(front) bearing compartment is built into the module and contains the support bear-
ings for the low spool and high spool stubshafts.
High Pressure Compressor The HP compressor is a ten stage, axial flow module. It is comprised of the drum
rotor assembly, the front casing which houses the variable stator vanes and the
rear casing which contains the fixed stators and forms the bleed manifolds. Diffuser / Combustor Module The combustion section consists primarily of the diffuser case, annular two piece
combustor, with 20 fuel injector and 2 ignitors. The high compressor exit guide
vanes and the No. 4 bearing compartment are also part of the module.
The main features of the module include a close-coupled prediffuser and combus-
tor that provide low velocity shroud air to feed the combustor liners and to mini-
mize performance losses. High Pressure Turbine The high pressure turbine is a two stage turbine and drives the HP compressor
and the accessory gearbox. Active clearance control is used to control seal clear-
ances and to provide structural cooling. Low Pressure Turbine The low pressure turbine is a five stage module. Active clearance control is used
to control seal clearances and to provide structural cooling. Accessory Drive Gearbox The accessory drive gearbox provides shaft horse power to drive engine and air-
craft accessories. These include fuel, oil and hydraulic pressure pumps and elec-
trical power generators for the EEC and fo r the aircraft. The gearbox also includes
provision for a starter which is used to drive the N2 shaft for engine starting.

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Figure 11: Engine Modules

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Module 31 (Fan Module) Module 31 is the complete Fan assembly and comprises:
• 22 wide-cord, titanium shroudless hollow fan blades
• 22 annulus fillers
• the titanium fan disc
• the front and rear blade retaining rings
The blades are retained in the disc radially by the dovetail root.
Axial retention is provided by the front and rear blade retaining rings.
Blade removal / replacement is 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. 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. Ice protection is provided by a soft rubber
cone tip.
The nose cone retaining bolt flange is faired by a titanium fairing which is secured
by 6 bolts.
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 locate d by the front and rear blad e 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 fille rs and the aerofoils, there
is a rubber seal bonded to each side of each filler.
Fan Disc 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.

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Figure 12: LP Compressor (Fan)

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Inlet Cone The Glass-fibre cone smoothes the airflow into the fan. It is secured to the front
blade-retaining ring by 24 bolts. A Fairing is attached to the front blade-retaining
ring by 6 bolts.
Balance weights must not be placed at these 6 bolt locations on the fairing.
The Nose Cone is balanced during manufacture by applying weights to its
inside surface.
The nose cone is un-heated. A soft rubber cone tip provides ice protection. As ice
builds up on the tip, it becomes un-balan ced and flexes. This causes the ice to be
dislodged from the rubber tip and is then ingested by the fan before it has built up
to a significant mass. The Nose Cone retaining bolt flange is faired by a titanium
fairing which is secured by six bolts.
The arrangement is shown below.
Take care when removing the Nose Cone retaining bolts. Balance weights
may be fitted to some of the bolts. The position of these bolts with their re-
spective weights must be marked before removal, so as to ensure they are
refitted to the same positio n.A special tool is used to remove the Inlet Cone
to prevent it from damage as shown below.

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Figure 13: Inlet Cone Removal
,/#!4).' 0).
/&&
).,%4 #/.%
/&&
(/,%
,/#!4).'
&!)2).'
7!3(%2
42)- "!,!.#%
7%)'(4
"/,4 /&&
"/,4 /&&
&2/.4 ",!$% 2%4!).).'
2).'
&,!.'%
).,%4 #/.%
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/0%.).'
/&&
05,,%2 ,%6%2
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&).$ 4(% 3)8 2%#4!.'5,!2 /0%.).'3 ). 4(% 2%!2 &!#% /& 4(% ).,%4 #/.% &,!.'% ). 4/ 7()#( 4(% 4)0 /& 4(% ,%6%2 )3 054 3%% 6)%7 !

!0,,9 %15!, 02%3352% #!2%&5,,9 &/27!2$ !4 %!#( /0%.).' ). 452. 7)4( 4(% ,%6%2 4/ 4(% &,!.'% 4/ "2%!+ ).4%2&%2%.#% &)4
!
!

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Front Blade Retaining Ring The Assembly is shown below.
The Front Blade Retaining Ring is secured to the Fan Disk by a ring of 36 bolts. A
second (outer ring) passes through the reta ining ring and permits the individual se-
curing of the Annulus Fillers by 22 bolts.
Both these sets of bolts must be removed before attempting to remove the Front
Blade Retaining Ring.
After the removal of the 22 annulus filler securing bolts and all 36 retaining ring
bolts, it is possible to remove the front bl ade retaining ring by the use of 6 ‘pusher
bolts being inserted into 6 threaded holes designed specifically for this purpose.
The fan blades and annulus filler positions are not id entified. For this reason
it is important to identify and make a note of the original blade and annulus
filler positions prio r to their removal.
When the Nose Cone is fitted, it is possibl e to identify the posi tions of blades num-
bers 1,2 and 3 by noting that the front blade retaining ring has etched on it’s outer
edge these blade number positions. These numbers are marked in a counter-
clockwise direction when viewing the engine from the front.
Having established the original positions of the blades it is important to number the
blades and their corresponding annulus filler by using an approved marker pen

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Figure 14: Front Blade Retaining Ring

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Fan Blade Removal / Installation Removal The Nose cone is secured to the front blade retaining ring by 18 bolts.
Be careful when removing the nose cone retaining bolts.
Balance weights may be fitted to some of the bolts. The position of these weights
must be marked before removal to ensure they are refitted to the same position.
The blade retaining ring is secured to the fan disc by a ring of 36 bolts. A second
(outer) ring of bolts passes through the re taining ring and screws into each of the
22 annulus fillers. Both ring s of bolts must be removed before attempting to re-
move the front retaining ring.
After all the securing bolts (22 + 36) have been removed the retaining ring can be
removed by srewing pusher bolts into the 6 threaded holes provided for this pur-
pose.
Balance weights, if required are located on the retaining ring.
The fan blades and annulus filler positions are not identified. For this reason it is
important to identify the bl ade and annulus filler position, relative to the numbered
slots in the fan disc, before disassembly.
Remove the annulus fillers on either side of the blade to be removed.
The annulus fillers can be removed as follows:
• lift the front end of the annulus filler 3 to 4 inches.
• twist the annulus filler through about 60 deg counter - clockwise
• draw the annulus filler forward to clear the blades
The blade to be removed can then be pulled forward to clear the dovetail slot in
the fan disc. Installation After the new blade and the annulus fillers are fitted, The front blade retaining ring
can be fitted.
The front blade retaining ring can only be fitted in one position which is determined
by tree off - set locating dowells on the fan disc.
When the retaining ring is fitted to the fan disc the lettet T, etched on the retaining
ring, identifies No 1 fan blade position.
Fan blade Inspection / repair are described in the AMM 72-31-11 Page block
800.
The moment weight of the fan blade is written on the root surface
Figure 15: Fan Blade Profile
HONEYCOMB CORE
CONVEX
SKIN
CONCAVE SKIN

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Figure 16: Fan Blade Removal / Installation

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Annulus Fillers After removal of the Front Blade retaining ring the Annulus Fillers can be removed
as follows:
• lift the front end of the Annulus Filler 3 to 4 inches
• twist the Annulus Filler through about 60 degrees counter-clockwise
• draw the Annulus Filler forward to clear the blades
Remove the annulus fillers on either side of the blade to be removed. The blade
to be removed can than be pulled forward to clear the dovetail slot in the fan disc.
Examine the outer surface of the Annulus Filler for cracks, nicks, dents and
scores.
Limits in the AMM can be applied to assess the damage for accept or reject.
If the surface coating of the annulus fille r is damaged to the point of requiring a
repair the AMM has a procedure that allows this to be done.
AMM ref 72-31-11-300-010 gives comprehensive instructions as to the correct
procedure for repair
When re-fitting the Annulus Fillers, it is extremely important that correct
location of the Annulus Fillers into the Rear Retaining Ring is achieved.
If the Annulus Filler is not correctly in stalled, it is possi ble that when the
Front Retaining Ring is subsequently torque tightened in place onto the Fan Disk,
it may result in the deformation and di splacement of the Rear Retaining Ring.
This could cause it to come into contac t with the inlet housing of LP Compressor
Module Reposition of the Annulus Filler Seals During the installation of the Annulus Filler it is possible to cause the sealing strips
to be incorrectly seated.
If this were to be left uncorrected, it is possible that the Fan Blade would be dis-
placed slightly prevented from it’s normal radial operating position.
This in turn would cause the Fan Module to become un-balanced and vibration
levels for the engine could be exceeded.
The task referenced above documents the procedure to eliminate this.
The task requires a stiff plastic strip to be used to reposition the seals if they ‘
rolled’ as shown in the diagram below.
Make sure the plastic strip has a smooth surface and edges. If you use a
strip with a rough edge surface or ed ges, damage to the seal can occur.
Make sure that you do not break the plastic strip and leave pieces of it in the Fan.
Pieces of plastic can damage the rubber.

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Figure 17: Annulus Filler

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72-31-11 Fan Blade Repair Fan Blade Inspection / Repair Before any repair is carried out, reference must be made to the AMM Chapter 72-
31-11 Page Block 800.
Repair Damage on the Low Pressure Compressor (LPC) Fan Blades by Local Ma-
terial Removal
• YOU MUST USE SILICON CARBIDE TYPE ABRASIVE WHEELS, STONES
AND PAPERS TO DRESS, BLEND AND POLISH THIS COMPONENT.
• IF THE MATERIAL SHOWS A CHANGE IN COLOR, TO DARKER THAN A
LIGHT STRAW COLOR, THE COMPONENT IS TO BE REJECTED.
• DO NOT USE FORCE WITH MECHANICAL CUTTERS, OR THE MATERIAL
WILL BECOME TOO HOT.
• LP COMPRESSOR FAN BLADES MUST BE REPAIRED AS SOON AS DAM-
AGE OR WEAR IS MONITORED, TO GET BACK LP COMPRESSOR EFFI-
CIENCY AND EXTEND THE ROTOR BLADE LIFE.
• THE MAXIMUM NUMBER OF DRESSED BLADES FOR A GIVEN THE LP
COMPRESSOR FAN BLADES SET IS THE EQUIVALENT OF THREE
BLADES DRESSED TO THE MAXIMUM LIMIT. ALL THE REMAINING
BLADES MUST NOT BE DRESSED.
• THE MAXIMUN NUMBER OF DRESSED BLADES MUST BE OBEYED, TO
PREVENT A RISK OF ENGINE VIBRATION. Procedure This repair lets you scallop the leading edge, remove damage from the airfoil sur-
face and if damage is found in Zone AD, then you must blend parallel with the
leading edge, to remove any material above the repaired area by material remov-
al. A. Chemically Clean the Blades 1. Use alkali cleaner, alkani cleaner (Material No. V01-339) or alkani cleaner and
prepare the solution (Ref. AMM TASK 70-11-50-100-010).
2. Wash the repaired area with a cloth soacked in the solution.
3. Use a cloth soaked in clean cold water until the area is fully cleaned.
4. If necessary repeat steps (2) and (3).
5. Wipe the area with a clean dry cloth.
B. Do a Local Penetrant Crack Test on the Damaged Blades 1. Use fluorescent penetrant and do a penetrant inspection of the damaged area
(Ref. SPM 702305). C. Examine the Blade Airfoil 1. Examine the blade airfoil for crack indications. Use X10 binocular under ultra
violet light.
If a blade is cracked, reject it.
2. Examine the blade for damage (Ref. TASK 72-31-11-200-010).
If a blade is damaged, do step (4.D.) that follows. D. Remove Local Damage on the Leading Edge
(Ref. Fig. 804 / TASK 72-31-11-991-174)
1. Remove damage on the leading edge by removal of minimum material.
Continue to remove damage until all the damage is removed. Use portable
grinding equipment.
If damage is shown in Zone AD, you must blend the damage parallel with
the blade leading edge, to remove any material above the repaired area.
If you blend in Zone AD, you can only have one scallop in Zone AC, Zone
AA and Zone AB, can each have a scallop, independently of the repair of
Zones AD and AC.
1. Remove damage as necessary on the airfoil surface by the removal of mini-
mum material. Continue to remove damage until all the damage is removed.
The maximum depth to remove the damage must not be more than 0.015 in.
(0.38 mm). The diameter of the repaired area is to be 50 times the depth.
2. Make smooth the repaired area‘s. Make sure all the damaged marks are com-
pletely removed and the surface finish is made the same as the adjacent ma-
terial. Use waterproof abrasive paper, waterproof abrasive paper and / or
waterproof abrasive paper.
Polish the repaired area‘s, to remove scratches and make the surface finish the
same as the adjacent material. Use waterproof abrasive paper, waterproof
abrasive paper (and / or waterproof abrasive paper.

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Figure 18: Fan Blade Repair Limits
:/.% !! :/.% !" :/.% !$ :/.% !#
0%2-)44%$ 3#!,,/03 /. "/4( 4(% ,% !.$ 4% !2% ./4 0%2-)44%$ 4(% 4% )3 0%2-)44%$ ",%.$%$ 9/5 #!. (!6% 47/ 3#!,,/03 /. "/4( 4(% ,% !.$ 4%
&!. ",!$% 3#!,,/0).' ,)-)43
,%!$).' %$'%
IN MM
IN MM
IN MM
IN MM
IN MM IN MM
IN MM IN MM
IN MM
IN MM
./ -/2% 4(!. /.% ",%.$%$ !2%! /. 4(% ,% !.$ /.% ",%.$%$ !2%! /. 3#!,,/03 !2% !,,/7%$ /. 4(% 4% !.$ ,% )& :/.% !$ (!3 ./4 "%%. /.,9 /.% 3#!,,/0 /. 4(% ,%!$).' %$'% ,% /2 42!),).' %$'% 4% )3
:/.%
!$
:/.%
:/.%
:/.%
!# !" !!
42!),).' %$'% 4/ 7(%.
&!. ",!$% ",%.$ $%4!),3
-!8 $%04( !% 6!2)%3 7)4( :/.%
$2%33 0!2!,,%, 7)4( ",!$% ,%!$).' %$'% !.$ 2%-/6).' $!-!'% ). :/.% !$
",%.$ !,, $2%33).' 3-//4(,9 ).4/ !)2&/),
X
$%04(
!%
!%

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Figure 19: Fan Blade Repair Limits
$2%33).' ).#,5$).' ",%.$).' ./4 4/ %8#%%$ 4()3 "/5.$!29
-!8 IN MM 352&!#% $2%33).'
$
",%.$).' /& 352&!#% $2%33).'
IN MM
352&!#%
-%!352%$ /. ",!$%
&!. ",!$% 352&!#% $2%33).' &/2 02%3352%
&!#% !.$ ./. 02%3352% &!#%
IN MM
X$
0%2-)44%$ 02/&),%
0/,)3( ! ,)44,% /&& !,, !2/5.$ 4(% ./3% 4(% &).)3(%$ 02/&),% -534 ')6% ! ./3% $/7. #/.$)4)/. /2 "% ). ! &5,, 2!$)53
&5,, 2!$)53
02%3352%
#/.#!6% 352&!#%
#/.6%8 352&!#%
35#4)/.
./4 0%2-)44%$ 02/&),%
4(%3% &!5,43 #!. "% 3%%. 7)4( 4(% .!+%$ %9%
!6/)$ &,!43
!6/)$ ! ./3% 50 4%.$%.#9
$/ ./4 -!+% 4(% ./3% 5.$5,9 ",5.4
$/ ./4 -!+% 4(% ./3% 6%29 3(!20

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E. Examine the LP Compressor Fan Blades 1. Visually examine and measure the dimensions of the scallop on the leading
edge and the airfoil surface. Make sure the maximum depth of the repair on the
airfoil surfaces is not more than 0.015 in. (0.38 mm). Discard the blades, if they
are not in the limits specified. Use workshop inspection equipment. F. Do a Local Penetrant Crack Test on the Damaged Blades 1. Use fluorescent penetrant and do a penetrant inspection of the damaged area
(Ref. SPM 702305). G. Identify the Repair 1. A log book entry is necessary when you have completed this repair.
Write VRS1506 in the engine log book.
2. At the next shop visit make a mark VRS1506 adjacent to the part number. Use
vibro-engraving equipment.
Blades repaired to this scheme, must be swab etched and inspected as
specified in the (Ref. EM 72-31-11-300-025) (VRS1026) and glass bead
peened at the next shop visit, to the instructions specified in the (Ref. EM
72-31-11- 300-016) (VRS1724).
Figure 20: LP Compressor Fan Blade
"%&/2%
!&4%2
490)#!, %8!-0,% /& $!-!'% "%&/2% !.$ !&4%2 3#!,,/0).'

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Figure 21: Fan Blade leading and trailing edge limits
11.500in.
(292.10mm)
0.75in.
(19.05mm)
Bt
3.00in.
(76.20mm)
Br
1.50in.
(38.10mm)
TRAILING EDGE
AREA F
At
LEADING EDGE
2.00in.
(50.80mm)
Ct
ANNULUS
LINE
Ar
3.00in.
(76.20mm)
Cr
MAXIMUM SERVICEABLE LIMITS FOR SURFACE DAMAGE DEPTH
ON CONVEX AND CONCAVE SURFACES.
Ar
0.008in. (0.20mm)
0.008in. (0.20mm)
0.008in. (0.20mm)
Br
F
At
0.025in. (0.63mm)
0.025in. (0.63mm)
Bt
0.008in. (0.20mm)
Ct

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Repair of the Fan Disk Rear Ramp During the removal operation of a fan blade, it is possible to dislodge the rear ramp
from its location in the ‘dov e-tail’ slot in the fan disk.
Great care must be taken to inspect the f an disk and the security of the rear ramps,
as they play an important role in providing a firm fixing and support for the individ-
ual fan blades.
Should it be discovered that a rear ramp has become separated from the disk it
must be refitted/replaced and a full descri ption of the task can be found in the
AMM task reference 72-31-12-300-010. This is summarised as follows:
Remove the stage 1 fan blade from the stage 1 fan disk assembly
Clean the disk and rear ramp bonding surfaces:
• Hand abrade the disk and rear ramp bonding area, using a scotch brite pad
(material No. V05-126) or garnet paper (Material No. V05-017)
• Swab degrease the disk and rear ramp bonding areas, using a clean lint-free
cloth made moist with methyl ethyl keytone (material No. V01-076)
Mating surfaces of the component must be scrupulously clean and contact
surfaces must not be touched by hand or otherwise contaminated. Bonding
must be carried out immediately following surface preparation
Bond the rear ramp to the disk:
• Apply masking tape to the rear ramp. Using masking tape (Material No. V02-
019) Note! The masking tape is used in order to allow the engineer to hold and
place the rear ramp accurately in the dovetail slot. See diagram on next page.
• Apply the adhesive to the disk and rear ramp bond areas. Use toughened
acrylic adhesive with initiator (Material No. V08-114) Use a small spatula or
trowel to apply the adhesive. Note The four ‘pips’ on the rear ramp, are to en-
sure adequate thickness of adhesive is maintained between the mating surfac-
es. See diagram on next page.
• Fix the rear ramp to the fan disk and remove the masking tape from the rear
ramp.
• Use finger pressure to hold the rear ramp in position for three minutes.
• Cure the adhesive for one hour at room temperature between 21 deg. C. and
25 deg. C.
• Visually and dimensionally examine the bonded rear ramp.
• Install the stage 1 fan blade to the fan disk assembly.

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Figure 22: Fan Disk Rear Ramp
SECTION
A-A
FOR REAR RAMP
DIRECTION OF INSERTION
DISK
BB
REAR FACE OF REAR RAMP
REAR FACE OF DISK
0.020 (0,50)
0.000 (0,00)
REAR RAMP
MASKING TAPE
FOUR PIPS
GLUE LINE THICKNESS TO BE
CONTROLLED BY FOUR PIPS ON
REAR RAMP
SECTION
C-C
AB
FLAT BOTTOM
PORTION
VIEW ON
BOND REAR RAMP WHERE MARKED
D
ALL DIMENSIONS ARE IN IN. (MM)
0.006 0,15
0.004 0,10
(())

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TAP Transient Acoustic Propagation Test E.g. after ingestion of birds, foreign obje cts or slush a TAP test of the LP Compres-
sor fan blades needs to be carried out within a specific timeframe.
• Do a transient acoustic propagation test (Ref. AMM TASK 72-00-00-200-011)
within 10 flight hours/5 flight cycles, whichever is sooner. Do an Inspection of the Fan Blades 1. Apply a small quantity (approximately p ea sized) of ultrasonic couplant (Mate-
rial No. V06-148) to the lower convex ai rfoil adjacent to th e annulus filler line.
2. Attach the probe to the fan blade.
3. Press the ON switch.
• DO NOT HOLD THE FAN BLADE WHEN YOU READ THE VALUE. YOUR
HAND WILL ABSORB SOME OF THE SOUND PULSE WHICH CAUSES A
FASTER DECAY RATE.
• MAKE SURE THAT NO LEADING EDGE AND/OR TRAILING EDGE PRO-
TECTION IS INSTALLED WHEN YOU READ THE VALUE. THE PROTEC-
TION WILL ABSORB SOME OF THE SOUND PULSE WHICH CAUSES A
FASTENER DECAY RATE.
4. Press the EXE switch.
a) Press the EXE switch. The display will show the value or message in ap-
proximately four seconds.
If the display shows the message COUPLING FAILURE, apply more
ultrasonic couplant (Material No. V06-148) and do the inspection again.
5. If the TAP-test display value is more t han 800 dB/sec., reject the fan blade.
6. If TAP-test display value is more than 700 dB/sec. but less than 800 dB/sec.,
the fan blades can stay in use for further five flight cycles. Reject the fan blades
after five flight cycles.
Figure 23: TAP Test
ACOUSTIC
EMISSION PROBE
ANNULUS FILLER LINE
ACOUSTIC EMISSION PROBE
EXEC
MENU
ON
OFF

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Fan Trim Balance There are two methods available to balance the fan, the ‘one shot’ and ‘trial
weight’ the method. Both use data gained from the Engine Vibration Monitoring
System (EVMS).
The one shot method allows balancing of the fan with fewer engine ground runs
required and has proved itself effective in service use.
If necessary a Vibration survey (Test No 8) may be performed to obtain the vibra-
tion characteristics of the engine.
Note:
• If vibration exceeds limits during the survey ground run, slowly bring engine
speed to idle and shutdown.
• Angles are counter clockwise viewed from the front of the engine.
Data: (speed, amplitude and phase angle) may be collected on ground or during
cruise flight, collection in flight is eit her automatic or for selected speeds and on
the ground may be manually selected.
Best results are obtained from data in the 80-90% N1 speed range with 85% N1
being the best single speed point, for gr ound running an average of correction. One Shot Method The following procedure may be used to trim balance an engine fan whilst mount-
ed on the aircraft wing. The data collection will be via the aircraft EVMU system.
Data may be collected during a ground run or in cruise flight.
Definitions
• Speed (N1) expressed as a percentage 100% = 5650 rpm. Note! (1% N1 =
56.5 rpm)
• Amplitude (U) indicated vibration levels expressed in Mils (P-P) from the EVMU
system.
• Phase Angle (A) indicated angle in degrees from the EVMU system.
• Phase Lag (B) dynamic phase lag of the LP system between phase angle and
true position of unbalance.
Mass Coefficient (K) value by which the amplitude must be multiplied to give cor-
rection mass required or a given speed
Fan Trim Balance with the EVMU (One Shot Method) This procedure can be used for consecutive fan trim balances if necessary. If con-
secutive fan trim balances wit h this method do not give significant results, carryout
a fan trim balance with the ‘Trial Weight’ method.
This information is contained in the EVMU and by accessing the EVMU Engine
Unbalance menu, it is possible to estab lish the necessary adjustments required to
eliminate out of balance situations.
Note:
Prior to carrying out any ad justments, the engineer must first confirm the accuracy
of the current status regarding the configuration of weights (position and part
number) that are already installed and recorded in the system.
To accomplish this it is necessary to phys ically verify the position and part number
of the balance weights already installe d onto the front blade-retaining ring.
Figure 24: Moment Weights
7 BALANCE WEIGHT
8 NUT
9 BALANCE WEIGHT
PULLER BOLT
10 BOLT (36 OFF)
PUSHER BOLT
(22 OFF)
11 BOLT
5 BOLT
4 FAN DISK
6 FRONT BLADE
RETAINING RING
14 LOCATING HOLE
(22 OFF)
13 THREADED
HOLE (6 OFF)
12 LOCATING
HOLE (3 OFF)
2 LOCATING PIN
(22 OFF)
3 SHOULDER
HEADLESS PIN
(3 OFF)

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Figure 25: Moment Weights

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Figure 26: Trim Balance CFDS Procedure
<LAST LEG REPORT
<LAST LEG REPORT
<PREVIOUS LEG REPORTS
<SYSTEM REPORT TEST
POST FLIGHT REP PRINT>
CFDS
<AVIONICS STATUS
<AIR COND
<AFS
<COM
<FIRE PROT
<RETURN
SYSTEM REPORT / TEST
<ELEC
F/CTL>
FUEL>
ICE & RAIN>
L/G>
NAV>
INST>
<RETURN
<PNEU
<APU
SYSTEM REPORT / TEST
ENG>
TOILETS>
<RETURN
<EIU 1
<FADEC 1A
<FADEC 2A
<EVMU
SYSTEM REPORT / TEST
ENG
EIU 2>
FADEC 1B>
FADEC 2B>
<ACC RECONFIGURATION
<ENGINE UNBALANCE
<FREQUENCY ANALYSIS
<RETURN
EVMU
PRINT*
<RETURNPRINT*
<ENG 1 FLIGHT DATA
LOAD
GROUND ACQN
TRIM
ENG 2>
ENG 2>
ENG 2>
ENG 2>
<ENG 1
<ENG 1
<ENG 1
EVMU
ENGINE UNBALANCE
<RETURNPRINT*
<FLIGHT DATA
<GROUND DATA
<MANUAL INPUT
EVMU
ENG1 CURRENT VIB DATA
<RETURNPRINT*
<ONE SHOT TRIM BALANCE
<TRIAL WEIGHT TRIM BAL
EVMU
ENG1 TRIM BALANCE
METHOD SELECTION
N1
RPM
3041
4199
4524
5088
ACCLRM
MIL
0.2
NO ACQUISITION
0.5
0.5
0.6
A
DEG
+0
+230
+236
+189
160408
D/M
03/01
03/01
03/01
03/01
<RETURNPRINT*
EVMU
ENG1 FLIGHT DATA
CONT>

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Figure 27:
<RETURNPRINT*
<FLIGHT V2500-A1
<FLIGHT V2500-A5
<FLIGHT SPECIFIC
EVMU
ENG1 ONE SHOT TRIM BAL
INFLUENCE COEFF SELECT
<RETURNPRINT*
<36 BOLT FLANGE
<24 BOLT FLANGE
EVMU
ENG1 INST CURRENT WGHTS
BALANCE SOLUTION CONT>
<RETURNPRINT*
<33 / 5AXXXX
<35 / N-A
CONT>
34 / 5AXXXX>
36 / 5AXXXX>
EVMU
ENG1 INST CURRENT WGHTS
36 BOLT FLANGE
<RETURNPRINT*
<19 / 5AXXXX
<21 / N-A
CONT>
20 / 5AXXXX>
22 / 5AXXXX>
<23 / N-A24 / 5AXXXX>
EVMU
ENG1 INST CURRENT WGHTS
24 BOLT FLANGE
<RETURNPRINT*
<36 BOLT FLANGE
<24 BOLT FLANGE
EVMU
ENG1 INST CURRENT WGHTS
BALANCE SOLUTION CONT>
<RETURNPRINT*
<36 BOLT FLANGE
<24 BOLT FLANGE
EVMU
ENG1 INST CURRENT WGHTS
BALANCE SOLUTION CONT>
<RETURNPRINT*
36 BOLT FLANGE 24 BOLT
01 / 5A0103
02 / 5A0127
03 / REMOVE03 / 5A0127
04 / 5A012710 / 5A0107
EVMU
ENG1 WEIGHTS TO CHANGE
SOL XX XXQZIN / XXXDEG
CONFIG UPDATE>
<RETURNPRINT*
<01 / 5AXXXX
<05 / 5AXXXX
<07 / 5AXXXX
<03 / N-A
08 / 5AXXXX>
02 / 5AXXXX>
04 / 5AXXXX>
06 / 5AXXXX>
EVMU
ENG1 INST CURRENT WGHTS
36 BOLT FLANGE
If CONFIG UPDATE is pressed, the EVMU
automatically updates the current weights
configuration, relying on the assumption that the
proposed weights have been installed.
If RETURN is pressed, this will quit the menu
without storing the new bolt configuration.

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Module 32 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 a axial motion to
the valve ring via 2 power arms.

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Figure 28: Booster Stage Bleed Valve
LOWER POWER ARM
ACTUATING ROD
MID ARM
BLEED VALVE
UPPER POWER ARM
OPEN CLOSE
CLOSE
OPEN
AA
FAN FRAME
BLEED VALVE
AA

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72-00 Engine Presentation
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Figure 29: Fan Case Section
,0 345" 3(!&4
#526)# 4%%4( ).,%4 '5)$% 6!.%
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&2/.4 &!)2).'
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6!.%
!"2!$!",% 25""%2
25""%2 &),,%2
)..%2 2).'
,0# &2/.4 #!3%
!!
!!

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Power Plant V2500A
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Module 40 HP Compressor The HP compressor has 10 stages. It utilises variable inlet guid e vanes at the inlet
to stage 3 and variable stator vanes at st ages 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 cas-
ing 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.
On the V2500-A1 the Inlet Guide Vanes and stages 3, 4, 5 & 6 are variable.
Figure 30: HP Compressor

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Power Plant V2500A
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Figure 31: HP Compressor

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Combustion Section The combustion section includes the diffuser section, the combustion inner and
outer liners, and the No 4 bearing assembly. Diffuser Casing The diffuser section is a primary structural part of the combustion section.
The diffuser section has 20 mounting pads for the installation of the fuel spray noz-
zles. It also has two mounting pads for the two ignitor plugs. Combustion Liner The combustion liner is formed by the inner and outer liners.
The outer liner is located by five locating pins which pass through the diffuser cas-
ing.
The inner combustion liner is attached to the turbine nozzle guide vane assembly.
The inner and outer liners are manufactured from sheet metal with 100 separate
liner segments attached to the inner surf ace. The segments can be replaced inde-
pendently.
Figure 32: Combustor Cut

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Figure 33: Combustion Section

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Common Nozzle Assembly (CNA) General The mixed exhaust system collects two flows of air.
The first is the cold airflow, which is the fan bypass air.
The second is the hot airflow which comes from the engine core.
The mixed exhaust system is made up of the common nozzle exhaust collector
and the engine exhaust cone.
• The common exhaust collector admits the hot and cold gas outflows. These
gas outflows then go out to the atmosphere through the common nozzle.
• The nozzle forms a convergent duct which increases the speed of the mixed
gas to give forward thrust.
• The engine exhaust cone forms the inner contour of the common nozzle ex-
haust collector. It is made of a welded inco 625 honeycomb perforated panel
for sound attenuation, an attachment ring and a closure panel.
• Interface seals provide sealing between the exhaust collector, the thrust re-
verser and the pylon.
The cold airflow exhaust is part of the thrust reverser system described in 78-30-
00. When the thrust reverser operates, t he cold and hot outflows divide, and go in
different directions.

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Figure 34: Common Nozzle Assembly
42!.3,!4).' #/7, 2%!2 3%!,
/54%2 $5#4
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3%!,
34254
500%2 3500/24
,/7%2 3500/24
34254
!
!

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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 start-
er 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
• Individually replaceable drive units
• Magnetic chip detectors
• Main gearbox 2 magnetic chip detectors
• Angle gearbox 1 magnetic chip detector
• De-oiler
• Pneumatic starter
• Dedicated generator / alternator
• Hydraulic pump
• Oil Pressure pump
Rear Face
• Fuel pumps (and Fuel Metering Unit FMU)
• Oil scavenge pumps unit
• Integrated Drive Generator System (I.D.G.)

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Figure 35: Angle and Main Gearbox
3934%- )$'3 $2)6% 0!$
).4%'2!4%$ $2)6% '%.%2!4/2
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34!24%2 $2)6% 0!$

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Drive Seal The sealol seal
The picture below shows a typical SEALOL SEAL (carbon drive seal) installation
(Starter).
This type of seals are used on the drive pads on the gearbox.
consists of the following parts:
• A mating ring (glazed face) with four lugs engaging the four corresponding
slots in the gearshaft ball bearing.
• A cover, secured to the bearing housing with nuts, to ensure constant contact
between the glazed face and the static part of the seal.
The sealol seals are matched assemblies. If one of the components is damaged,
replace the complete seal!

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Figure 36: Drive Seals

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Borescoping General Hand Cranking A access to crank the HP compressor manually is provided at the front face of the
gearbox between the Starter and the deticated alternator (PMA).

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Figure 37: Manual Handcranking
2 CRANK COVER 3 WASHER
4 NUT
6 EXTERNAL GEARBOX MODULE
5 STARTER IDLER GEAR
1 PACKING
5 NUT (2 OFF)
(2 OFF)
4 WASHER
3 STARTER IDLER GEAR
2 ADAPTOR
MODULE
GEARBOX
1 EXTERNAL

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Borescope Inspection of the HP Comp. Borescope ports are provided to give access for visual inspection of the compres-
sor and the turbine. For further information and limits refer to AMM 72-00-00.
Inspection/Check Procedure
• Install the tool to turn the HP system.
• Prepare the borescope equipment for use as given in the makers instructions.
• Carefully put the borescope probe into th e access port of the stage of the com-
pressor you want to examine.
Use an 8mm probe for ports X, A, B and a 5.5mm probe for ports C, D, E, F
& G and a flexible borescope for inspection of the heatshield assemblies.
• Whilst turning the HP system, examine each blade in turn for:
– Nicks & Tears
– Cracks
–Dents
– Tip damage & discolouration
Blade numbers & dimensions are shown for each stage.
• Examples of blade damage limits are in AMM 72-00-00
• On completion of the inspection remove the borescope probe from the engine
and refit the access port covers as described on the next page.
• Remove the tool used to turn the HP system & return the engine to normal.
Figure 38: Borescope Inspection Equipment
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!.',% +./"
&/2 50 !.$ $/7.
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&/2 2)'(4 !.$ ,%&4
!.',% +./"

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Borescope Access Note 1:
IAE recommends that only the stage 3 & 12 HP compressor blades are examined
with the engine on-wing.
Note 2:
Access port D should not be used on engines that are pre SBE72-0033 as dam-
age can be caused to the borescope equipment.
• Remove the required borescope access part covers X, A, B, C, D, E, F, G, by
removing the attaching bolts. The diagram below shows which stage are ac-
cessed through each port.
• Remove the old jointing compound from around the access ports and access
port covers using a non-metallic scraper and a lint free cloth made moist with
cleaning fluid.
• Prior to installation of the borescope access port covers it Is necessary to apply
jointing compound. The procedure to be taken is:
Access ports X, A, B & C
• Apply a thin layer of jointing compound to the mating faces using a stiff bristle
brush. Do not apply within 0.12 to 0.16in (3 to 4mm) of access port.
• Wait 10 minutes, install access port cover & attach with bolts. Torque load to
between 85 - 105 lbf in.
• Re-torque again to same figures after 2 minutes then remove excess jointing
compound.
Access ports D, E, F & G.
• Do not require jointing compound.
Figure 39: Borescope Access Booster

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Figure 40: HP Compressor Borescope Access

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73 Engine Fuel and Control - V2500A

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73-00 Fuel System Presentation General The fuel system enables delivery of a fuel flow corresponding to the power re-
quired and compatible with engine limits.
The system consists of:
• the two stage fuel pump with low pressure & high pressure elements,
• the engine fuel cooled oil cooler (FCOC),
• the fuel filter
• the fuel diverter and return to tank valve.
• the integrated drive generator (IDG) fuel cooled oil cooler (FCOC),
• the fuel metering unit (FMU),
• the fuel distribution valve,
• the fuel flow transmitter,
• 20 fuel nozzles, Description and Operation Distribution The fuel supplied from aircraft tanks flows through a centrifugal pump (LP stage)
then through the Fuel Cooled Oil Cooler and then through a filter and a gear pump
(HP stage).
The fuel from the HP pump is delivered to the Fuel Metering Unit (FMU) which con-
trols the fuel flow supplied to the fuel no zzles (through the fuel flow meter and the
fuel distribution valve).
The FMU also provides hydraulic pressure to all hydraulic system external actua-
tors. These include the Booster Stage Bleed Valve actuators, Stator Vane Actua-
tor, ACOC air modulating valve and HPT/LPT Active Clearance Control valve. Low
pressure return fuel from the actuators is routed back into the fuel diverter valve.
The fuel diverter and return to tank valv e enables the selection of four basic con-
figurations between which the flow paths of the fuel in the engine are varied to
maintain the critical IDG o il, engine oil and fuel temp eratures within specified lim-
its. The transfer between configurations is determined by a software logic con-
tained in the EEC.
Controlling The Fuel Authority Digital Electronic Co ntrol (FADEC) system provides full range
control of the engine to achieve steady state and transient performance when op-
erated in combination with aircraft s ubsystems. The FADEC is a dual channel
EEC with crosstalk and failure detection capability. In case of specific failure de-
tection, the FADEC switches from one channel to the other.
The FADEC System operates compatibly with applicable aircraft systems to per-
form the following:
• Control of fuel flow, stator vanes and bl eeds to automatically maintain forward
and reverse thrust settings and to prov ide satisfactory transient response.
• Protect the powerplant from exceeding limits for N1, N2, maximum allowable
thrust, and burner pressure.
• Control of the HPT 10th stage cooling air, and low and high turbine active clear-
ance control systems.
• Control of fuel, engine and IDG oil temperature.
• Control of the thrust reverser.
• Automatic sequencing of start system components.
• Extensive diagnostic and maintenance capability.

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Figure 1: Fuel System Schematic

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73-10 Fuel Distribution Components 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) A filter 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.
b) 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 pres sure switch is de - energized which causes
the caution to go off.
The differential pressure switch si gnal is fed directly to the SDAC. Fuel Temperature Thermocouple (refer to 73-20 Heat Management System)
The measured temperature is transmitted to the EEC (Electronic Engine Control)
and used for the Heat Management System.. Fuel Diverter & Return Valve General The fuel diverter and return valve (FD & RV) is a primary unit in the heat manage-
ment 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 act-
ing as a fuel cooler.

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Figure 2: Fuel Filter Diff. Press. Switch/FCOC Fuel Temp. Thermocouple

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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 dis-
charge ports. The ten discharge ports ar e 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 noz-
zle.
The two fuel manifolds which remain full help supply fuel for the next engine start.
Figure 3: FDV Location

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Figure 4: Fuel Distribution Valve

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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 meteri ng 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 t he 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 co mbustion 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
high-velocity air as it enters the comb ustion chamber. The fuel nozzle forms the
atomized mixture of fuel and air into the correct pattern for satisfactory combus-
tion.
The design of the fuel nozzle results in fast vaporization of the fuel through the full
range of operation. This results in decreas ed emissions, high combustion efficien-
cy, and good start quality.
The high-velocity flow of fuel prevents formation of coke on areas where fuel
touches metal. Heatshields installed also prevent formation of coke.
Figure 5: Fuel Nozzle
(%!43()%,$
!)2 &,/7
&5%, &,/7
)..%2 (%!43()%,$
,!34 #(!.#% &),4%2

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Figure 6: Fuel Distribution Tubes

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Fuel Pump General The LP / HP fuel pumps are housed in a si ngle pump unit which is driven by a com-
mon 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 hydrome-
chanical components and for combustion in the burner.
The unit consists of a LP centrifugal bo ost stage which feeds an HP single stage,
two gear pump.
The housing has provision for mounting the fuel metering unit (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 stag e. 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 uni t, 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
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.
There are no mechanical inputs to, or outputs from the FMU.

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Figure 7: Fuel Metering Unit

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Fuel Metering Unit Fuel metering is achieved by the Fuel Metering Valve and the Pressure Drop Reg-
ulator and Spill Valve, which act to gether in the following sequence:
Signals from the EEC cause the torque motor to change position, which directs
fuel servo pressure to re-position the Fuel Metering Valve.
This changes the size of the metering or ifice through which the fuel passes which
in turn changes the pressure drop across the metering valve.
The change in the pressure drop is sensed by the Pressure Drop Regulator which
will re-position the spill valve and so incr ease/decrease the fuel flow through the
fuel metering valve until the pressure drop is restored to its datum value.
The increase/decrease in fuel flow causes the engine to accelerate/decelerate un-
til the actual EPR is that demanded by the EEC signal.
Movement of the Fuel Metering Valve is transmitted through a rack and pinion
mechanism to drive a dual output position resolver. The resolver output is fed back
to the EEC.
The EEC automatically corrects changes in fuel density. Bi-metallic washers locat-
ed in the pressure drop governor and spill valve assembly provide automatic com-
pensation for changes in fuel temperature.
The three main functions of the FMU are:
• metering the fuel supplies to the fuel spray nozzles.
• overspeed protection for both the LP (N1) and HP (N2) rotors.
• isolation of fuel supplies for starting/ stopping the engine.
These three functions are carried out by three valves arranged in series, as
shown:
• the Fuel Metering Valve
• the Overspeed Valve
• the Pressure Raising and Shut Off Valve.
The position of each valve is monitored and positional information is transmitted
back to the EEC.
This ensures that the EEC always knows that the valves are in the commanded
position.
FAIL SAFE POSITION OF THE METERING VALVE TORQUE MOTOR:
" MINIMUM FUEL FLOW CONDITION "
Overspeed Valve Operation The overspeed valve is spring loaded to the closed position, it is opened by in-
creasing 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 asignal 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 ar ound the overspeed valve to prevent engine
flame out.
The overspeed valve is hydraulically latche d in the closed position, thus prevent-
ing 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.
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" Pressure Raising and Shut off Valve 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 se-
quences if the sequence has to be stopped for any reason.
The EEC’s ability to close the shut of f 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 "

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Figure 8: Fuel Metering Unit Schematic

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Figure 9: FMU - Engine Shut Down

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Figure 10: FMU - Engine Running Figure 11: FMU - Engine Overspeed

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The engine fuel supply system has two fuel shut off valves.
• one PRSOV in the FMU
• One LP - fuel shut off valve on the front wing spar. Low Pressure Fuel Shut Off Valve 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 th is configuration lets 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 RIB 8 and RIB 9).
Each LP valve has an actuator 9QG (10QG). 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 sourc-
es:
• the 28VDC BATT BUS supplies the motor 1
• the 28VDC BUS 2 supplies the motor 2.
If damage occurs to the electrical circuit, it is necessa ry to make sure that the
valve can still operate. Thus the electric al 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 - Aquisition C oncentrators
(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
• a valve spindle
• a mounting flange.
The LP fuel - valve actuator has two electr ical motors which drive the same differ-
ential - 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 t he other motor does not op-
erate.
The actuator drive shaft has a see/feel indicator where it goes through the actuator
body. The see/feel indicator gives an indi cation of the valve position without re-
moval of the fuel LP fuel valve.

73-00-17
Power Plant V2500A
73-00 Fuel System Presentation
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 12: LP Fuel Shut-Off Valve

73-00-18
Training Manual
A319/A320/A321
Power Plant V2500A
73-00 Fuel System Presentation
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
HP & LP Fuel SOV Control The HP fuel shut off valve control is fully electrical.
It is performed from the engine panel in the cockpit as follows:
Opening of the HP fuel PRSOV:
it is controlled by the EEC: the EEC receives the commands from the MASTER
control switch and ignition selector switch.
Closure of the HP fuel PRSOV:
it is controlled directly from the M ASTER control switch in OFF position PRSOV Fuel Shut Off Control The FADEC control system contains a fuel shut - off in the FMU, which acts
through a 2 position torque motor to close the pressurizing valve:
The fuel shut - off is direct-hardwired to the MASTER control switch.
This tourque motor operated PRSOV is powered by the 28VDC.
• Loss of power supply does not lead to change the selected HP fuel shutoff
valve position.
• The cockpit command " OFF " has priority over the EEC command. LP Fuel Shutoff Valve Control The LP fuel shut-off system has two independ ent electrical control circuits for each
LP fuel - valve. They connect through a control relay to these related switches:
• the ENG MASTER switch
• the FIRE PUSH switch.
When the No. 1 ENG MASTER switch is set to ON, it disconnects a 28VDC supply
from the relay 11QG (HP FUEL SOV SOL P / B SW). The relay 11QG de - ener-
gizes and connects a 28VDC supply (through the ENG 1 FIRE PUSH switch) to
the " open " side of the LP fuel - valve actuator.
The actuator then opens the LP fuel - valve.
When the No. 1 ENG MASTER switch is set to OFF, it connects a 28VDC supply
to the relay 11QG. The relay energizes and connects a 28VDC supply (through
the ENG 1 FIRE PUSH switch) to the " close " side LP fuel - valve actuator. The
actuator then closes the LP fuel - valve.
If the ENG 1 FIRE PUSH switch is operated:
• it disconnects the 28VDC supply to the " open " side of the LP fuel - valve ac-
tuator
• it connects a 28VDC supply to the " close " side of the LP fuel valve actua tor
the LP fuel - valve moves to the closed position.
The LP fuel - valve opens (closes) when the ENG MASTER switch is set to
ON (OFF). But the operation of the engine FIRE PUSH switch always over-
rides an ON selection and closes the valve.
It is also commanded open via the relay 11QG when the C / B of the HP Fuel
SOV is pulled, (Relay 11QG (12QG) deenergized).

73-00-19
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73-00 Fuel System Presentation
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Figure 13: HP and LP Fuel Shutoff Valve (SOV)
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73-00-20
Training Manual
A319/A320/A321
Power Plant V2500A
73-00 Fuel System Presentation
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
73-20 Heat Management System Presentation General Heating and cooling of fuel, engine oil and IDG oil is accomplished by the Fuel
Cooled Oil Cooler (FCOC), the Air Cooled Oil Cooler (ACOC) and the IDG cooler
under the management of the EEC.
FUEL TEMPERATURE:
The fuel temperature is measured at the exit of the filter.
OIL TEMPERATURTE:
The engine oil temperature is measured upstream of the ACOC.
The IDG oil temperature is measured at IDG oil cooler exit.
The system is designed to provide adequate cooling, to maintain the critical oil and
fuel temperatures within specified limits, whilst minimisi ng the requirement for fan
air offtake.
Three sources of cooling are available:
• the LP fuel passing to the engine fuel system
• the LP fuel which is returned to the aircraft fuel tanks
•fan air
There are four basic configurations between which the flow paths of fuel in the en-
gine L.P. fuel system are varied. Within each configuration the cooling capacity
may be varied by control valves which form the Fuel Diverter and Back to Tank
Valve.
The transfer between modes of operation is determined by software logic con-
tained in the EEC. The logic is generated around the limiting temperatures of the
fuel and oil within the system together wi th the signal from the aircraft which per-
mits/inhibits fuel spill to aircraft tanks. Operation The measured temperature is transmitted to the EEC (Electronic Engine Control).
In response to the measured temperature, the EEC sends the signal to the fuel
diverter valve. The fuel diverter valve is used to reduce too high fuel temperature.
The excess of high pressure fuel flow from the FMU (Fuel Metering Unit) and re-
turn fuel from control actuator are plum bed to the diverter valve which normally
turns the flow to the FCOC exit.
Fuel Temp. Thermocouple The Fuel Temperature is measured by the thermocouple at the fuel exit of the
FCOC (Fuel Cooled Oil Cooler).
The thermocouple is composed of stainless steel sheathed sensing portion, stain-
less steel installing flange with seal spigot and electrical connector.
The control of fuel temperature is done by the fuel diverter valve which is installed
upstream of the FCOC. IDG Oil Cooler Temp. Thermocouple IDG Fuel Cooled Oil Cooler oil temperatur e is measured at the IDG Oil Cooler Exit
by a thermocouple.
The termocouple gives an electrical output in relation to the temperature of the oil
in the fuel cooled IDG oil cooler.
This temperature information is send to the EEC and is used for the heat manage-
ment system. ACOC Oil Temp. Thermocouple The oil temperature is measured at the ACOC inlet by a thermocouple. The ther-
mocouple is composed of stainless steel sheathed sensing portion, stainless steel
installing flange with seal sp igot and electrical connector.
The temperature is transmitted to the EEC (Electronic Engine Control). In re-
sponse to the measured temperature, the EEC sends the signal to the modulating
air valve. ACOC Modulating Air Valve The modulating air valve regulates air flow to the ACOC. Oil heated by the engine
passes through the ACOC and then to the FCOC. The air valve is modulated by
the EEC to maintain both oil and fuel temperatures within acceptable minimum
and maximum limits. Minimum oil temperature limits are used such that the oil may
be used to prevent fuel icing with the use of FCOC. Maximum limits have been
established to avoid breakdown of engine oil and to avoid excessively high fuel
temperatures.

73-00-21
Power Plant V2500A
73-00 Fuel System Presentation
Training Manual
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for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 14: HMS Main System Components

73-00-22
Training Manual
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Power Plant V2500A
73-00 Fuel System Presentation
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Copyright by SRTechnics
Fuel Diverter & Return Valve General The FDRV configuration allows four modes of operation according to electrical sig-
nals from the EEC (based on fuel and oil temperature measurements transmitted
by thermocouples). Description The fuel diverter and return valve is installed on the FCOC.
The FDV is a two - position selector valve which has two pistons in a sleeve.
The two pistons are mechanically connected and make two valve areas which are
referred to as valve A and valve B. The FRV has a main valve and a pushing piston
in a sleeve. This main valve is a half - area piston - type valve which moves valve
to change the metering port area. The main valve has two valve functions that are
referred to as valve C and valve D.
The EEC gives the electrical signal to the FDRV to change the position of the
valves. The FDRV gives a feedback signal to the EEC to transmit the position of
valves in the unit. The fuel flow chan ges with the position of the valves.
Thus, the fuel flow can be controlled through the FDRV and the EEC. Fuel Return Valve The EEC operates the dual-wound torque motor to control the servo pressure.
This servo fuel pushes the main valve.
The pressure balance between two sides of the main valve (Valves C and D) gives
the direction and the speed of the valve movement.
Then the valve changes the direction of the fuel flow and controls the metering port
area.
FAIL SAFE POSITION:
" FRV CLOSED, NO RETURN TO TANK (MODE 3 or 5) Fuel Diverter Valve The EEC energizes the solenoid valve to open the servo fuel flow.
The switch assemblies transmit the EEC the valve position when the solenoid is
de - energized.
FAIL SAFE POSITION:
" FDV SOLENOID DE - ENERGIZED " (MODE 4 or 5)
Return to Tank Modes
HMS Mode 1 (Normal Mode) This is the normal mode and is shown below. In this mode all the heat from the
engine oil system and the IDG oil system is absorbed by the LP fuel flows. Some
of the fuel is returned to the aircraft tan ks where the heat is absorbed or dissipated
within the tank.
This mode is maintained if the following conditions are satisfied:
• Engine not at high power setting (Take Off and early part of climb (not below
25,000ft).
• Cooling spill fuel temperature less than 100 deg C.
• Fuel temperature at pump inlet less than 54 deg C. HMS Mode 4 Mode 4 is the mode adopted when the burned fuel flow is low.
For example;
• Low engine speeds.
• High HP fuel pump inlet temperature.
In this mode the fuel/oil heat exchang er is operating as a fuel cooler.
The excessive heat is passed to the engine oil, the ACOC extracts the heat from
the oil that has been heated up by the hot fuel.
The ACOC modulating valve is fully open.

73-00-23
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A319/A320/A321
for training purposes only Sep08/Technical Training
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Figure 15: Return to Tank Mode 1 Figure 16: Return to Tank Mode 4

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No Return to Tank Modes 3 and 5
HMS Mode 3 Mode 3 shown below is the mode that is adopted when the requirements for fuel
spill back to tank can no longer be satisfied i.e.
• Engine at high power setting (below 25,000ft).
• Spill fuel temperature above limits (100 deg C).
• Tank fuel temperature above limits (54 deg C).
In this condition the burned fuel absorbs all the heat from the engine and I.D.G. oil
systems.
If however, the fuel flow is too low to provide adequate cooling the engine oil will
be pre-cooled in the air/oil heat exchanger, by a modulated air flow, before pass-
ing to the fuel/oil heat exchanger.
This is the preferred mode of operation, when return to tank is not allowed. HMS Mode 5 Mode selected when system condition demand as in mode 3 but this is not permit-
ted because IDG oil temperature is excessive or return to tank is not permissible
due to the high return fuel temperat ure. The ACOC valve is fully open.
This mode is adopted if the conditions exist.
In case the oil temperature cannot be kept within the limits the FADEC sys-
tem will increase the engine speed (FAIL SAFE MODE OF OPERATION.
.

73-00-25
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Figure 17: NO Return to Tank Mode 3 Figure 18: NO Return to Tank Mode 5

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73-00 Fuel System Presentation
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Copyright by SRTechnics
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 Bolted to the outlet face of the air/oil heat exchanger.
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.

73-00-27
Power Plant V2500A
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Figure 19: Air Modulating Valve

73-00-28
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73-00 Fuel System Presentation
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Copyright by SRTechnics
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 an d outlet ports. One set of ports is used
for the flow of the fuel to or from the fu el 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.
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 IDG Oil Cooler Temp. Thermocouple (refer to 73-20 Heat Management system)
This temperature information is send to the EEC and is used for the heat
management system.

73-00-29
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Figure 20: IDG FCOC Oil Cooler

73-00-30
Training Manual
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73-00 Fuel System Presentation
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Indicating 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 filter condition through a dedicated hardwired pressure switch.
The fuel system is monitored by:
• 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 asso-
ciated 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 trans-
mits the information to the ECAM system for display. Fuel Filter Clogging Indication
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:
• 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 pressu re switch is de-energized which causes
the caution to go off.
The differential pressure switch signal is fed directly to the SDAC through the
hardware.

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for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 21: Fuel System Indication
ENG 1(2) FUEL FILTER CLOG
EPR
EGT
N1
N2
C
%
%
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6
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6
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73-00-32
Training Manual
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Power Plant V2500A
73-00 Fuel System Presentation
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
FADEC Presentation FADEC = Full Authority Digital Engine Control
General The Full Authority Digital Engi ne Control system consists of an Electronic Engine
Control plus a Fuel Metering Unit, sensors and peripheral components. Electronic Engine Control The EEC consists of two channels (A and B) with crosstalk. Each channel can
control the various components of the engine systems.
They are permanently operational. one channel is in command while the other is
in standby. In case of failure of the operational channel, the system automatically
switches to the other one.
The channel selection strategy is based on channel health criteria. The com-
mand channel alternates each engine start. Interfaces The EEC receives air data parameters from the Air Data Inertional Reference Sys-
tem (ADIRS), and operational commands from the Engine Interface Unit (EIU).
It also provides the data outputs nescessary for the Flight Management and Guid-
ance Computers (FMGCs), and the fault message to the EIU for aircraft mainte-
nance data system.
Each EEC channel directly receives the Thrust Lever Angle (TLA).
The EEC transmits the thrust parameters and TLA to the FMGCs for the autothrust
function. Sensors Various sensors are provided for engine control and monitoring.
Pressure sensors and thermocouples are provided at the aerodynamic stations.
The primary parameters are Engine Pressure ratio (EPR = P4.9/P2), N1 and N2
speeds, Exhaust Gas Temperature (EGT) and metered Fuel Fuel Flow (FF). Fuel Metering Unit (FMU) In the FMU, three torque motors are activated by the EEC. These provide the cor-
rect fuel flow, overspeed prot ection and Engine Shut Down.
In case of an overspeed, an incorporated valve reduces the fuel flow.
The fuel Pressure Raising Shut Off Valve is controlled by the EEC through the
FMU, but it is closed directly from the corresponding ENG MASTER lever when
set to OFF.
The functions of the FADEC are also reset when the ENG MASTER lever is
set to OFF.
Compressor Airflow and Turbine Clearance Control The EEC controls the compressor airflow and the turbine clearance through sep-
arated sub systems.
It also monitors the engine oil cooling through an air/oil heat exchanger servo
valve.
Compressor airflow control:
• Booster Stage Bleed Valves (BSBV).
• Variable Stator Vanes (VSV).
• 7th and 10th stage handling bleed valves.
Turbine clearance control:
• HP and LP Turbine Active Clearance Control (ACC) valves.
• 10th stage make-up air valve.
Engine oil cooling:
• Air Cooled Oil Cooler (ACOC)servo valve.

73-00-33
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Figure 22: FADEC Presentation IAE V2500

73-00-34
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Power Plant V2500A
73-00 Fuel System Presentation
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Copyright by SRTechnics
Fuel Metering Unit The fuel metering unit (FMU) provides fuel flow control for all operating conditions.
Variable fuel metering is provided by the FMU through EEC commands by a
torque motor controlled servo drive. Position resolvers provide feedback to the
EEC. The FMU has provision to route excess fuel above engine requirements to
the fuel diverter valve through the bypass loop. Thrust Reverser Hydraulic Control Unit The EEC controls the thrust reverser operation through a Hydraulic Control Unit
(HCU)
Each EEC channel will energize the solenoids of an isolation valve and a direc-
tional valve included in the HCU to prov ide deployment and stowage of the thrust
reverser translating sleeves. Start and Ignition Control Each channel can control the starter valve operation, the fuel Pressure Raising
Shut - Off Valve opening and the ignition during the engine start sequence. Fuel Diverter and Return Valve The EEC manages the thermal exchange between the engine oil, IDG oil and en-
gine fuel system by means of a Fuel Diverter and Return Valve.
Part of the engine fuel can be recirculated to the aircraft tanks by means of a return
valve included in the fuel diverter valve module.
The EEC controls the operation of the Fuel Diverter and Return Valve according
to the engine fuel temperature (T FUEL) and the IDG oil temperature and the en-
gine oil temperature (T OIL). Engine Parameter Transmission for Cockpit Display The FADEC provides the necessary engine parameters for cockpit display
through the ARINC 429 buses output. Engine Condition Parameter Transmission Engine Condition monitoring is provided by the ability of the FADEC to transmit
the engine parameters through the ARINC 429 bus output.
The basic engine parameters available are:
• WF, N1, N2, P5, PB, Pamb T4.9 (EGT), P2, T2, P3 and T3.
• VSV, BSBV, 7th and 10th stage bleed commanded positions HPT/LPT ACC,
HPT cooling, WF valve or actuator position
• status and maintenance words, engine serial number and position.
In order to perform a better analysis of engine condition, some additional param-
eters are optionally available. These are P12.5, P2.5 and T2.
FADEC System Maintenance Fault Detection The FADEC maintenance is facilitated by internal exte nsive Built in Test Equip-
ment (BITE) providing efficient fault detection.
The results of this fault detection are contained in status and maintenance words
according to ARINC 429 specification and are available on the output data bus. Non Volatile Memory In flight fault data is stored in FADE C non volatile memory and, when requested,
is available on an aircraft centralized maintenance display unit. Communication with CFDS Ground test of electrical and electronic parts is possible from cockpit, with engines
not running, through the CFDS.
The FADEC provides engine control system self testing to detect problems at LRU
level.
FADEC is such that no engine ground run for trim purposes is necessary after
component replacement.

73-00-35
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Figure 23: FADEC Presentation V2500

73-00-36
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Power Plant V2500A
73-00 Fuel System Presentation
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Copyright by SRTechnics
FADEC Functions The FADEC system operates compatibly with applicable aircraft systems to per-
form the following functions:
1 GAS generator control for steady state and transient engine operation
within safe limits
• Fuel flow control
• Acceleration and deceleration schedules
• Variable Stator Vane (VSV) and Booster Stage Bleed Valve (BSBV) schedules
• Turbine clearance control (HP / LP)
• 10th stage cooling air control
• Idle setting.
2 Engine limits protection
• Engine overspeed protection in terms of fan speed and core speed to prevent
engine running over certified red lines
• Engine turbine outlet gas temperature monitoring. (EGT)
3 Power management
• Automatic engine thrust rating control
• Thrust parameter limit computation
• manual power management through constant ratings versus throttle lever re-
lationship
– take-off / go-around at full forward throttle lever position
– flex take-off at constant intermedia te position whatever the derating is
– other ratings (max continuous, max climb, idle, max reverse) at associated
throttle lever detent points.
• Automatic power management through direct engine power adjustment to the
autothrust system demand.
4 Automatic engine start sequencing
• Control of starter air valve ON / OFF
• Control of HP fuel valve (ON / OFF on ground, ON in flight)
• Control of fuel schedule
• Control of ignition (ON / OFF)
• EPR, N1, N2, WF, EGT monitoring
• Abort / Recycle capability on ground.
5 Thrust reverser control
• Control of thrust reverser actuation (deploying and stowing)
• Control of engine power during reverser operation.
• Engine idle setting during reverser transient
• Control of maximum reverse power at full rearward throttle lever position.
• Restow command in case of non commanded deployment.
• Redeploy command in case of non commanded stowage.
6 Engine parameters transmission for cockpit indication
• Primary engine parameters
• Starting system status
• Thrust reverser system status
• FADEC system status.
7 Engine condition monitoring parameters transmission.
8 Detection, isolation, accommodation and memorization of its internal sys-
tem failures.
9 Fuel return & diverter valve control
FADEC controls the ON / OFF return to the aircraft tank in relationship with:
• Engine oil, IDG oil and fuel temperatures
• Aircraft fuel system configuration
• Flight phases.

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Figure 24: FADEC Architecture

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Engine Control Pushbuttons and Switches Engine Mode Selector Position CRANK:
• selects FADEC power.
• allows dry and wet motoring (ignition is not availiable).
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 • controls the start valve (when the mode selector is in IGNITION / START or
CRANK position). FADEC GND PWR P/B Position ON:
• selects FADEC power N1 Mode P/B Position ON:
• switches EEC from EPR Mode to N1 Mode

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Figure 25: Engine Control P/B‘s and Switches
ON
OFF
OOOO
NN
OOOOOO
FFFF
FFF
MODE
NORM
2
ENG
MASTER 1 MASTER 2
CRANK
IGN
START
ON
OFF
ENG
2
ENG
1
FIRE
FIRE F
AUAU
LT
FIRE FIRE F
AUAU
LT
1
ON ON ON
ON
ENG ENG
MAN STARTN1 MODE
12 12
ON ON
ENG
FADEC GND PWR
12
C MAINTENANCE PANEL 50VU
B OVERHEAD PANEL 22VU
A CENTRAL PEDESTAL115VU

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Figure 26: Engine Circuit Breakers

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Figure 27: Engine Circuit Breakers

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Failures and Redundancy Improved reliability is achieved by utilising dual sensors dual feedback.
• Dual sensors are used to supply all EEC inputs exept pressures, (single pres-
sure transducers within the EEC provide signals to each channel-A and B).
• The EEC uses indentical 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 in put 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 re quired 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
whitch 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. 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 lim-
its. In addition, the FADEC unit monito rs EGT and sends an appropriate indication
to the cockpit display in case of exceedance of the limit.
The FADEC unit also provides surge recovery. Overspeed Overspeed protection logic consists of overspeed limiting loops, for both the low
and high speed rotors, which act directly upon the fuel flow command. Supple-
mentary electronic circuitry for overspeed pr otection 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 max-
imum 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.
Figure 28: Stall and Surge

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Figure 29: FADEC Processing and Fault Logic

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Power Management Autothrust Mode The autothrust mode is only available between idle and maximum (MCT) when the
aircraft is in flight.
After take-off the lever is pulled back to the maximum climb position. The auto-
thrust function will be ac tive and will provide an EPR target for:
• Max climb thrust
•Optimum thrust
• An aircraft speed (Mach number)
• A minimum thrust. Memo Mode In the memo the thrust value is frozen to the last EPR actual value, and will remain
frozen until the thrust lever is moved man ually or autothrust is reset with the auto-
thrust pushbutton switch.
When the autothrust function is disengaged while the thrust lever is in MCT/ FLX
or CL (Maximum Continuous / Flexible Take-Off or Climb) detent, the thrust is
locked until the thrust le ver is moved manually.
Memo mode or Thrust locked is entered automatically from autothrust mode
when:
• The EPR target is invalid,
• Or one of the two instinctive disconnect pushbutton switches on the thrust le-
vers is activated,
• Or autothrust signalis lost from EIU. Manual Mode This mode is entered any time the conditions for autothrust or memo modes are
not present. In this mode, thrust lever se ts an EPR value proportional to the thrust
lever position up to maximum take-off thrust. Flexible Take-Off Rating FLEXIBLE TAKE-OFF rating is set by the assumed temperature method with the
possibility to insert an assumed temperat ure value higher than the maximum one
certified for engine operation. (30 deg C.)
Autothrust Activation / Deactivation The autothrust function (ATHR) can be engaged or active.
The engagement logic is done in the Flight Management Computer (FMGC) and
the activation logic is implemented into the EEC.
The activation logic in the EEC unit is based upon two digital discretes:
ATHR engaged,
ATHR active
from the FMGC, plus an analog discrete from the instinctive disconnect pushbut-
ton on the throttle.
The ATHR function is engaged automatically in the FMGC by auto pilot mode de-
mand and manually by action on the ATHR pushbutton located on the Flight Con-
trol Unit (FCU).
The ATHR de-activation and ATHR disengagement are achieved by action on the
disconnect pushbutton located on the thrott le levers or by depressing the ATHR
pushbutton provided that the ATHR was engaged, or by selection of the reverse
thrust.
If the Alpha Floor condition is not present, setting at least one throttle lever forward
of the MCT gate leads to ATHR deactivation but maintains ATHR engaged.
If the Alpha Floor condition is present, the ATHR function can be activated regard-
less of throttle position.
The thrust is controlled by the throttle lever position and ATHR will be activated
again as soon as both throttles are set at or below MCT gate.
When ATHR is deactivated (pilot’s action or failure), the thrust is frozen to the ac-
tual value at the time of the deactivation. The thrust will be tied to the throttle lever
position as soon as the throttles have been set out of the MCT or MCL positions.
AUTOTHRUST IS ONLY ACTIVE IN EPR MODE. IN RATED & UNRATED
N1 MODE AUTOTHRUST IS LOST.

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Figure 30: Auto Thrust Definition
!4(2 053("544/.
,!4#().'
2%6%23%2
,%6%2
,%6%2
#/.42/,
).34).#4)6% !54/4(2534 $)3#/..%#4 053("544/.

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EPR Setting Requirements EPR The EEC uses closed loop control based on EPR or, if EPR is unoptainable, 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:
• Ambient temperature (T amb)
• Engine air inlet temperature (T2)
• Altitude (ALT)
• Mach number (Mn)
• Throttle Resolver Angle (TRA).
• 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 cu rrent thrust setting mode. If
the fault is no longer present, the EEC will switch to the primary control mode
(EPR). If the fault later reoc curs, reversion back to N1 mode (rated or unrated) will
result.

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Figure 31: Power Setting Requirements Schematic

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Rated N1 Setting Requirements Rated N1 The loss of either the P2 or the P 4.9 signal will cause an automatic reversion to
the rated N1 closed loop control.
This is a alternate control mode which ut ilizes to control the thrust automatically.
It is a despatchable mode but autothrust is not available when operating in this
mode. The rated N1 mode can also be manually selected by actuating the related
N1 MODE P/B switch (one per engine) that is located on the overhead panel.
The inputs required for Rated N1 control are:
• T2 and
• the Throttle Resolver Angle (TRA).
The processing of the N1 error signal is the same as for EPR error signal.

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Figure 32: Rated N1 Mode

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Unrated N1 Setting Requirements Unrated N1 The loss of the T2 signal will cause auto matic reversion to unrated N1 closed loop
control.
Max N1, N1 thrust lever, N1 mode and N1 raiting limit indications on the upper
ECAM are lost.
The input required for unrated N1 control is:
• the Throttle Resolver Angle (TRA).
The unrated N1 thrust setting requires the thrust to be set manually to an N1
speed. An overboost can occur in the unrated N1 thrust setting at the full forward
thrust lever position. Use of unrated N1 thrust setting overboost above normal rat-
ed thrust is not recommended and will result in reduced engine life.
The maximum N1 must therefore be determined from charts in the Flight Crew Op-
erating Manual (FCOM).
It is a non-despatchable mode and autothrust is not available when operating in
this mode.
The processing of the N1 error signal is the same as for the rated N1 error signal.

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Figure 33: Unrated N1 Mode

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FADEC Fault Strategy General The Electronic Engine control (EEC) system is dual, the two channels are equal.
Failures are classified as class 1, 2, 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 oc-
cur. 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 sourses:
• Local or cross- channel through the Cross channel Data Link
Some sensors can directly be synthetized by the corresponding channel Single Input Signal Failure There is no channel changeover for input signal failure, as long as the Cross
Channel Data Link is operativ.
Faults are not latched.
Automatic recovery is possible. Dual Input Signal Failure If dual input signal failure occurs, the system runs on synthetized values of the
healthiest channel.
The selected channel is one having the least significant failure. Single Output Signal Failure If an output failure occurs, there is an aut omatic switchover to the standby active
channel. T/S Action One Channel - most likely LRU failure.
Complete Output Signal Failure In case of complete output failure there will be no curren t flow through torque mo-
tors or solenoids. The associated component will be the " FAIL-SAFE " position.
If the EEC power supply is lost, the components will go into"FAILE-SAFE"
position.

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Figure 34: FADEC Single Input Signal Failure

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Component Fail Safe States *
If there is a failure of the thrust revers er hydraulic control unit directional valve
while the reverser is deployed, th e reverser will remain deployed.
COMPONENTS:
FAIL SAFE STATE:
METERING VALVE
MIN FLOW
VARIABLE STATOR VANE ACTUATOR
VANES OPEN
2.5 BLEED ACTUATOR (BSBV)
BLEED OPEN
7TH STAGE HANDLING BLEED VALVES
BLEED OPEN
10TH STAGE HANDLING BLEED VALVE
BLEED OPEN
HPT ACC VALVE
VALVE CLOSED
LPT ACC VALVE
VALVE PARTIALLY OPEN - 45%
ACOC AIR VALVE
OPEN
10TH STAGE ”MAKEUP ” AIR VALVE
OPEN
FUEL DIVERTER VALVE
FMU RETURN FLOW THROUGH FCOC (MODE 4 OR 5 )
SOLENOID DE-ENERGIZED
RETURN TO TANK VALVE
CLOSED ( MODE 3 OR 4 )
IGNITION
ON
STARTER AIR VALVE
CLOSED
P2/T2 PROBE HEAT
OFF
THRUST REVERSER CONTROL UNIT *
REVERSER STOWED

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Loss of Inputs from Aircraft
EIU SIGNALS:
NO ENGINE STARTING.
NO AUTOTHRUST ON BOTH ENGINES.
NO REVERSE THRUST
MODULATED IDLE NOT AVAILABLE.
CONTINUOUS IGNITION
ADC SIGNALS:
EEC USES ENGINE SENSORS.
BOTH TLA:
IN REVERSE: 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 SE-
LECT MCT AT SLAT RETRACTION
AUTOTHRUST CAPABILITY.
ONE TLA:
THE EEC USES THE REDUNDANT SENSOR.
BOTH 115V AC:
NO IGNITION
NO P2/T2 PROBE HEATING
BOTH 28V DC:
NO START RUN ON ALTERNATOR ABOVE 10% N2
DISAGREEMENT BETWEEN TRA:
ON GROUND:SET FORWARD IDLE
IN FLIGHT: SELECT LARGER VALUE BUT LIMIT THIS
TO MCT
ON REVERSE:SELECT REVERSE IDLE.

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Idle Control • 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 su fficiently short accelleration 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 re-
verse idle detent position.
• Bleed Idle = Bleed demand.
Bleed Idle command will set the fuel flow requested for ensuring correct aircraft
ECS system pressurization, wing anti ice and engine anti ice pressurization
(Pb-"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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Figure 35: Idle Control Requirements
Approach
Idle
Bleed
Idle
Reverse
Idle
EEC
N2 Idle
Setting
HMS
TLA (REV. IDLE)
LGCIU
1 / 2
ZONE
CONT.
0
1
2
3
FULL
0
1
3
FULL
2
EIU
WOW (GRD)
THRUST
LEVERS
LANDING
GEARS
SLAT /
FLAP
LEVER
SFCC
1 / 2
EIU
AIR
LEVER NOT ZERO
Min Idle
EIU
WING ANTI ICE
ENG ANTI ICE
EIU FAULT
PACKs
PACK
CONT.
1 / 2
ECS DEMAND
ENGINE FUEL TEMPERATURE

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N1 Speed Table
20-
.

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Figure 36: Ground Idle Speed Diagram N2
7600
50,8%
8000 53,5%
8400
56,1%
8800
58.8%
9200
61,5%
9600
64,2%
N2 ROTOR SPEED ( RPM / % )
–80 –60 –40 –20 0 +10 +20 +30 +40 +50
AMBIENT TEMPERATURE ( DEG. C. )
V2530-A5 SLS / STD GROUND IDLE ( NO OFFTAKES )
+15
57,5%

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FADEC Power Supply EIU Power Supply The EIU is powered from the aircraft electr ical power, no switching has to be done. Electronic Engine Control (EEC) Power Supply The EEC is supplied from the aircraft electrical power when engine is shutdown,
then from the EEC generator when the engine is running.
• aircraft electrical power when N2 <10%.
• EEC generator power when N2 >10%. Powering N2 <10% Each channel is independently supplied by the aircraft 28 volts through the Engine
Interface Unit.
A/C 28 VDC permits:
• automatic ground check of FADEC before engine running
• engine starting
• powering the EEC while engine reaches 10% N2.
The EIU takes power from the same bus bar as the EEC. Powering N2 >10% As soon as engine is running above 10% N2, the EEC generator can supply di-
rectly the EEC.
The EEC generator supplies each channel with three-phase AC. Two TRU’s in the
EEC provides 28VDC to each EEC channel. Auto Depowering The FADEC is automatically depowered on ground, through the EIU after engine
shutdown.
EEC automatic depowering on ground:
• after 5 mn of A/C power up.
• after 5 mn of engine shutdown
An action on the ENG FIRE P/B provides EEC power cut off.
FADEC Ground Power Panel For maintenance purposes and MCDU engine tests, the FADEC Ground Power
Panel permits FADEC power supply to be restored on ground with engine shut
down.
When the corresponding ENG FADEC GND POWER P/B is pressed "ON" the
EEC is powered again.
Also the FADEC is repowered as soon as the engine MODE SELECTOR or
the MASTER LEVER is selected.

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Figure 37: FADEC Power Supply

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FADEC LRU‘S Electronic Engine Control (EEC)
Data Entry Plug The Data Entry Plug (DEP) provides discrete inputs to the EEC. Located to the
Junction 6 of the EEC it provides un ique engine data to channel A and B.
The data transmitted by the DEP is:
• EPR Modifier (Used for power setting)
• Engine Rating
• Engine Serial No.
If the data inputs of the data entry plug J6 are lost, then an automatic revi-
sion 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 jump-
er 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 Bulle tins or after engine overhaul, when the
EPR modifier code may need to be changed. This is accomplshed by changing
the configuration of the jumper leads in accordence 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 con nectors are tight. 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 connec-
tor to 32 Ibf in.
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. EEC DEP Tester After modifing the DEP a electrical wiring test on the data entry plug assembly-
must be performed with the tester below, to make sure the pins and jumpers are
proberly installed.

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Figure 38: EEC/ Data Entry Plug

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Electronic Engine Control (EEC) Harness (Electrical) and Pressure Connections Two identical, but separate electrical ha rnesses provide the input/output circuits
between the E.E.C. and the relevant sensor/control actuator, and the aircraft in-
terface.
The harness connectors are ’keyed’ to prevent misconnection.
Single pressure signals are directed to pressure transducers
• located within the E.E.C.
• the pressure transducers then supply digital electronic signals to chan-
nels A and B.
Electrical Connections
Front Face
Rear Face
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 itlet probe)
P2.5 booster stage outlet pressure
P5 (P4.9) L.P. Turbine exhaust pressure (P5 (P4.9) rake)
P12.5 fan outlet pressure (fan rake)
Harness Connector Plug Identification
J1 E.B.U. 4000 KSA
J2 Engine D202P
J3 Engine D203P
J4 Engine D204P
J11 Engine D211P
J5 Engine D205P
J6 Data Entry Plug
J7 E.B.U. 4000 KSB
J8 Engine D208P
J9 Engine D209P
J10 Engine D210P

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Figure 39: Electronic Engine Control (EEC)

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FADEC Sensors FADEC LRU‘S Sensors Engine Sensors T4.9 (EGT) Sensor
(Ref. 77-20-00)
N1 Sensor
(Ref. 77-10-00)
N2 Sensor
(Ref. 77-10-00)
Engine Oil Temperature Sensor
(Ref. 79-30-00)
P2/T2 Sensor
(Ref. 77-00)
P3/T3 Sensor
P4.9 (P5)

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Figure 40: FADEC Sensors

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FADEC LRU‘S Sensors P3/T3 Sensor The P3/T3 sensor monitors the pressure and temperature at the exit of the HP
compressor.
The combined sensor houses two thermocouples and one pressure inlet port.
Each thermocouple provides an independant electrical signal, proportional to tem-
perature, to one channel of the Electronic Engine Control (EEC). Purpose The purpose of the P3/T3 sensor is to provide performance data to the EEC for
starting and during transient and stea dy state operation of the engine.
Figure 41:

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Figure 42: P3/T3 Sensor
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P12.5 Sensor The P12.5 sensor is a pressure tapping at the top of the fan case. It monitors the
pressure behind the fan stator. This pressure is used for trend monitoring.
The pressure tapping is also used for the cooling air supply of the dedicated alter-
nator(see Fig.114). P2.5 / T2.5 Sensors These two sensors are located in the intermediate case. They are monitoring the
pressure and temperature between the two compressors. T2.5 is used for system
scheduling, P2.5 is used for trend monitoring.

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Figure 43: P2.5 / T2.5 Sensors

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FADEC Test General To get access to the FADEC SYSTEM REPORT / TEST menu the FADEC GRD
PWR must be switched "ON". Then press the line key adjacent to CFDS - SYS-
TEM REPORT / TEST - NEXT PAGE - ENG 1A (1B),(2A),(2B). FADEC Previous Legs Report This CFDS menu function gives access to the faults which have been detected
and stored during the previous 64 flight legs.
The Cells indicate if the failure was detected in the ground memory or the flight
memory.

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Figure 44: Previous Legs Report

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FADEC Troubleshooting Report The trouble shooting menu has 4 submenus:
•FLIGHT DATA
• GROUND DATA
• AIRCRAFT DATA
• EEC CONFIGURATION Flight Data This menu gives additional failure data (temperatures, pressures, RPM, etc.)
when a fault occured during the flight. This data is saved in a CELL. Each CELL
provides 2 menu pages of troubleshooting informations. The cell allows a identifi-
cation which CFDS FAULT message belongs to which troubleshooting data (eg.
Ground Scanning menu.)
In the example a OSPXCF (OVERSPEED CROSS CHECK FAILURE) is indicat-
ed. Ground Data This menu gives additional failure data (temperatures, pressures, RPM, etc.)
when a fault occured on ground. This data is saved in a CELL.
The cell allows a identification which CFDS FAULT message belongs to which
troubleshooting data (eg. Ground Scanning menu.) FADEC Failure Types Definition WRAP - Around Failure (WAF) A detected failure in the circuitry of a system. The EEC checks for continuity.
If failed in one channel:
• EEC switches to the other channel (the ability to switch is based on relative hel-
th of the other channel)
If failed in both channels:
• specific output is depowered (exception - solenoids are depowered in groups)
T/S ACTION:
Most likley a loose connector or chaf fed harness next LRU and finally EEC. Track-Check Failures (TKF) Failure of the system to follow the commands of the EEC.
The EEC compares feedback position against commanded position.
If failed in one channel:
• EEC switches to the other channel (the ability to switch is ba sed on relative hel-
th of the other channel)
If failed in both channels:
• Healthiest channel continues to command actuator.
T/S ACTION:
one channel - most likely LRU failure.
both channels - most likely mechanical failure, check LRU/moving mechanism.
Cross Check Failures (XCF) A detected difference in the feedbacks fr om the LRU LVDT‘s or microswitches.
The EEC compares channel A against Channel B.
Failure of TRA: EEC has specific fault accomodation based on previous value.
Failure of Reverser: EEC will select mo st stowed and will not allow a deploy.
Failure of Temperature sensors: EEC will use fail safe value.
T/S ACTION:
Most likely a LRU problem, next check harness then EEC Input Latched Failed (ILF) (Single Input Signal Failure)
There is no channel changeover for input signal failure, as long as the Cross
Channel Data Link is operativ.
Faults are not latched. Thus automatic recovery is possible.

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Figure 45: Trouble Shooting Report

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Figure 46: Flight Data / Ground Data
FADEC 1B
FLIGHT DATA
CELL: 31 FAULT: WOFWAF
PG:01
RPM: N1 = 5326 N2 = 14392
DEG C: T5 = 554.0 T2 = 26.0
TCJC = 42.0FLTPH = 3
PSIA: PB = 458.5 P2 = 14.62
MN = .117 HOURS = 571.0
FADEC FAULT CELL
N1 RPM
T5 Temperature
( T4.9 EGT )
Cold Junction Temperature
( Actual Temp. in EEC )
Air Pressure on Eng. Station 3
( PB = Burner Pressure )
Mach Number
Fault Code
N2 RPM
T2 Temperature ( Eng. Inlet )
Flight Phase
Total Air Pressure ( Eng. Station 2 )
EEC Operating Hours
Note: The Abbreviations used in the GROUND DATA are the same.
Page one of the Cell 31

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Figure 47: Flight Data / Ground Data
FADEC 1B
FLIGHT DATA
CELL: 31 FAULT: WOFWAF
PG:02
ALT: = 336.0 FT EPRI = 1.562
SVA : = 1.906 INCH INCOM = 1
FF = 11162 PPHBACKUP = 0
B 25 = 1.218 INCH LEG = 398.0
WOW = 1
FADEC Fault Cell
Standart Altitude
Stator Vane Actuator
( Feedback )
Fuel Flow
2.5 Bleed Actuator Feedback
Weight on Wheels
1 = Yes ( Ground )
0 = NO ( Flight )
Fault Code
EPR ( indicated )
Channel in Control
1 = Yes , 0 = No
N1 Mode
1 = Yes
0 = No ( EPR Mode )
Flight Legs
Note: The Abbreviations used in the GROUND DATA are the same.
Page two of the Cell 31

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FADEC System Test The FADEC SELF TEST should really be known as the FADEC SYSTEM TEST.
The test and results can be split into three categories described as follows. Output Driver Test This is a systen maintenance test that performs a wraparound (continuity) test of
all the EEC output driver lines and associated component wiring.
There are three possible results as follows:
4. Output Driver Test Failed - Indicates that a continuity fault was found.
5. Output Driver Test Passed - Indicates that no wraparound fault was found.
6. Output Driver Test No Run - Indicates that the test was not run because the test-
ed channel was not capable of powering the outputs. Input / lnternal Test This is the FADEC (EEC) internal check to verify that the local channel interface,
input and output circuits are functional prior to entering MENU MODE.
There are three possible results as follows:
7. Input / Internal Test Failed - Indicates t hat the activity monitor circuit test failed
or the local channel was unable to provide power to any Output or there were in-
terface or input fault.
8. Input / lnternal Test Passed - Indicate s that the activity monitor circuit passed
and that no interface or input faults were set prior to entry into menu mode.
9. Input / Internal Test No Run - Indicate s that the local cannel was not capable of
powering its outputs or that the EEC has not spent the minimum of 30 seconds in
normal mode. Pressure Sensor Test This is an internal measurement of the pressure sensors (P2, P5, Pb, PMX) in the
EEC via the local channel to make sure they are within a specified tolerance of
each other.
The three possible results are as follows:
10. Pressure Sensor(s) Failed - Indicates that an interface or range failure (from
normal mode) is set for any pressure sensor (hard failures).
11. Pressure Sensor(s) Agree - Indicates that the static pressure sensor test ran
and that all the pressure sensors are within tolerances.
12. Pressure Sensor(s) Disagree - Indicates that the static pressure sensor test
ran and any two pressure sensors were not within the specified tolerances.

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Figure 48: FADEC Self Test

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FADEC Ground Scanning This menu shows the faults which are present on ground. More information can
be obtained using the troubleshooting menu.
This menu must also be used to indicate which faults were detected in the other
FADEC TEST menus (eg. Starter Valve Test, Reverser Test, etc.)
Figure 49: Ground Scanning

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FADEC Class 3 Fault Report This menu shows all class 3 faults of the FADEC system which have to repaired
after 200 hours or during an A-maintenance check.
Figure 50: FADEC Class 3 Fault Report

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Scheduled Maintenance Report At the aircraft level :
• Level “A” faults are considered class ” 1” faults with the associated specific in-
formation to the flight crew (ECA M warnings, advisory information...).
• Level “B” faults are regrouped under the generic ECAM warning “ENG X MI-
NOR FAULT” on the A330/A340 programs and under class 2 fault messages,
with a maintenance status, on the A3 19/320/32 1 programs.
• Level “C” faults are now covered by the powerplant “Scheduled Maintenance
Report”.
This engine “Scheduled Maintenance Report” has been created by Airbus Indus-
trie and the engine manufacturers in order to fit the dispatch time limitation asso-
ciated to the engine level “C” faults ( REF. A320/A321 SIL 73-017 ).
Faults are not latched. Thus automatic recovery is possible.
Originally all “ Short Time ” (level B – 150 FH) and “ Long Time ” (level C – 500
FH) engine faults were annunciated through Class 2 status messages and thus
had to be repaired within 10 days/150 FH. This was penalizing operators who
could not take advantage of the 500 FH interval associated with “ Long Time ”
faults.
So that the operators are no longer encumbered with this situation, the manufac-
turers developed a modification such that the Long Time level C faults no longer
appear as Class 2 messages.
A new facility “ Scheduled Maintenance Report “ (SMR) was introduced.
The following Airbus modifications introdu ce ECU/EEC software standards on the
CFM56-5B, V2500-A5 and PW6000 that cause ONLY Level C (time limited) faults
to be reported in the SMR (refer to figure 1). Unlimited faults no longer appear in
the SMR: they are indicated in the dedicated FADEC Class 3 report (access
through SYSTEM REPORT/TEST - ENGINE).
this is not applicable for V2500-A1 where BOTH time limited and unlimited
engine faults are reported in the SMR
SMR Time limited faults must be corrected within:
• 500 FH of previous task accomplishment for V2500-A1 engine
• 600 FH of previous task accomplishment for V2500-A5 engine and PW6000
engine.
• 1200 FH of previous task accomplishment for CFM56-5A and 5B engines.
Figure 51: SMR Menu
<><
>
<
<
<
<
<
>

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Figure 52: Scheduled Maintenance Report

V2500-A1 EEC Standards
EIS
“SMR“
Menu

“SMR”
Menu

Long Time Dispatch faults (unasterisked)
+
Unlimited Class 3 Faults (asterisked)
Long Time Dispatch
Faults
Unlimited Class 3 Faults

“CLASS 3
Faults”
Menu
“CLASS 3
Faults”
Menu
and subsequent
and subsequent
V2500-A5 EECStandards
SCN 10A

Unlimited Class 3
Faults
“CLASS 3
FAULTS”
Menu
V2500-A5
SCN 12C

Unlimited Class 3
Faults
+
V2500-A1

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Engine Interface Unit EIU Presentation Two EIUs are fitted on each aircraft, one for engine 1, one for engine 2
Each EIU, located in the electronics bay 80VU, is an interface concentrator be-
tween the airframe and the corresponding FADEC located on the engine, thus re-
ducing the number of wires. EIUs are active at least from engine starting to engine
shutdown, they are essential to start the engine.
The main functions of the EIU are:
• to concentrate data from cockpit panels and different electronic boxes to the
associated FADEC on each engine,
• to insure the segregation of the two engines,
• to select the airframe electrical supplies for the FADEC,
• to give to the airframe the necessary logic and information from engine to other
systems (APU, ECS, Bleed Air, Maintenance). EIU Input Description EIU Input from the EEC The EIU acquires two ARINC 429 output data buses from the associated EEC
(one from each channel) and it reads data from the channel in control. When some
data are not available on the channel in control, data from the other channel are
used.
In the case where EIU is not able to identify the chann el in control, it will assume
Channel A as in control.
The EIU looks at particular engine data on the EEC digital data flow to interface
them with other aircraft computers and with engine cockpit panels. EIU Output to the EEC Through its output ARINC 429 data bus, the EIU transmits data coming from all
the A/C computers which have to communicate with the EEC, except from ADCs
and throttle which communicate directly with the EEC.
There is no data flow during EIU internal test or initialization.
Figure 53: EIU Location

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Figure 54: EIU Schematic

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EIU Interfaces
SIGNALS
PURPOSE
WING ANTI-ICE SWITCH
ENGINE BLEED COMPUTATION LOCIG
ENGINE FIRE P/B SIGNAL
FADEC ENGINE SHUTDOWN LOGIC
LOW OIL PRESSURE SWITCH (AND GROUND)
-COCKPIT WARNING SIGNALS
-HYDRAULIC MONITORING
-WINDOW AND PROBE HEATING SYSTEM
-AVIONIC VENTILATION SYSTEM
-RAIN REPELLENT SYSTEM
-CIDS,CVR,DFDR
FADEC GROUND POWER P/B
FADEC POWER SUPPLY LOGIC
LGCIU 1 AND 2 (GROUND SIGNAL)
THRUST REVERSER AND IDLE LOGIG
SFCC 1 AND 2
ENGINE FLIGHT IDLE COMPUTATION LOGIC
SEC 1 ,2 AND 3
THRUST REVERSER INHIBITION CONTROL
FLSCU 1 AND 2
HEAT MANAGEMENT SYSTEM FUEL RETURN V ALVECONTROL
ENGINE SELECTED
ENGINE 1 OR 2 INDENTIFICATION
OIL PRESSURE,OIL QUANTITY AND OIL TEMPERA TURE
INDICATION ECAM
NACELLE TEMPERA TURE
INDICATING (ECAM)
START VALVE POSITION (FROM EEC)
ECS FOR AUTOMATIC PACK VALVE CLOSURE, DURINGENGINE START
N2 GREATER THAN MINIMUM IDLE (FROM EEC)
FUNCTIONAL TEST INHIBITION OF THE RADIO AL TIMETER TRANSCEIVER
-BLUE HYDRAULIC SYSTEM PUMP CONTROL
ENGINE START FAULT SIGNAL
ILLUMINATION OF FAULT LIGHT ON THE ENGINE STARTPANEL
APU BOOST DEMAND SIGNAL (EIU)
MAIN ENGINE START MODE TO THE APU ELECTRONICCONTROL BOX
TLA IN TAKE-OFF POSITION (MIN. T/O N2, FROM EEC)
PACK CONTROLLER FOR INLET FLAP CLOSURE
-AVIONIC EQUIPMENT VENTILATION CONTROLLER
( CLOSED CIRCUIT CONFIGURATION )
-CABIN PRESSURIZATION COMPUTER PRE-PRESSUR-
IZATION MODE
THRUST REVERSER (FROM SEC 1,2 AND 3 )
THRUST REVERSER INHIBITION RELAY

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CFDS System Report/Test EIU This Page shows the menu of the Engine Interface Unit (EIU).
The EIU is a Type 1 System.
The EIU is availlable in CFDS back up Mode.
Figure 55: EIU Menu

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LAST Leg Report Last leg Report
Here are Displayed the Internal EIU Faillu res that Occured during Last Flights.
LRU Indentification Shows the EIU part number.
Figure 56: Last Leg Rep./ LRU Indentification

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Ground Scanning This Page gives the EIU Faillures still presend on Ground.
• RTOK means Re - Test Ok, you can ignore this Fault
Figure 57: Ground Scanning

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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 or1.
The DISCRETE OUTPUT SIMULATION can operate systems and compo-
nents without special indication on th e MCDU. Make allways sure that the
working areas are clear!
For the simulation refer to AMM 73-25-34, (TASK 73-25-34-860-041).
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 to"APU BOOST" discrete output status:
"APU BOOST"becomes "1" and the EIU sends the APU BOOST command to the
59KD ECB.
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.
• APU BST2 line key has no boost effect on the APU.
THE APU BOOST FUNCTION SHOULD NOT BE USED UNLESS STRICT-
LY REQUIRED FOR TROUBLE SHOOTING PURPOSE THUS TO AVOID
PREMATURE CORE PERFORMANCE DETERIORATION. IF REALLY RE-
QUIRED, DO NOT OPERATE THE APU FOR MORE THAN ONE (1)
MINUTE IN BOOST MODE CONDITION.
NOTE THAT FOR TROUBLE SHOOTING AN ECS SYSTEM MALFUNC-
TION, THE ENGINE BLEED SHOULD BE PREFERRED. Simulation: " FAULT " Use this key to simulate a disagree between the position and the command of the
HP fuel valve through the MCDU.
FAULT discrete output status becomes "1" and FAULT legend of the 5KS1(2) an-
nunciator light comes on.
Simulation: "LOP GND 1 " Use this key to simulate OIL LOW PRESS & GND for these systems through the
MCDU:
Blue/Yellow-main-hydraulic-pressure power warning-indicating, WHC2, PHC2,
Green-main-hydraulic PWR RVSR indicating, FCDC1, FCDC2.
REMOVE THE PROTECTIVE COVERS FROM THE PROBES BEFORE
YOU DO THE TEST.
• B(Y) ELEC PUMP LO PR warning message inhibition stops
• The PHC2 controls a low probe heating level for pitot 2
• The WHC2 controls a low windshield (F/O) heating level
• The 3DB1 and 3DB2 rain repellent valve can open
The LOP GND1 discrete is used to inhibit the Flight Control System test
through the CFDS. Access to this menu is prohibited by the CFDS architec-
ture as long as you work on the EIU DISCRETE OUTPUTS menu. Simulation: "LOP GND 2 " Use this key to simulate OIL LOW PRESS & GND for these systems through the
MCDU:
PHC1, PHC3, WHC1, AEVC, DFDR and CVR.
When the line key adjacent to LOP "LOP GND2 " discrete output status becomes
GND2 "0".
• The PHC1 and PHC3 control a low probe heating level for pitots 1 and 3
• The WHC1 controls a low captain windshield heating level
• The CVR and DFDR are set to on
When you simulate LOP GND2 to "0" the horn is inhibited if the airflow ex-
traction is low in the avionics compartment. Simulation: " T/R INHIB " To simulate the authorization of closure of the thrust reverser directional control
valve solenoid (through the relay 14KS1(2)) through the MCDU.
T/R INHIB discrete output status becomes "1" and the 14KS1(2) inhibition relay is
energized. This permits the energization of the directional-control-valve solenoid

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Figure 58: Discrete Outputs Simulation

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EIU CFDS Discrete Outputs Simulation 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 status FUEL PN
becomes "1" and the zone controller 8HK will receive the HP FUEL VALVE 1(2)
open condition.
The zone controller uses the HP fuel va lve position to make the bleed status
on label 061. Then it sends it to th e EEC through the EIU (label 030). This
input can change the bleed status only if the PRV opens (engine in opera-
tion). Simulation of " 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.
The PACK FLOW control valve 1(2) require a muscle air pressure to open. Simulation of " N2 > IDLE " To simulate "N2 > IDLE" for the following systems:
XCVR radio altimeter 25A
Blue main hydraulic power
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).
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 OUT-
PUTSmenu.
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 be-
comes "1"
On the ECAM PRESS page check that the inlet and extract skin air valves close.

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Figure 59: Discrete Outputs Simulation

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EIU Discrete Outputs Many systems get the engine "on" or "off" signal. This signal is switched via the
Oil Low Press and Ground relay. The relay is directley triggert from the EIU.
Low Oil Pressure Switching via EIU
• To CIDS (23-73)
• To DFDRS INTCON Monitoring (31-33)
• To CVR power Supply (23-71)
• To Avionics Equipment Ventilation (21-26)
• To WHC (30-42)
• To PHC (30-31)
• To FCDC (27-95)
• To Blue Main Hydraulik PWR(29-12)
• To Valve Rain RPLNT. (30-45)
• To Green Main HYD PWR RSVR Indicating (29-11)
• To Yellow Main HYD PWR RSVR Indicating (29-13)
• To Blue Main HYD PWR RSVR Warning / Indicating (29-12)

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Figure 60: EIU Discrete Outputs

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74 Ignition - V2500A

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74-00 Ignition System Presentation General System Operation Dual ignition is automatically selected for:
• all inflight starts
• manual start attempts
• continuous ignition
Single alternate ignition is selected for ground auto starts. System Test The system can be checked on the ground, with the engine shutdown, through the
CFDS maintenance menu. Ignition System Components The system comprises:
• one ignition relay box
• two ignition exiter units
• two igniter plugs - located in the combustion system adjacent to No‘s 7&8 fuel
spray nozzles.
• two air cooled H.T. ignition connector leads (cooling is provided by fan air).
Ignition Relay Box
The ignition sytem utilises 115V AC supp lied from the AC 115V normal and stand-
by bus bars to the relay box.
The 115V relays which are used to connect / isolate the supplies are located in the
relay box and are controlled by signals from the EEC.
The same relay box also houses the relays which control the 115V AC sup-
plies for P2/T2 probe heating.
According to M.E.L. the IGN. system A is required as minimum!

74-00-3
Power Plant V2500A
74-00 Ignition System Presentation
Training Manual
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Copyright by SRTechnics
Figure 1: Ignition System Component

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Training Manual
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Power Plant V2500A
74-00 Ignition System Presentation
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Copyright by SRTechnics
Ignition Starting - Operation Description The ignition circuit is supplied with 115 VAC - 400Hz. The electrical power is sup-
plied 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 be-
cause:
• the two ignition systems are independent
• the EEC selects alternately ignition system A or B. 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 be-
tween 10%-16% depending on TAT and keeps it energized until N2 reaches 43%.
For inflight restart the EEC selects simult aneously 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 ignitors) In flight after en-
gine restart, if the selector is maintain ed 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 au-
tomatically the sequence by closing the star ter shut-off valve and the HP fuel shut-
off valve and deenergizing the ignitors. Ignition during Alternate Start Sequence
(Manual Start Procedure) When a manual start sequence has been activated by the EEC (ENG/MODE se-
lector switch in IGN/START position and the ENG/MAN START pushbutton switch
selected to ON) the EEC energizes both ignition exciters.
The deenergization of the ignition excite rs is automatically commanded by the
EEC when engine N2 speed reaches 43%. (Starter cut-out)
Positioning of the MASTER control switch to OFF, during that starting sequence,
results in ignition exciter deenergization.
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. Automatic selection The EEC selects automatically the continuous ignition in some specific conditions:
• 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 run-
ning, 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 121VU

74-00-5
Power Plant V2500A
74-00 Ignition System Presentation
Training Manual
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Copyright by SRTechnics
Figure 2: Ignition and Starting System Eng. 1
MAN START
FILTER
115VAC
1JH
5A
ENG 1&2/IGN/SYS A 49VU
2JH1
3A
2KS1
3A
4KS1
3A
3JH1
3A
ENG/1GN 1/SYS B 121VU
FIRE P/BSW
1WD
3KC
ON
OFF
CRANK
NORM
IGN/START
6KS
ON
9KS1
MAN START 11KS1 EIU2 86VU
FILTER
ECAM
LABEL 031 16 15 14 13 12 11
ON
OFF
CRANK
NORM
IGN/
START
RELAY BOX
A
B
A
A1
A2
B1
B2
4000KS1 EEC1
A
B
ENGINE 1
PERMANENT
MAGNETIC
ALTERNATOR
B
A
STARTER
VA LV E
CHA
CHB
CHA
CHB
EXCITER A
EXCITER B
IGNITER
PLUG A
IGNITER
PLUG B
FILTER
FILTER
FILTER
4100 KS
401XP
ESS BUS
115VAC
901XP
STAT INV
28VDC
401PP
ESS BUS
28VDC
301PP
BAT BUS
BUS BAR
115VAC
103XP
BUS 1

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74-00 Ignition System Presentation
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Ignition System Test Igniter Plug Test The operation of the igniter plugs can be checked on the ground, engine not run-
ning, through the maintenance MENU mode of the FADEC.
The test will be performed by selecting th e corresponding IGNITOR TEST page in
the MENU and positioning the MASTER control switch to ON to have the 115VAC
power supply to the relevant engine.

74-00-7
Power Plant V2500A
74-00 Ignition System Presentation
Training Manual
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Copyright by SRTechnics
Figure 3: FADEC Ignition Test

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74-00 Ignition System Presentation
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Ignitor Test Operational Test of the Ignition System with CFDS
Each ignition system must be individually selected to be tested.
For the test procedure, refer to AMM TASK 74-00-00-710-041
During the test, an aural check of the ignitor plug operation has to be done.

74-00-9
Power Plant V2500A
74-00 Ignition System Presentation
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Figure 4: FADEC Ignition Test Cont.

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74-00 Ignition System Presentation
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Ignition Test without CFDS For the test procedure, refer to AMM TASK74-00-00-710-041-01
During the test, an aural check of the ignitor plug operation has to be done.
MAKE SURE THAT THERE IS ZERO PSI AT THE STARTER VALVE IN-
LET BEFORE YOU PUSH THE MAN START P/B.
READ THE PRESSURE ON THE ECAM START PAGE.

74-00-11
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Figure 5: Ignition Test without CFDS

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74-00 Starting 80-00
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74-00 Starting 80-00 General Starting Schematic The starting system of the engine utilizes pr essurized air to drive a turbine at high
speed. This turbine drives the engine high pressure rotor through a reduction gear
and the engine accessory drive system.
The air which is necessary to drive the starter comes from:
• either the APU
• or the second engine
• or a ground power unit.
The starter supply is controlled by a starte r shut-off valve (SOV) pneumatically op-
erated and electrically controlled. In case of failure, the SOV can be operated by
hand.
The starter valve closes when the N2 speed reaches 43 %.
The starter centrifugal clutch disengag es when N2 speed is higher than 43%.
Engine starting is controlled from the ENG start panel 115VU located on center
pedestal and ENG/MAN START switch on the overhead panel.
The starting sequence may be interrupted at any time by placing the MASTER
control lever in OFF position which overrides the FADEC. When the MASTER
control lever is in OFF position the HP fu el shut off valve is closed and the engine
is stopped.
Two procedures are applicable for engine starting: A. Normal Starting Procedure (automatic) The starting sequence is fully controlled by the FADEC and is selected when the
ENG/MODE/CRANK/NORM/IGN START selector switch is in IGN/START posi-
tion and the MASTER control lever in ON position. Start can be aborted on ground
only by the FADEC in case of failure. B. Alternative Starting Procedure This sequence controlled by the pilot is as follows:
• the ignition selector switch in IGN/START position and MAN START pushbut-
ton switch command the starter shut-off valve,
• the MASTER control lever controls the HP fuel shut-off valve.
No start abort by the FADEC in case of failure.

74-00-13
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74-00 Starting 80-00
Training Manual
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Figure 6: Starting System Schematic
%.').%
4/ .!#%,,%
!.4))#% 6!,6%
34!24%2
34!24%2
6! ,6 %
)0
(0
)0#
(06
026
4/ (9$2!5,)# 2%3%26/)2 02%3352):!4)/.
%.').% /.,9
/06
&!6
&)2%
7!,,
02%#//,%2
/6%2"/!2$
4,4
4#4
00
"-#
/4(%2
!#
3934%-3

74-00-14
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Copyright by SRTechnics
Starting Components Starter Motor The pneumatic starter motor is mounted on the forward face of the external gear-
box and provides the drive to rotate the H.P. compressor to a speed at which light
up can occur.
Attachment to the gearbox is done by a V-clamp adaptor.
The starter motor is connected by ducting to the aircraft pneumatic system.
The starter motor gears and bearings are lu bricated by an integral lubrication sys-
tem.
Servicing features include:
• oil level sight glass
• oil fill plug
• oil drain plug with magnetic chip detector Starter Motor - Operation The starter is a pneumatically driven turbin e 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 sing le 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 tur-
bine, and leaves through the air exhaust. The reduction gear train converts the
high speed, low torque rotation of the turb ine to low speed, high torque rotation of
the gear train hub.
The ratchet teeth of the gear hub engage the pawls of the output drive shaft to
transmit drive to the external gearbox, which in turn accelerates the engine H.P.
compressor rotor assembly.
When the air supply to the starter is cut off, the pawls overrun the gear train hub
ratchet teeth allowing the turbine to coast to a stop while the engine H.P.
turbine compressor assembly and, therefore, the external gearbox and starter out-
put 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, allowin g the pawls to be pulled clear of the gear
hub ratchet teeth to disengage the output drive shaft from the turbine.
Starter Air Control Valve The starter air control valve is a pneumatically operated, electrically controlled
shut-off valve positioned on the lower right hand side of the L.P. compressor (fan)
case.
The start valve controls the air flow from the starter air duct to the starter motor.
The start valve basically comprises a butt erfly type valve housed in a cylindrical
valve body with in-line flanged end connec tors, an actuator, a solenoid valve and
a pressure controller.
A micro switch provides valve position feed back information to the FADEC.

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Training Manual
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Figure 7: Starting Components

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Starter Air Control Valve Description The start air control valve is a pneumatica lly operated, electrically controlled shut-
off valve positioned on the lower right hand side of the L.P. compressor (fan) case. Manual Operation The starter air valve can be opened/ clos ed manually using a 0.375 inch square
drive. Acces is through a panel in the R. H. fan cowl. A valve position indicator is
provided on the valve body.
A micro switch provides valve position feed back information to the FADEC.
Do not operate the valve manually without positive duct pressure.
FAIL SAFE POSITION:
"SOV CLOSED"

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Training Manual
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Figure 8: Starter Air Control Valve

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74-00 Starting 80-00
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Start Air Control Valve Test Start Air Control Valve Test via CFDS The start air control valve operation may be tested via CFDS.
Refer to AMM Task 80-13-51-710-040.

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74-00 Starting 80-00
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Figure 9: Starter Valve Test via CFDS

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74-00 Starting 80-00
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Start Air Control Valve Test (Fault Detected) AMM Starter Valve Test ata 80-13-51 p507

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Power Plant V2500A
74-00 Starting 80-00
Training Manual
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Figure 10: Starter Valve Test via CFDS

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Cranking-Description Air Supply The air necessary for the starting comes from the duct connecting engine bleed
and the precooler.
The air necessary for the starter is supplied by either:
• the other engine through the crossbleed system
• the APU and in that case, all the air bled from the APU is used for starting
• an external source able to supply a pressure between 30 and 40 psig. Dry Cranking (Test No 1) Requirement
A dry motoring of the engine will be needed when:
• it is necessary to eliminate any fuel accumulated in the combustion chamber
• a leak ckeck of engine systems is needed.
To perform this operation, the starter is engaged and the engine is motored but
the HP fuel shut off valve remains cl osed and both ignition systems are OFF.
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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Figure 11: Dry Cranking Procedure

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Training Manual
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74-00 Starting 80-00
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Wet Cranking Wet Cranking (Test No 2) A wet motoring will be needed when the integrity of the fuel system has to be
checked.
If such a test is performed, both ignition systems are off (also pull the circuit break-
ers) and the starter is engaged to raise N2 up to the required speed of 20%.
The MASTER control switch is moved to ON and the exhaust nozzle of the engine
carefully monitored to detect any trace of fuel. On the ECAM the FF indication
shows approx. 180kg initial fuel flow.
When the MASTER control switch will be returned to th e OFF position to shut-off
the fuel, also the starter valve closes. The EEC automatically reengages the start-
er at 10% N2 and the engine should be motored for at least 60 seconds to elimi-
nate entrapped fuel or vapor.
The motoring can be performed for a maximum of three consecutive cycles (2 of
2 minutes and 1 of 1 minute with a cooling period of 15 seconds between each
cycles).
After three cycles or 4 miutes of contin uous cranking, stop for a cooling period of
30 minutes.

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74-00 Starting 80-00
Training Manual
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Figure 12: Wet Cranking Procedure

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Training Manual
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Power Plant V2500A
74-00 Starting 80-00
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Automatic Start The automatic start mode gives the EEC full control to automatically sequence the
starter air valve, ignition relays and the fu el on / off torque motor. Upon receipt of
the appropriate start command signals from the engine interface unit (EIU), the
EEC commands, in sequence:
• the starter air valve
• ignition exiter relay(s),
– alternatively selected for each ground start
– both selected for inflight or manual starts
• fuel on function of the torque motor which opens the shutoff valve.
During a normal start, the starter air valve and ignition exciter are automatically
turned off by the EEC at a predetermined N2 speed of 43%
Starter assist will be comanded by the EEC for inflight starts at low MACH num-
bers where windmilling conditions are insufficient for engine starting.
(The EEC has input data necessary to activate starter assist function where nec-
essary.)
In case a Auto Start is initiated and one thrust lever is not in idle position a
ECAM warning is triggert. The start sequence will contiue and the engine will
accelerate to the trust lever position. EEC Auto Start Abort The autostart procedure commences only when the engine is not running, the
mode selector set to IGN/START and the master switch is ON.
Intermittent mode selector position or m anual start push button switch selection
has no effect on autostart sequence once the autostart procedure is initiated.
Switching the master switch OFF during an autostart will close the fuel and starter
air valves and turn the ignition system off. It also resets the EEC.
The automatic start abort function is only available when N2 speed is below 43%
and in case of:
• Start valve failure
• Ignition failure
• Pressure Raising Shut Off Valve failure
• Hot start
• Hung start
• Surge
• EGT >250 deg C when restart (max 2 min)
• Loss of EGT
The oil pressure is not monitored during Auto Start!
The EEC automatically shuts off fuel, igniti on, and starter air and provides the ap-
propriate fault indication to the cockpit. (Auto Start Fault)
Autostart fault messages will be disp layed until approximately idle speed.
The EEC’s ability to shut off fuel is i nhibited above 43% N2 on the ground and at
all conditions inflight. In case of an automatic start abort, the EEC re-opens the
start valve when reaching 10% N2 for a 30 second dry motoring cycle to clear fuel
vapor and to cool the engine.
Then the operator has to select the Master switch to the OFF position by a com-
mand indicated on the ECAM page ("Master lever OFF").
The operator then has to decide to perform a new engine start or troubleshoot the
system.

74-00-27
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Figure 13: Automatic Start Procedure

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Power Plant V2500A
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Manual Start The engine manual start panel, used for manual start, is located on the overhead
panel and is composed of two manual start push button switches (one per engine).
The manual start mode limits the authority of the EEC so that the pilot can se-
quence the starter, ignition an d fuel on/off manually. This includes the ability to dry
crank or wet crank.
During manual Start operation, the EEC Auto Startabort feature is not available
and conventional monitoring of the start parameters is required.
The EEC continues to provide fault indications to the cockpit.
The manual start procedure commences when the mode selector is set to:
IGN/START,
the manual start push button switch is se t to ON and the master switch is OFF.
The starter air valve is then commanded open by the EEC.
When the master switch is turned ON (at 22% N2) during a manual start, both ig-
nitors are energized (IGN A/B) and fuel is turned on (Intial FF 180 KG/H).
Intermittent mode selector position has no effect on the manual start sequence
once the manual start procedure is initiated.
The starter air valve can be closed by selecting the manual start push button
switch OFF at any time prior to turning the master switch ON.
Once the master switch is turned ON, the manual start push button switch has no
effect on the start.
When the master switch is turned OFF, the control commands the HP fuel valve
closed, the starter air valve closed and t he ignitors off and the EEC is resetted.
Figure 14: ECAM Start Pages

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Figure 15: Manual Start Procedure

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Continuous Ignition With engine running, continuous ignition can be selected via the EEC either man-
ually using the rotary selector or automa tically by the Full Authority Digital Engine
Control (FADEC).
Figure 16: Continuous Relight Logic

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Figure 17: Continuous Ignition Logic
1
3
2
4 31-54-06
401XP.B
115VAC
ESS BUS
24-58-11
1
3
2
4 31-54-06
901XP.A
115VAC
STAT INV
BUS
24-58-14
1
3
2
4 31-54-06
204XP.A
115VAC
BUS 2
24-58-02
103XP.A
115VAC
BUS 1
24-58-02
SCH03
14F
14H
5
1
5B
1C
1A
1B
1F
AB
AB
AC
AC
AA
AA
14B
14D
4
6
9A
5A
5C
14K
15J
15A
15C
AB
AB
AC
AC
AB
AB
AA
AA
28VDC
SUPPLY SWITCHING
UNSD
UNSD
1WD
SCH02
P(G)
S(J)
R(H)
A(C)
A(C)
ENG/APU
26-21
UNSD
SCH01
1A
115VU210
76-12
CRANK
AUTO
IGNITION
SCH08
1(21)
2(22)
3(23)
4(24)
ENG/MODE/CRANK
AUTO IGN/IGN
115VU210
73-25
SCH08
C3
A
A
C2
C1
A
A
UNSD
MAN ENG START
22VU212
73-25
1KS1
EIU-1
1KS2
EIU-2
86VU128
A
B
A
B
A
B
EEC1 (EEC2)
INPUT A1
EEC1 (EEC2)
INPUT B1
BUS A
"FAULT"
LIGHT
ON
SCH13 (SCH21)
21(3)
A
A
45(2)
10(1)
A
A
8LP (19LP) BOARD
ANN LT TEST
& INTFC
70VU126
33-14
4100KS
RELAY BOX
SCH18
27 27
D
D
C
C
67 67 45 45
A
A
E
E
D
D
A
A
E
E
C
C
73-25
J9
J9
CH.A
CH.B
NM NM
J3
J3
J9
J9
CH.A
CH.B
R
<G
J3
J3
R
<G
WIRING DIAGRAM
74-31-01
74-31-02
4000JH12 1 3
A
A
EXCITER-
IGNITION, A
UNSD
4000JH2
2
1
3
A
A
EXCITER-
UNSD
73-25
SCH04
PN <P <D <P <DLK LKNP
J7
J7
J1
J1
J7
J7
J1
J1
T<P<QT<P<Q
J8
J8
J9
J9
J2
J2
J3
J3
A
B
A
B
A
B
A
B
A
B
A
B
TO EEC ARINC
INPUTS & OUTPUTS
73-25 SCH10
EIU-1 (EIU-2)
BUS A INPUT
EIU-1 (EIU-2)
BUS A INPUT
73-25 SCH10
EEC B
OUTPUT 1
EEC B
OUTPUT 2
73-25 SCH10
EEC A
OUTPUT 2
EEC A
OUTPUT 1
5KS1 (5KS2) ANN-
ENG/1 (2) FIRE/FAULT
SCH09
7 4
A
A
98 56
A
A
R
A
FIRE
FAULT
UNSD
115VU210
73-25
CH.B
CH.A
CH.B
CH.A
78 34 56 12
DISCRETE
SOLENOID
DISCRETE
SOLENOID
CHANNEL B
CHANNEL A
26-12
SCH01
9
8
9
8
85VU127
73-25
<I
<I
A
A
3A
2A
5A
4E
OFF
ON
1JH C/B
ENGINE/1 AND 2/
ING/SYS A
49VU210
2JH1 (2JH2) C/B
ENGINE/ING/ENG1 (ENG2)/
SYS A/BAT
121VU212
FIRE PNL
210
3KC (2KC) CTL SW-
ENG/MASTER 1(2)
6KS SEL SW-
9KS1 (9KS2) P/BSW-
ENG/MAN START/1(2)
436 (446)
4001JH1
PLUG-IGNITER, A
454 (464)
454 (464)
IGNITION, B
454 (464)
4000KS EEC
436 (446)
4005KS VALVE-
436 (446)
PNEUMATIC STARTER
3JH1 (3JH2) C/B
ENGINE/ING/
ENG1 (ENG2)/SYS B
121VU212
4001JH2
PLUG-IGNITER, B
454 (464)

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Figure 18: Engine Opearating Limits
Engine Rating N1 N2 EGT Max EGT Cont. EGT Start Pre start
EGT
N1 Vib N2 Vib
V2533-A5 5650 14950 650 610 635 250 5.0 5.0 V2530-A5 5650 14950 650 610 635 250 5.0 5.0 V2528-D5 5650 14950 635 610 635 250 5.0 5.0 V2527-A5 5650 14950 635(E/M) 610 635 250 5.0 5.0 V2525-D5 5650 14950 620 610 635 250 5.0 5.0 V2500-A1 5465 14915 635 610 635 250 5.0 5.0 V2524-A5 5650 14950 635 610 635 250 5.0 5.0 V2522-A5 5650 14950 635 610 635 250 5.0 5.0
E is enhanced performance. M is for the corporate A319 jet.
The following operating limits apply to all engine ratings for the oil system.
Min start Min to
1.3EPR
Min to T/O Max trans Max limit Minimum Maximum
Oil Pressure 60 psi ISA
dependant
Oil temperature -40 deg.c -10 deg.c 50 deg.c 156 deg.c
amber
165 deg.c red

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75 Engine Air - V2500A

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75-00 System Presentation General • Nacelle Compartement and Accessory Cooling
• Bearing Compartment Cooling and Sealing
• HP TurbineCooling
• HP / LP Turbine Clearance Control System (ACC)
• Ignition System Cooling (REF, ATA 74)
75-30 Compressor Control
• LP Compressor Airflow Control System
• HP Compressor Airflow Control System
75-40 Nacelle Temperature Indicating
The external air system consits of the following subsystems:
• Fuel control system air bleed
• HP / LP turbine active clearance control
• High energy igniter harness cooling air
• Engine bleed air.
The internal air system consits of:
• Propulsion airflow (secondary & primary flows)
• Bearing compartments pressurizing air
• Cooling air FADEC Compressor and Clearance Control General The engine compressor and clearance control system are provided with servo
valves operated by fuel pressure, but the HP compressor handling bleed valves
are operated by pneumatic pressure.
The actuators have two feedback signals, one for channel A one for channel B,
exept for the HP compressor handling bleed valves which do not have any position
feedback.
There is a cross-talk between the two channels, so that each channel knows the
position sensed by the other channel.
Compressor Control General The booster stage bleed valve, the variabl e 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. Booster Stage Bleed Valve (BSBV) Control The BSBV position is controlled by the EEC. 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 thro ugh 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. 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 compres-
sion with the high pressure stages. This va riation of vane position changes the ef-
fective angle at which the air flows acro ss the compressor blades and vanes. The
VSV angle determines the compression characteristics (direction and velocity) for
any particular stage at compression. HP Compressor Bleed Valves The 7th and 10th stages bleed valves maintain a more stable operation of the
compressor.

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Figure 1: Compressor Control Schematic

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75-31 LP Comp. Air Flow Sys. 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:
• Engine start.
• Engine transient operation. Description
General the airflow control system includes:
1. Two bleed-valve actuating rods
2. Pisten Jack Fork End
3. An LPC bleed-master actuator
4. An LPC bleed-slave actuator
5. Intermediate Structure A booster bleed valve and actuating mechanism The airflow control system automatically op erates 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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Figure 2: Booster Stage Bleed Valve System

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BSBV Actuating Mechanism Booster Bleed Valve and Actuating Mechanism
Description The bleed valve and actuating mechanism is a sub - assembly which 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.

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Figure 3: BSBV and Actuating Mechanism

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75-32 HP Comp. Air Flow Sys. VSV System Components The four stages of variable incidence st ators comprise inlet guide vanes to stage
3 and stages 3, 4 and 5 stator vanes. General The purpose of this system is to posit ion the Inlet Guide Vanes (IGV) and stator
vanes, using a fuel driven hydraulic actuat or, in response to electrical signals pro-
vided 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
T 2.6 (synteziesed 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 regu-
lated by accurate control of high pressure fuel flow to one or other side of a differ-
ential area piston. The piston has an ex ternally 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 LV DT completes the electronic control loop
by providing a signal of actuator posit ion 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 VSV Actuator Dual wound torque motors convert electrically isolated drive signals from each
channel of the Electronics Engine Control (EEC) into hydraulic drive signals to po-
sition the actuator piston. If power to the st ator 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
• actuator/crankshaft drag link
• crankshaft (steel)
• four crankshaft/unison ring drag links
• four unison rings
• spindle levers (titanium)
• variable IGVs 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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Figure 4: VSV System Components

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VSV Rigging Variable Stator Vane System (VSVS)
Actuator Installation / Rigging Before the actuator is removed it is important that the VSV crankshaft assembly is
locked in order to prevent damage to the stator vanes.
Rig pins are provided to lock the cran kshaft and the actuator, as shown below.
After the fuel supply and return tubes have been disconnected the crankshaft
should be rotated to align the rig pin holes in the input lever and the front bearing
housing.
Spanner (Wrench) flats are provided on the crankshaft for this purpose. Installing
the rig pin locks the crankshaft assembly with the actuator and vanes in the high
speed position (actuator fully retracted).

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Figure 5: VSV Actuator Rig

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Handling Bleed Valves Handling bleed valves are fitted to the H.P. compressor to improve engine starting,
and prevent engine surge when the compressor is operating at off-design condi-
tions.
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 as a function of N2 and
T2.6 (N2 corrected).
When the bleed valves are open, H.P. compressor air bleeds into the fan duct
through ports in the inner barrel of the 'C' ducts.
The servo air used to operate the bleed valves is H.P. compressor delivery air
known as P3 or Pb.
The bleed valves are arranged radially around the H.P. compressor case as
shown below.
Silencers are used on some bleed valves.
All the bleed valves are spring loaded to the open position and as a result will al-
ways be in the correct position (open) for starting.

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Figure 6: HP Compressor Bleed Valves

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Handling Bleed Valves Function 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 clos e the valve. The valve is kept in the
open position by servo air (P3) supplied fr om the solenoid control valve (solenoid
de-energised). The bleed valves will be closed at the co rrect 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. Operating Schedule The schedule for one bleed valve - 7C - is shown, in detail, below. Steady State It can be seen that the valve will be co mmanded 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 belo w the transient closing speed the valve will
remain open until the acceleration timer expires (30 seconds).
During decelerations the valve will be co mmanded open whenever N2 is below the
transient opening speed. The valve remains open until the deceleration ceases
and a deceleration time, 2 seconds, expires.
The transient regime is slightly modifi ed for operation above 15000 ft but op-
erates 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 command-
ed open, if the speed is below the open speed, and will remain open until the en-
gine restabilises.
During an engine deceleration the reverse operation occurs and the bleed valve
opens.
Handling bleed valves (surge bleed)
The bleed valves and the solenoid control valves all operate in the same manner.
FAIL SAFE POSITION: "7th and 10th OPEN".

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Condition N2C26 7A status
7B status
7C status
10 status
SS ONLY
SS ONLY
SS & TR
SS ONLY
starting <8623 open
open - closes before
reaching idle
open - closes on reaching
idle
open - closes before
reaching idle
idle/taxi 8623 open
closed
closed*
closed
take off acceleration 8623<N2C26<12100 open>closed
closed
opens on detection of
acceleration, then closes
at mid-power
closed
take off (including derates)
Begin T/O 12100, 90% de-
rated 12044, 80% derated
11965
closed
closed
closed*
closed
climb
Begin 11869, Mid 12142, End
12294
closed
closed
closed*
closed
cruise (mid) ~12100 closed
closed
closed*
closed
end of cruise deceleration 12000<N2C26<10819 closed>open
closed
opens on detection of
deceleration, then closes
closed
top of descent 10819 open
closed
closed*
closed
mid descent 10211 open
closed
closed*
closed
end of descent 8509 open
closed
closed*
closed
approach 9085<N2C26<11560 open
closed
closed*,***
closed
touchdown 9745 open
closed
closed*
closed
reverse 12135
open (if N2C26 below
certain threshold)
closed
open (if N2C26 below
certain threshold)
closed
idle/taxi <8623 open
closed
closed*
closed
surge recovery NA
open (if N2C26 below
certain threshold)
closed
open (if N2C26 below
certain threshold)
open (if N2C26 below
certain threshold)
* bleed valve will open in response to throttle lever angle variation
** the holding condition varies based on aircraft weight, landin g runway altitude, airport traffi c, typical mission etc. the EE C does not have a unique TRA position for holding
conditions. generally a 30% maxmum take off thrust is used for holding condition power setting.
*** bleed valve will open when approach mode is selected and engine switches from low to high idle

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Figure 7: HBV OPEN/CLOSED Schematic

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Bleed Valve Locations The bleed valves are arranged radially around the HP compressor case as shown
below.

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Figure 8: Bleed Control Valve Solenoids

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Handling Bleed Valve Malfunctions A engineering order (010169) is released to cover this problems.
7TH / 10TH STAGE HANDLING BLEED VALVES STICKING
Hung starts or starting stalls experien ced 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 valve‘s on:
• the ground and airstart capability,
• the engine operability (surge free operation)
• the engine performance (EGT, fuel consumption)
have been assessed and are summarized in the following tables:
A bleed test set is provided to check the bleed valves and solenoid valves
for proper function.
Figure 9: HDLG Bleed Valves Malfunction Tables

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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 sup-
plements 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, 10th Stage "Make-Up" Air System
Introduction The make up air discharges into the area around No4 bearing housing and sup-
plements 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 energise / deenergise 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 de-energised) 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.
The E.E.C. will keep the ai r valve open at all engine operating phases except
cruise. The valve incorporates 2 micro switches for transmitting valve posi-
tion to the E.E.C channel A & B.
The "fail safe" position is valve open, solenoid de-energised.
Figure 10: HPC 10 Cooling Tubes 4 Off

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Figure 11: HP Turbine Cooling Air Schematic

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Figure 12: Stage10 to HPT Air Control Valve

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Figure 13: Air Systems Schematic

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Turbine Cooling Control The EEC controls the actuation of an Active Clearance Control (ACC) valve for the
HP and LP turbine active clearance control and a 10th stage make-up air valve for
supplementary internal cooling of the turbines. HP Turbine (10th Stage) Cooling Air Control The HP turbine cooling air valve (make up air valve) supplies supplemental air
(from HPcompressor 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. HPT/LPT Active Clearance Control (HPT/LPT ACC) 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 opening and closing of the ACC system by monitoring input
signals of:
• Corrected N2.
• Altitude.
The dual track LVDTs will send feedback signals to the EEC of the ACC system
operation. Operating Schedule The graph shown below represents the conditions of engine operation and the ef-
fect it has on the modulating air valves position. Position A At position A the engine is shut down. This is also the failsafe position.
• HPT ACC valve is closed.
• LPT ACC valve is at –44%. Position B This position represents idling conditions.
• HPT ACC is closed.
• LPT ACC is closed.
Position C This position represents a typical take off condition. This position is altitude de-
pendent.
• HPT ACC is starting to open.
• LPT ACC is at 70%. Position D and E These positions represent typically cruise and top of descent conditions. This po-
sition is altitude dependent.
• HPT ACC at D is 30% and at E is fully open.
• LPT ACC is fully open at points D and E. Fail Safe When there is no torque motor current or no fuel servo pressure, the actuator pis-
ton 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)

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Figure 14: Turbine Cooling Control Schematic

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HPT / LPT Active Clearance Cont. Sys. 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 Clear-
ance Control Valves which direct air for both HP and LP turbine case cooling.
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 al-
low a permanent cooling of the HP turbine.
In case of a valve removal / installation the same configuration must be pro-
vided on the new valve.
If the plugs must be removed, there is a storage bracket provided on the ac-
tuator rod. Do not throw the plugs away !
Figure 15: ACC Valve

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Figure 16: LPT / HPT Active Clearance Control Valve

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HPT / LPT Cooling Manifolds HP Turbine Manifold The assembly consists of a left and righ t hand tube assemblies which are a simple
push fit into the manifold.
Air outlet holes on the inner face of the t ubes direct the air onto the HP turbine cas-
ings. LP Turbine Manifold The assembly consists of a upper and lower tube assemblies with integral mani-
folds, both ends of the cooling tubes are sealed.
Air outlet holes on the inner surfaces direct the air onto the LP turbine cases.

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Figure 17: HPT / LPT Cooling Manifolds

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Nacelle Ventilation Ventilation is provided for the fan compartment Zone 1, and the core compartment
Zone 2 to:
• prevent accessory and component overheating
• prevent the accumulation of flammable vapours. Zone 1 Ventilation Ram air enters the zone through an inlet located on the upper L.H. side of the air
intake cowl. The air circulates through the fan compartment and exits at the ex-
haust 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 e ffected in the same way as when the engine is running.

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Figure 18: Nacelle Ventilation

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75-41 Nacelle Temperature 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 Trans forms the Information to digital Form.
The EIU Transmits the Data to the ECAM System.
On Classic Aircraft the nacelle temperature is displayed if the system is not in en-
gine starting mode and one of the two te mperatures reaches the advisory thresh-
old.
A advisory indication will be created on the engine system page when the temper-
ature reaches approx. 300 - 320°C.
On enhanced aircraft the nacelle temperature indication is permanently displayed.
Figure 19: Classic Lower ECAM Indication
Figure 20: Enhanced Lower ECAM Indication

14

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'7 ' ,/!$

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Figure 21: Nacelle Temperature Sensor

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76 Engine Controls - V2500A

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76-00 Engine Controls Throttle Control System General The throttle control system consist of:
• the throttle control lever
• the throttle control artificial feel unit (Mechanical Box)
• the thrust control unit
• 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. Thrust Control Unit The Thrust Control Unit contains two resolvers, eac h of which sends the thrust le-
ver position to the Electronic Engine Control. The extraction current for the resolv-
ers is provided by the EEC. Autothrust Disconnect Pushbutton The autothrust instinctive disconnect pushbutton can be used to disengage the
autothrust function. Thrust Levers General The thrust levers comprises:
• a thrust lever which incorporates stop devices and autothrust instinctive dis-
connect pushbutton switch
• a graduated fixed sector
• 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). Reverse Thrust Latching Lever To obtain reverse thrust settings, the re vers thrust laching lever must be lifted.
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:

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Figure 1: Engine Thrust Lever Control
ENGINE THRUST LEVER CONTROL
REVERSE THRUST
LATCHING LEVER
RESOLVER 1
RESOLVER 2
CHANNEL A
CHANNEL B
– FUEL
METERING
VALVE
AUTOTHRUST
DISCONNECT PB
THRUST LEVER
REVERSE THRUST
LATCHING LEVER
MECHANICAL
BOX
THRUST CONTRO
L
UNIT
FMU
EEC

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Bump Rating Push Button(A1 Engined Aircraft only) 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 re-rated 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 g enerate a digital signal to both EECs 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.
Inflight, 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 inflight (EPR or rated N1 mode) under the following
conditions.
• Bump rating initially selected on the ground.
• TO/GA thrust lever position set.
• Airplane is within the takeoff envelope.
The bump rating is a non-standard rating and is only available on certain desig-
nated operator missions.
Use of the bump rating must be recorded. This information is fo r tracking by main-
tenance personnel.
Figure 2: Bump Push Bottons

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Figure 3: Flat Rated Thrust Control and Modification Inputs

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Artificial Feel Unit (Mechanical Box) The Throttle control artificial feel unit is located below the cockpit center pedestal.
This artificial feel unit is connected to en gine 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 syste m which provides a load feedback to the
throttle control lever.
This artificial feel 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 transmit ted 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 fr iction brake assembly shaft. This move-
ment 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.

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Figure 4: Mechanical Boxes
MECHANICAL BOX(ES)
MECHANICAL
BOXES
ADJUSTMENT
SCREW
RIGGING
POINT
An adjustment screw
is provided at the
lower part of each
mechanical box to
adjust the artificial
feel.
DETENT FORCE
ADJUSTMENT
LEVER
ASSEMBLY
FRICTION
ADJUSTMENT
SCREW
BELLCRANK
ASSEMBLY
FRICTION
BRAKE
ASSEMBLY
GEAR
ASSEMBLY
CASING

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Throttle Control Unit The throttle control unit comprises:
• an input lever
• mechanical stops which limit the angular range
• 2 resolvers whose signals are dedicated to the EEC (one resolver per channel
of the EEC)
• 6 potentiometers fitted three by three. Their signals are used by the flight con-
trol system
• a device which drives the resolver and the potentiometer
• 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 as sembled face to face. Each sector drives
itself a set of one resolver and three potentiometers.
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:
• less than -47.5 deg. TRA or
• greater than 98.8 deg. TRA as resolver position signal failure.
The EEC incorporates a resolver fault accomodation logic. This logic allows en-
gine operation after a failure or a complete loss of the throttle resolver position sig-
nal.

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Figure 5: Thrust Control Units
THRUST CONTROL UNIT(S)
RESOLVER
RIGGING
POINT
ELECTRICAL
CONNECTORS
– 2 units
Each unit consists of :
– 2 resolvers
– 6 potentiometers.
3 COUPLED POTENTIOMETERS
C
C
C
C
TOOTHED
SEGMENTS
ONE
RESOLVER
CONNECTORS
3 COUPLED
POTENTIOMETERS
CONTROL
LEVER
CROSS SECTION

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Rigging The throttle control levers must be at the idle stop position to perform the rigging
procedure.

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Figure 6: Thrust Control System Rigging

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AIDS Alpha Call Up of TRA Using the Aids Alpha call up it is possibl e to check both TRA (Thrust Resolver An-
gle)
Figure 7: Alpha Call-up TRA

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77 Indicating - V2500A

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77-00 Engine Indicating Presentation Indication General Primary Engine Display The primary engine parameters listed below are permanently displayed on the En-
gine and Warning display (E/WD):
• Engine Pressure Ratio (EPR)
• Exhaust Gas Temperature (EGT)
• N1 (low rotor speed)
• N2 (high rotor speed)
• 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 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 be-
comes amber.
Only in case of certain system faults and flight phases a warning message ap-
pears on the Engine Warning Display. Secondary Engine Display The lower display shows the secondary engine parameters listed below. The en-
gine page is available for display by command, manually or automatically during
engine start or in case of system fault:
• 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:
– OIL FILTER COLG
– Fuel FILTER CLOG
– No. 4 BRG SCAV VALVE with valve position
Some engine parameters also displayed on the CRUISE page

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Figure 1: Engine ECAM Indications
T.OAUTO BRK
SIGNS ON
SPLRS ARM
FLAPS T.O
T.O CONFIG NORMAL
T.O INHIBIT
IGNITION
LDG LT
EPR
EGT
N1
N2
C
%
%
89.
0
89.
0
82.
6
82.
6
464464
4
10
4
10
4
8
4
8
1.2
1.4
1.6
1
1.2
1.4
1.6
1
1.
296
1.
296
1.
296
3040 3040
6700
2
FLX56 C
KG/H
KG
FLAP
FOB:
FF
SF
ENGINE
F.USED
207KG215
OIL
PSI
QT
25
VIB
VIB
(N1)
(N2)
300 300
0
25
0 0
0
14
.5
140
14
.5
166 166
140
1.
6
1.
6
0.
3
0.
3
C
TAT
SAT
+10
+10
C
C25
03
H
GW54700KG
PSI 350 PSI
AB
ON
OFF
ONO OOFOFF
FF
MODE
NORM
2
ENG
MASTER 1 MASTER 2
CRANK
IGN
START
ON
OFF
ENG
2
ENG
1
FIRE
F
AU
L
T
FIRE F
AU
L
T
1
ON ON ON
ON
ENG ENG
MAN STARTN1 MODE
12 12
MASTER
WARN MASTER
CAUT
Upper E/WD Lower ENG System Display
RH lower Overhead Panel
Overhead Maint. PanelAttention Gatters
ENG Control
Panel
ON ON
ENG
FADEC GND PWR
12

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77-10 Power Indicating 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 this 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 ERP value to the ECAM. Each of the two chan-
nels 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 th rust lever position.
5. REV indication
Appears in amber when one reverser is unstowed or unlocked or inadvertenly
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 asso-
ciated EPR rating is displayed in blue. In MREV no EPR value is displayed.
Thrust limit mode is displayed in digita l form, it indicates the mode which the EPR
limit value will be computed.
• In flight (or on ground with ENG stopped):
– The selected mode corresponds to the detent of the most advanced thrust
lever position
– 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 re-
ceiving the most advanced thrust lever position).
1.When a thrust lever is set between two positions the EEC selects the
rating limit corresponding to the highest mode.
2.When idle is selected the EEC selects CL
3.When M REV is selected, the EPR rating limit value is re placed by am-
ber crosses (M REV mode is limited by N1)
• On ground (with engines running)
– With engines running, on ground, whatever the lever position is, this limit
corresponds to: TO GA thrust limit.
– With engine running, on ground, if FLX mode is selected, FLX EPR is dis-
played 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, the analog indication is replaced by
amber crosses, the analog indicator and the marks on the circle disappear, the cir-
cle becomes amber.

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Figure 2: EPR Indication - Upper ECAM Display Unit
T.OAUTO BRK
SIGNS ON
SPLRS ARM
FLAPS T.O
T.O CONFIG NORMAL
T.O INHIBIT
IGNITION
LDG LT
EPR
EGT
N1
N2
C
%
%
89.
0
89.
0
82.
6
82.
6
464464
4
10
4
10
4
8
4
8
1.2
1.4
1.6
1
1.2
1.4
1.6
1
1.
296
1.
152
1.
503
FLX35 C
1.
520
TOGA MCT CL MREV
x
OR
OR
OR
OR
REV
5
4
2
3
6
1

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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 meas-
ures 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 accom-
plished by two resistance-sensing elements housed in the P2/T2 sensor body.
Each channel of the Electronic Engine Control (EEC) monitors one of these resist-
ance elements and converts the resistance measurement to a temperature equiv-
alent.
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. Th e heater is a hermetically sealed, co-
axial 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.
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 pres-
sure. 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 manifold is co nnected 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 sign al (proportional to pressure) that can be
processed and used by the EEC to control the engine.

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Figure 3: P2 / T2 and P4.9 Sensor

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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 CFDS Test
menu.
The relay box also contains the 115v Ignition relays.
FAIL SAFE POSITION:
"PROBE HEATER OFF"

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Figure 4: P2/T2 Heater Schematic

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FADEC P2/T2 Heater Test Figure 5: P2/T2 Heater Test

.%84 0!'%

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77-20 Temperature EGT Indication EGT Indicator 1. Actual EGT
Normally displayed in green.
Pulses amber up to MCT when EGT 610 C.
Pulses red when EGT 650 C.
EGT index pulsing amber must be disregarded when using TO or FLX
thrust.
2. 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:
– Engine start up, instead a amber mark is placed at 635 C
– Take Off sequence.
3. 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.
4. Red cross line
is set at the max EGT over temperature achieved during the last leg. The red
cross line will disappear thro ugh corresponding DMC’ s
– MCDU action or by the next T/ O.
Figure 6: EGT Indication

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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.
– 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.
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.
The thermocouples are connected, in parallel, to the junction box for each chan-
nel, from where two indepent signals are send to the EEC. Each signal is an av-
erage of the four probes.

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Figure 7: EGT System

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77-10 Power N1 and N2 Indication N1 Indication The low pressure rotor speed signal is used in the EEC for engine control compu-
tation and for ECAM visual display.
1. Actual N1
Displayed normally in green.
Pulses red if N1 exceeds 100%.
Pulses amber when N1 exceeds the N1 rating limit, in N1 MODE.
2. Max permissible N1
is 100 %. At 100 % a red band begins.
If the RPM exceeds 100 % index and numeric value pulses red.
3. Red cross line
is set at the max N1 over speed value achieved during the last leg.
4. White circle
N1 command corresponding to the thrust lever (angle) position (predict N1) ap-
pears 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 must 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 disp layed on E / WD in the MEMO display.
5. 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.
6. 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
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.
7. Actual N2
Digital indication normally green.
It is overbrightness and grey boxed duri ng engine start sequence up to 43 %
(starter cut out).
Turns red if N2 exceeds 100 % and a red "X" appears. The red "X" will disap-
pear through corresponding DMC’s - MCDU action or by the next T/O.
General: A failure titl e will be displayed on E / WD on the MEMO display.
If a failure occurs on any indication displayed, the analog indication is replaced by
amber crosses, the analog indicator and the marks on the circle disappear, the cir-
cle becomes amber.
ON ON ON
ON
ENG ENG
MAN STARTN1 MODE
12 12

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Figure 8: N 1 and N2 Speed Indication

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31 Indicating Max Pointer Reset (N1, N2 & EGT) Monitoring of the Relevant Display of the Engine Parameters N1, N2, EGT, and FF indications of both engines are monitored internally and ex-
ternally. The DMC compares the N1 signal received from the EEC 1 with the feed-
back signal which reflects the displa yed position of the N1 needle -
In order to grant dissimilarity with the engine 2 monitoring process the DMC com-
pares the N1 signal from the EEC 2 with the feedback signal representing the N1
digital value.
The same applies to the EGT parameters indications, but with the displayed posi-
tion of the engine 2 EGT needle and the engine 1 EGT digital feedback value.
As for the N2 and FF parameters, the DMC compares the direct signal from the
EEC with the displayed digital value.
In case of detected discrepancy, a CHECK amber message is displayed just be-
low the relevant parameter indication.
In addition the FWC’s perform an external monitoring between the feedback sig-
nals (that correspond to the displayed values and the signets that are directly re-
ceived by the FWC’s from the EEC‘s
Should a discrepancy occur, for one or more parameters, a CHECK amber mes-
sage is displayed under the relevant indication
The FWC’s generate a caution
• single chime
• master caution Light
• message on the upper ECAM DU: ENG 1 (2) N1(N2/EGT/FF) DISCREPANCY Max Pointer Reset (N1, N2 & EGT) The Max pointers for N1, N2 and EGT can be reset using the CFDS menu IN-
STRUMENTS. The menu for the EIS 1,2,3, (DMC 1,2,3) must be selected.
The memory cells which store the possible exceedance are reset either by press-
ing the GENERAL RESET line key or automatically at the next take off. Read-Out / Reset of the Engine Red Line Exceedances The DMC connected to the upper ECAM DU monitors primary parameter indica-
tions of both engines.
Should an exceedance occur, the DMC memorizes in its BITE memory the maxi-
mum value reached during the Last Flight Leg.
The values of the N1, N2, EGT red lines and transitory overlimit values are stored
in 2 independent tables, one per engine.
Read out of this engine parameter exceedance can be performed via the DMC
MCDU menu. With the function engines the parameters can be selected either for
engine 1 or 2.
A reset of the red line limits have to be performed on all 3 DMCS.
N1 Red Line Exceedance The N1 red line is represented by an arc shaped red ribbon situated at the end of
the scale.
If the N1 actual value exceeds the N1 re d line (even for a short period of time), a
small red line appears across the N1 scale and then stays at the maximum value
which has been reached.
This indicates a N1 exceedance condition. Should this condition occur, the small
red line disappears only after a new take-off or after a maintenance action through
the MCDU DMC reset. N2 Red Line Exceedance The N2 indications are displayed in digital form only. 100% N2 correspond to
14460 RPM. Should N2 actual exceeds the N2 red line value, a red cross appears
next to the digital indication. This red cross disappears only after a new take off or
a DMC reset. EGT Red Line Exceedance The EGT indications are provided in the sa me form as for the N1 indications. The
same applies to changes in color and EGT exceeding indications. However it has
to be noticed that the amber linie (EGT MAX) is variable. 635 deg. C at engine start
and 610 deg. C afterwards. Red line Limit is 650 deg.C.

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Figure 9: Max Pointer Reset

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77-10 Power N1 Indication 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 be-
low.
• One sensor provides the Engine Vibration Monitoring Unit with N1 analog sig-
nals (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 con-
nected 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 who has only one pole piece. 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.
• 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.
Figure 10: N1 Speed Sensor

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Figure 11: Fan Speed & Trim Balance Sensor, N1 Terminal Block

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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 reliabilit y 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 independ-
ing windings, two of which provide three phase frequency AC electric power to
respectively channel "A" and "B".
The third winding provides a single phas e 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 indica-
tion.
The stator and rotor are sealed from the gearbox by a shaft seal. If a shaft seal
failure occurs and the altern ator fills with engi ne oil, the alternat or will continue to
function normally.
To maintain the temperature of the dedic ated alternator at an acceptable level the
alternator incorporate an integral cooling air manifold using fan air.

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Figure 12: Engine Dedicated Alternator

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77-30 Analyzers Vibration Indication An engine vibration monitoring unit monitors the N1 and N2 levels of both engines. General The engine vibration measurement system comprises:
• one transducer on each engine with 2 piezoelectric accelerometers.
• an Engine Vibration Monitoring Unit
• two vibration indications N1 and N2.
The engine vibration system provides the following functions:
• vibration indication due to rotor unbalance via N1 and N2 slaved tracking filters
• excess vibration (above advisory level of 5 units)
• fan balancing (phase and displacement)
• shaft speed (N1 and N2)
• storage of balancing data
• initial values acquisition on request (option)
• BITE and MCDU communication
• accelerometer selection
• frequency analysis when the printer is available.
Only one accelerometer is used at a time (A or B).
The same accelerometer is not used for two successive flights. The change-
over occurs at power-up or on special request (MCDU) on the ground. Interfaces The EVMU interfaces with the ECAM and the CFDS
CFDS interfaces: Maintenance fault messages.
The N1 and N2 vibrations of the left an d right engines are displayed on the engine
and cruise pages.

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Figure 13: Vibration Indication
0.8
1.2
0.8
1.2
CFDIU
SDAC2
SDAC1
VIBRATION indications:
VIB N1
VIB N2
0.8 0.9
THE VIBRATION INDICATIONS
OF THE LP AND HP ROTORS ARE
DISPLAYED IN GREEN.
PULSING
ADVISORY
ABOVE 5
PULSING
ADVISORY
ABOVE 5
1.2 1.3
8080
140 160
VIB SENSOR A
VIB SENSOR B
Ded. Gen.
P
owersupply
115V AC

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Engine Vibration Monitoring Unit (EVMU) Description The signal conditioner is composed of:
• 2 channel modules
• 1 balancing module
• 1 data processing module
• 1 power supply module.
These modules are removable parts from the signal conditioner and are repairable
subassemblies. 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 componen t due to rotor first order unbalance.
The N1 and N2 signals are used to:
• drive the tracking filters, and
• 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:
• the 2 engine accelerometers (1 per engine)
• 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:
• SDAC1 and 2 for cockpit indication
• the CFDIU
•the DMU
• and printer (if installed) for maintenance purposes. Power Supply Module The power supply module receives the 115VAC/400Hz power. It provides the oth-
er modules with the necessary voltages.
Power Supply The EVMU is supplied with 115V/400Hz by the busbar 101XPA, 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:
• the non-volatile memory
• the timers
• the analog-to-digital converter
• the ARINC 429 transmitter and receivers
• the tacho generators.
During the power-up sequence of the BITE, the following parts of the EVMU sys-
tem are checked:
• N1 and N2 NB velocity
• unbalance data
• N1 and N2 tacho frequencies
• accelerometer signals.
Any detected failure is stor ed in the non-volatile memory with GMT, the date and
other reference parameters.

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Figure 14: EVMU Schematic

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Components The vibration transducer including two inde pendent 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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Figure 15: Vibration Sensors

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CFDS System Report / Test The Centralized Fault Data System (CFDS) enables access to the system.
The first menu sent to the MCDU is the main menu. The various functions are de-
tailed here after. Last Leg Report The EVMU sends the list of the LRUs which have been detected faulty during the
last leg. Previous Leg Report The EVMU sends the list of the LRUs which have been detected faulty during the
legs (maximum 64) previous to the last leg. The faults detected are the same as
for the last leg report. LRU Identification The EVMU sends the EVM unit part number Test The test item allows initiation of a complete c heck of the EVM system.
If no failure has been detected, the message "TEST OK" is displayed.
If any failure has been detected the failed LRU is displayed.

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Figure 16: CFDS System Report / Test EVMU

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CFDS System Report /Test Engine Unbalance Menu This menu permits for both engine, to command unbalance data storage during
next flight and the read out of the stored data. It also permits to effectuate balanc-
ing for a selected engine with both accelerometers.
Measurement of the unbalance data
The EVMU measures the position and the amplitude of the rotor unbalance of
each engine. It provides this informati on, when available, to the output bus.
Storage of unbalance data
If requested, the system can store the balancing data during the cruise phase
when stabilized conditions are reached (the actual N1speed does not fluctuate
more than plus or minus 2% during at least 30s). For every stored measurement
the stabilized conditions s hall be met once more again.
This test can be done during an engine run-up in order to obtain vibration
measurement for different N1 speeds. Refer to AMM ATA 77-32-34.
To get access again to the system report / test menu ENG, refer to AMM 31-
32-00.

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Figure 17: Unbalance Data

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CFDS System Report /Test Engine Unbalance Menu The EVMU acquired unbalance data can be cleared with the clear menu.

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Figure 18: Unbalance Data

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CFDS System Report /Test Frequency Analysis Menu This menu enables a request for a frequency analysis of the acceleration signal.
The results of the frequency analysis are sent to the printer. Frequency Analysis The EVMU can perform a frequency analysis if requested from the MCDU on the
ground. The EVMU makes the analysis at a selected N1 or N2 speed and uses
any valid accelerometer (A or B). The maximum frequency analysis is 500 Hz and
the frequency increment between adjacent spectral lines is 4 Hz. On the printer it
shown in semi-graphic form.
The frequency analysis may be performed during cruise
(flight phase = 6) or when the airc raft is on ground, engin(s) running
(flight phase = 2,3 or 9) Frequency Analysis Report When the speed and phase are those shown on the MCDU, the printer will auto-
matically print the Frequency Analysis Report.
The printer gives the vibration in "IPS Peak" (Inch per seconds), every 4 HZ and
in frequency range from 0 - 500 Hz.
For interpretation of the frequency analys is report, contact the IAE representative.

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Figure 19: Frequency Analysis

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CFDS Accelerometer Reconfig. This menu allows selection of the accelerometer A or B or the auto switch mode
alternate to be used for the next flights.
The EVMU indicates which accelerometer is in operation.

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Figure 20: Reconfig. of the Accelerometer

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78 Exhaust - V2500A

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78-00 Reverser System Introduction Description The thrust reverser comprises a fixed in ner and a movable outer (translating) as-
sembly.
The translating cowl is moved by four hydraulically operated actuators which are
pressurized by the pumps mounted on each engine.
The air is discharged through cascades.
The reverser is controlled through the FADEC system from the cockpit by a lever
hinged to the corresponding throttle control lever.
The thrust reverser system comprises:
• a hydraulic control unit (HCU)
• four actuators with internal lock for lower actuators
• three flexible shafts
• two linear variable differential transformers located on each upper actuator
• two proximity switches located on each lower actuator
• two thrust reverser cowls comprising a fixed structure and 2 translating sleeves
latched together.

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Figure 1: Thrust Reverser stowed / deployed

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Thrust Reverser System Description General The thrust reverser is actuated in resp onse to signals from the Engine Electronic
Control (EEC). Selection of either stow or deploy from the cockpit generates a sig-
nal to the engine EEC which in turn, supplie s signals to the thrust reverser hydrau-
lic control unit. Thrust Reverser Deployment Thrust reverser deployment is initiated by rearward movement of the reverser le-
ver which inputs a signal, via a dual resolver, to the EEC.
The EEC supplies a 28 volt signal to the isolation valve and directional control
valve solenoids mounted in the HCU.
The supply of the signal to the directio nal control valve solenoid is also dependent
if aircraft is on ground (weight on wheels ) and upon the closure of the aircraft per-
mission switch (T/R inhibition relay) in that line. This s witch is closed by the Throt-
tle Lever Angle signal via the spoiler/elevator computer and the Engine Interface
Unit energization of the isolation valve solenoid and the directional control valve
solenoid allows hydraulic pressure into the system. This event being relayed to the
EEC by the pressure switch mounted in the HCU.
Pressure in the lower actuators releases the locks and these events are signalled
to the EEC by the Proximity Switches (lock sensors). As the pistons move rear-
ward to deploy the reverser, the Linear Variable Differential Transformer (LVDT)
on the upper actuators monitors the movement and informs the EEC when the
translating sleeve is fully deployed, the Proximity Switches and LVDTs remain ac-
tive and the isolation valve remains energized. Thrust Reverser Stowage Stowage of reverser is initiated by forward movement of the piggyback levers
which signal this intent to the EEC. The si gnal to the directional control valve so-
lenoid is then cancelled by the EEC and permission switch, allowing pressure to
remain only in the stow side of the actuators. The pistons then move forward until
stowing is complete and the lower actuator locks are engaged after which the iso-
lation valve solenoid is de-energized and the reverser is locked in the forward
thrust mode.
During normal reverser operation the isolation valve remains energized for
a period of five seconds after the LVDTs have registered fully stowed to en-
sure full lock engagement and completion of the stow cycle.
Inadvertent Stowage/Deployment In either case the LVDT sensors would detect a movement the EEC would exe-
cute auto-restow or auto-redeploy.
This occurs when the LVDTs sense uncommanded movement greater than 10%
of actuator full travel.
When auto-restow is initiated the EEC signals the isolation valve to open.
Pressure is returned to the system and with the directional control valve in its stow
position the reverser is returned to its stowed condition.
Following auto-restow the isolation valve would remain energized for the remain-
der of the flight.
If the reverser travel exceeds 15% of its travel from the fully stowed position then
the EEC will command idle.
Following restow, full power is again obtainable.
When auto redeploy is initiated to counteract inadve rtent stow, the EEC will com-
mand the isolation valve to close and maintain it closed until forward thrust has
been reselected. This action will prevent further movement in the stow direction by
virtue of the large aerody namic loads on the translating sleeves which will normal-
ly be sufficient to deploy the reverser. If the reverser travel exceeds 22% of its trav-
el from the fully deployed position then the EEC will command idle power. T/R Components Monitored by CFDS The following components are monitored by the CFDS:
• HYDRAULIC CONTROL UNIT (HCU)
• STOW SWITCH LOWER ACTUATOR R/H
• STOW SWITCH - LOWER ACTUATOR L/H
• LVDT -THRUST REV UPPER ACTUATOR R/H (DEPLOY)
• LVDT - THRUST REV UPPER ACTUATOR L/H (DEPLOY) Thrust Reverser Independent Locking System General **ON A/C 116-199,
An independent locking system is designed to isolate the thrust reverser from the
aircraft hydraulic system. This system cons ists of thrust reverser Shut-Off Valve
(SOV) upstream of the Hydraulic Control Unit (HCU), a filter and associated
plumbing, mounting and electrical supply. The SOV is electrically actuated from
an independent signal from the SEC (Spoiler Elevator Computer), bypassing the
FADEC command circuit.

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Figure 2: Reverser System Schematic
AND
>50%
A/C on GND
from EIU

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Thrust Reverser System Cascades The cascades are designed to direct the fan air to provide the reverse thrust for
the engine.
• There are 16 cascades installed.
• The cascades are not interchangeable.

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Figure 3: Reverser Installation

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Thrust Reverser Hydraulic Supply Thrust Reverser Operation The thrust reverser is operated by aircraft hydraulic pressure.
The reverser hydraulic control unit (HCU) directs hydraulic pressure to the actua-
tors.
The EEC controls the HCU and the reverser operation. Thrust Reverser Manual Deployment Non Return Valve (By-pass) During manual deployment the non return valve must be set in the bypass position
to allow the hydraulic from the actuators to go back to return.
Access to the non return valve is gained by removing the pylon access panel on
the left hand side.

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Figure 4: Reverser Hydraulic Supply

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Thrust Reverser Independent Locking System **ON A/C 116-199, General An independent locking system is designed to isolate the thrust reverser from the
aircraft hydraulic system. This system cons ists of thrust reverser Shut-Off Valve
(SOV) upstream of the Hydraulic Control Unit (HCU), a filter and associated
plumbing, mounting and electrical supply. The SOV is electrically actuated from
an independent signal from the SEC (Spoiler Elevator Computer), bypassing the
FADEC command circuit. Component Location The SOV and the filter are located under the pylon. (Ref. Fig. 001) Component Description Shut-Off Valve The thrust reverser Shut-Off Valve (SOV) is a 3 port, two position spool valve.
It is controlled by a solenoid driven 3 port, two position normally open pilot valve.
Electrical power is supplied to the SOV through the fan electrical feeder box. Filter and Clogging Indicator It is used to filter the fluid from the ai rcraft hydraulic system. The filter is a flow-
through cartridge-type filter. The clogging indicator monitors the pressure loss
through the filter cartridge and has a pop-out indicator to signal when it is neces-
sary to replace the filter element. Two spring-loaded magnetic pistons keep the
pop out indicator in retracted position. Th e lower magnetic piston monitors the dif-
ferential between the filtered and unfiltered fluid pressure across the filter element.
As the differential pressure increases, the piston compresses its spring and
moves away from the upper magnetic piston. At a preset displacement of approx-
imately 2 mm, the upper magnetic piston spring overcomes the magnetic force
and drives the pop-out indicator from it s retracted position. The filter assembly
contains a check valve to permit the removal of the canister and the change of the
filter element with a minimum of spillage.
Figure 5: Shut-Off Valve

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Figure 6: T/R Independent Locking System (**On A/C 116-199)

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Reverser Hydraulic Control Unit Reverser Hydraulic Control Unit (HCU)
General The hydraulic control unit controls hydraulic fluid flow to the thrust reverser actua-
tors.
Control and feedback signals are exchanged with the EEC.
The HCU is mounted on the pylon over the engine centerline, just forward of the
C-duct and is accessible from the left side.
The hydraulic control unit includes the following items:
• isolation solenoid valve solenoid,
• isolation valve,
• directional control valve solenoid,
• directional control valve,
• pressure switch,
• filter and clogging indicator (pop out). Isolation Valve The solenoid operated isolation valve isol ates the thrust reverser actuation sys-
tems from the remaining hydraulic network on the engine. The isolation valve so-
lenoid is a dual coil valve solenoid connected to both channels of the EEC.
The isolation valve is in the closed positi on while the thrust reverser is in the
stowed position. Upon actuation of the thrust reverser system, the isolation valve
solenoid is energized and the isolation valve is opened. Directional Control Valve The solenoid operated directional control va lve directs high pressure hydraulic flu-
id to the correct end(s) of the actuators to either stow or deploy the translating
sleeve. The directional control valve solenoid is a dual wound solenoid connected
to both channels of the EEC. The directional control valve solenoid is energized
when the deploy command is given and provides hydraulic fluid at hydraulic pump
supply pressure to both ends of the actuators through the directional control valve
to initiate deployment of translating sleeve.
Pressure Switch The pressure switch provides signals to th e EEC to indicate when there is hydrau-
lic pressure downstream of the isolation valve. The pressure switch is closed at
pressure between 798 and 1450 psi and is opened at a minimum pressure of 798
psi. Filter and Clogging Indicator The hydraulic control unit filter is used to f ilter the fluid supply from the aircraft hy-
draulic system. The filter is a flow throug h cartridge type filter. The clogging indi-
cator monitors pressure loss through the filter cartridge and features a pop- out
indicator to signal when it is necessary to replace the filter element. Manual Lockout Lever With the manual lockout lever it is possib le to shut the hydraulic supply to the re-
verser by closing the isolation valve in the HCU. The lever can be secured in the
lockout position with a pin.(this is also a part of blocking the reverser.)
This must always be done when working on the reverser system!

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Figure 7: Hydraulic Control Unit (HCU)

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HCU in Forward Thrust Position In the initial stowed position with the reverse stow control selected in the cockpit,
the hydraulic pressure is applied to the input of the HCU. All reverser hydraulic
systems are pressurized at the return pressure as long as the aircraft is in flight
and no signal is sent to open the isolation valve solenoid.

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Figure 8: HCU Schematic

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HCU Deploy Sequence Description 1. When reverse thrust is selected in th e cockpit, the EEC ensures that deploy-
ment is permitted. In that case, the elec trical power (28VDC) is sent to the iso-
lation valve solenoid and to the directional valve solenoid.
2. When the isolation valve is opened and the directional control valve solenoid
is energized, hydraulic pressure (3000 psi) moves the directional control valve
to supply hydraulic pressure to the head end of the actuator to unlock the ac-
tuators, and then extending the actuators.
3. As soon as both lock sensors indicate unlocked for more than 0.2 seconds (in-
dicating that translating sleeves are “unl ocked sleeves” signal is sent by these
sensors to the EEC. In the cockpit an am ber REV indication is displayed in the
middle of the EPR dial or the ECAM display unit.
4. Each translating sleeve arriving at 95 pe rcent of its travel is slowed down until
completely deployed through hydraulic actuator inner restriction. This event is
indicated to EEC when both Linear variable Differential Transformers (LVTD)
detect this position. REV indication changes to green.
When the thrust reverser is in the deployed position, the isolation valve re-
mains energized to maintain the hydraulic pressure in the actuators to pre-
vent vibration. If an uncommanded stow movement is detected, the EEC will
de-energize the isolation valve. This will lead to a thrust reverser redeploy
due to aerodynamical forces on the blocker doors.

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Power Plant V2500A
78-00 Reverser System
Training Manual
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Copyright by SRTechnics
Figure 9: HCU Deploy Sequence

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Power Plant V2500A 78-00 Reverser System
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Copyright by SRTechnics
HCU Stow Sequence Description 1. When translating sleeves stowing is selected, the EEC ensures that stowing is
permitted. In that case the EEC de-energizes the directional valve solenoid.
When one translating sleeve is less th an 95% deployed, REV indication chang-
es to amber.
2. Hydraulic pressure is supplied to the rod end of the actuator, the head is con-
nected to return. A flow limiter controls hydraulic actuator piston retraction
speed.
3. When both translating sleeves are at 0% from their stowed position, they set
the proximity switches (lock sensor) wh ich send the "stowed sleeves" informa-
tion to the EEC. The REV indication disappears.
4. The actuators move until stowing is co mplete and the lower actuator locks are
engaged after which the isolation valve solenoid is de-energized and the re-
verser is locked in the forward thrust mode position.

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78-00 Reverser System
Training Manual
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Figure 10: HCU Stow Sequence

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Power Plant V2500A 78-00 Reverser System
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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 un commanded 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 H CU remains energized for the rest of the
flight. In FWD thrust, if the EEC detec ts any un commanded 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 counte ract inadvertent stow, the EEC will command
the isolation valve to close and maintain it closed until FWD thrust has been rese-
lected. The air aerodynamic load on the translating sl eeves will normally be suffi-
cient 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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78-00 Reverser System
Training Manual
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Figure 11: Auto Restow / Redeploy

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Hydraulic Actuators The actuator base is attached to a torque ring and the end of the piston is attached
to the translating sleeve. As hydraulic pres sure builds up in the actuator, the piston
extends. This moves the translating sleeve aft to the deploy position.
In the retract mode, the piston retracts which moves the translating back to the
stow position.
The Upper actuators (2) have internal LVDT.
The Lower actuators (2) have a manual unlocking handle and proximity switches. Flexshaft Installation Synchronization System
Flexible Shafts Three flexible shafts connect the four ac tuators together to synchronize the speed
with which the actuators operate and the T/R sleeves on each side of the engine.
This synchronization keeps the top and bottom of the sl eeve travelling at the same
rate so the sleeve will not tilt and jam. The synchronization also keeps the two
translating sleeves moving together so reverse pressure in the secondary air flow
is equal on both sides of the engine.
The flexible shafts are installed inside the extend (deploy) hydraulic hoses. The
shaft engages a worm gear at the base of the actuator that translates the turning
action of the actuator piston as it moves out or in.
A cross-over shaft connects the two upper actuators.
Another shaft connects the upper and lower actuators on each side.

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Training Manual
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Figure 12: Flexible Drive Shafts

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Power Plant V2500A 78-00 Reverser System
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Hydraulic Actuators Description Four actuators are used for each thrust reverser, two actuators are used for each
translating cowl.
• the lower actuators incorporate an integral lock mechanism which holds the
piston in the fully stowed position.
• the upper actuators incorporate an integral Linear Variable Directional Trans-
former (LVDT) to indicate piston positi on, and thus translating cowl position, to
the EEC.
All actuators use hydraulic snubbing at t he end of the deploy stroke to slow down
the actuators over the final part of the deploy stroke. All actuators also incorporate
the necessary deploy stroke mechanical stops. Upper Nonlocking Actuator The two upper actuators are identical and in conjunction with the two lower locking
actuators, control movement of the fan re verser translating elements in response
to hydraulic inputs from the HCU.

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Training Manual
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Figure 13: Upper Nonlocking Actuator

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Lower Locking Actuators The two lower looking actuators are identi cal and in conjunction with the two upper
actuators, control movement of the fan re verser translating elements in response
to hydraulic inputs from the hydraulic control unit (HCU).
The actuators incorporate an integral lock mechanism to hold the piston rod when
the actuator is in the fully stowed position.
The lock releases on rising hydraulic pressure when deploy is commanded via the
HCU. The lock mechanism incorporates a manual release facility and proximity
switch for electrical lock position feedback to the EEC.
Figure 14: Locking Actuator Operation

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78-00 Reverser System
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Figure 15: Lower Locking Actuator

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Power Plant V2500A 78-00 Reverser System
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Thrust Reverser Manual Deploy / Stow Manual Deploy/Stow The thrust reverser may be deployed/stowed manually for maintenance - trouble-
shooting operations.
The procedure is summarised below, the full procedure, warnings and cautions
may be found in the MM ATA 78-30.
• open and tag the CB’s listed in the MM.
• open the L. and R. hand fan cowls.
• move the thrust reverser hydraulic control unit deactivation lever to the de- ac-
tivated position and insert the lockout pin.
• disengage the locks on the two locking actuators. Insert pins to ensure locks
remain disengaged.
• position the non return valve in the by pass position (deploy only-not necessary
for stow operation).
• insert 3/8 inch square drive speed brace into external socket, push to engage
drive and rotate speed brace to extend/retract translating cowl as required.
do not exceed max. indicated torque loading.

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78-00 Reverser System
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Figure 16: Reverser Manual Operation

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Thrust Reverser Deactivation Deactivation The procedure is summarised below, the full procedure is described in the MM 78-
30-00 P.407.
• if the thrust reverser is deployed, it has to be stowed manually.
• install the lock out pin in the deactivation lever of the hydraulic control unit.
• remove the translating cowl deactivation pins (2) from their stowage and insert
them in the deactivation position.
T / R Lockout pin installation
When fully inserted in the deactivation position the pins will protude approx.
0.8" to provide visual indication of "lock out".
Figure 17: Deactivation Pin

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Figure 18: T/R Deactivation

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FADEC CFDS Reverser Test Reverser Testing via MCDU Via MCDU it is possible to operate the reverser on ground with engines OFF to
make sure the system operation is o.k.
For the TEST refer to:
MM Task 78-31-00-710-41 Operational Test of the Thrust Reverser System with
the CFDS. Description For the test hydraulic power must be switched on depending which reverser sys-
tem will be tested. (Green ENG 1, Yellow END 2).
All the test steps are wri tten on the MCDU. If the test is active the REV UNSTOW
warning appears on the engine warning display.
Movement of the throttle into the reverse idle position will deploy the reverser. Re-
turning the throttle to the FWD idle position will re stow the reverser. During the test
also the REV indication in the EPR indicator must be checked.
The actual position of the T/R is also indicated on the MCDU.
Make sure the travel ranges of the thrust reversers are clear.
For safety reasons the Test time duration is limited to 60 sec.

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Figure 19: FADEC T/R Test (NO FAULT)

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FADEC T/R Test (Fault Detected)

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FADEC T/R Test (NOT O.K.) For safety reasons the time for the test is limited.
If the Test procedure is not performed within 15 seconds (moving the Throt-
tle Lever to reverse) the test will be in terrupted and a new Test must be ini-
tiated.
The duration of the complete T/R operational Test (opening & closing) is lim-
ited to 60 seconds.
If this time is exceeded the test will be interrupte d and a new Test must be
initiated.
Figure 20: FADEC T/R Test (NOT O.K.)

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79-00-1
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79-00
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79 Oil - V2500A

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Power Plant V2500A
79-00 Oil System
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79-00 Oil System Oil System Presentation System Description The lubrication system is self-contained and thus requires no airframe supplied
components other than certain instrumentation and remote fill and drain port dis-
connectors on the oil tank. These ports ar e used to refill the oil tank promptly and
precisely by allowing the airlines to quic k-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
oil, IDG oil and fuel temperatures are maintained at acceptable levels, to the bear-
ings. 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 th e entry of each compartment to prevent
blockage by any debris / carbon flakes in the oil.
The scavenge oil is then piped, either dire ctly 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 th e oil tank. A valve allows oil to bypass
the scavenge filter when the filter differen tial pressure exceeds 20 psi. A differen-
tial pressure warning switch. set at 12 psi gives cockpit indication of impending
scavenge filter bypass.
The oil pressure is measured as a differential between the main supply line pres-
sure, upstream of any restrictors, and the pressure in the No.4 bearing compart-
ment 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 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.
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 No. 4 bearing compart-
ment to protect the seals by maintaining low pressure differentials across compart-
ment walls and minimizes air leakage into the bearing chamber.
The scavenge valve pressure transducer senses the pressure present in the scav-
enge 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 ap-
proximately 450 psi to protect downstream components.

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79-00 Oil System
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Figure 1: Oil System Schematic 01

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Power Plant V2500A
79-00 Oil System
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Lubrication System Components The lubrication system consits of four subsystems:
• the lubrication supply system
• the lubrication scavenge system
• the oil seal pressurization system
• the sump venting system.
System Monitoring and Limitations
The operation of the engine oil system may be monitored by the following flight
deck indications.
• engine oil pressure
• engine oil temperature
– MINIMUM STARTING: - 40° C
– MIN. PRIOR EXCEEDING IDLE: -10° C
– MIN. PRIOR TAKE OFF: 50° C
– MAX CONTINIOUS: 155° C
– 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
– RED LINE LIMIT: 60 PSI
– AMBER LINE LIMIT: 80 PSI
• scavenge filter clogged.
No. 4 compartment scavenge valve inoperative.

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Figure 2: Oil System Schematic 02
AB CD E FG H
AB
DE
A
A
A
A
POWER SUPPLY
1 TO 9VDC FOR 0 TO 400PSIA ^1 1PSIA
P
A
ENG 1 (ENG 2)
1 32
AUNSD
P
60PSID
DE - OILED AIR OVER - BOARD
TWO - POSITION SCAVENGE VALVE
DE - OILER
REED SWITCH OPENS
A
A
23
ENG 1 (ENG 2)
STRAIN
GAGE
A
A
A
A
POWER SUPPLY
1 TO 9VDC FOR
12
43
101PP (202PP) BUS 1 (BUS 2) 28VDC
RESTRICTOR VENT AIR / OIL OIL PRESSURE OIL SCAVENGE
32EN1 (2EN2) ENG 1 (ENG 2) OIL / PRESS 121VU212
1
2
3
4
5
BRG
COMPT
BEARING COMPARTMENT
BRG
COMPT
546213
UNSD
A
A
EXTERNAL GEAR BOX
ANGLE
GEAR BOX
RESERVE SPILL RESTRICTOR
P
A
AUNSD
132
P
P
12
PSID
SCAVENGE
FILTER
STRAINERS
CHIP
DETECTOR
SCAVENGE
PUMPS
ENG FUEL
DE-
AERATOR
FAN
AIR
FUEL COOLED OIL COOLER
AIR COOLED OIL COOLER
PRESSURE
FILTER
101PP (202PP) BUS 1 (BUS 2) 28VDC
31EN1 (1EN2) ENG 1 (ENG 2) OIL QTY 121VU212
A
A
A
A
POWER SUPPLY
1 TO 9VDC FOR 0 TO 27,5 QUARTS QUARTS
12
34
STRAINER
PRESSURE
PUMP
OIL TANK
PRESSURE TRIMMING VA LV E
PRESSURE
RELIEF VA LV E
PRESSURISING
VA LV E
J
-60 TO 250°C
N°
0,5 TO
0,6
0 TO 400PSID 1PSID
K L
TO SENSOR 4006KS
TO SENSOR 4007KS
ENG 1 (ENG 2) 4005EN TRANSDUCER-
ENG 1 (ENG 2)
4006EN INDICATOR-
4003EN XMTR - OIL PRESS
4000EN SW - LOW OIL PRESS
ENG 1 (ENG 2)4001EN SW - SCAVENGE FILTER DIFFERENTIAL PRESS
ENG 1 (ENG 2) 4002EN XMTR-OIL QTY
NO.4 BEARING PRESSURE
POS NO.4 SCAVENGE VALVE
4004EN THERMOCOUPLE SCAVENGE-OIL TEMP ENG 1 (2)
210PSID WHEN P
SUCTION
1KS1
85VU127
1KS2
86VU127
EIU 1 EIU 2
THE EIU`S COMPUTE THE
FOLLOWING VALUES:
- OIL QUANTITY
- OIL TEMPERATURE
- OIL PRESSURE
- OIL LOW PRESSURE
- NUMBER 4 BEARING
SCAVENGE PRESSURE,
AND TRANSMIT THEM
THROUGH ARINC BUS
FOR INDICATING AND
WARNING GENERATION.
EIU

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79-00 Oil System
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Oil System Bearings and Gears Lubrication Bearings and gears require oil for:
• Lubrication.
• Cooling.
• Vibration suppression. Front Bearing Compartment (Bearings no. 1, 2, 3) The following bearings and gears are located in the front bearing compartment:
• Ball bearing no.1. (LP Thrust)
• Roller bearing no.2. (LP Radial)
• Ball bearing no.3. (HP Thrust)
Description
The bearing chamber utilises 1 hydraulic seal and 2 carbon seals to contain the
oil within the bearing chamber.
The front and rear seal of the LPC booster has stage 2.5 air passing across the
seals in order to prevent oil loss.
The hydraulic seal has HPC8 air passing across the seal in order to prevent oil
loss between the two rotating shafts.
The bearings and gears are fed with oil by utilising oil je ts that liberally allow oil to
enter the bearing area.
The front bearing compartment has:
• Oil fed from the pressure pump.
• Scavenge oil recovery by the scavenge pumps.
• Vent air outlet to allow the sealin g air to escape to the de oiler.

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Figure 3: Front Bearing Compartment No.1, 2, 3

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Centre Bearing Compartment (Bearing no.4) The following bearing is located in the centre bearing compartment:
• Roller bearing no.4. (HP Radial) Description The centre bearing compartment is the hottest compartment in the engine.
In order to maintain the bearing at an acceptable operating temperature HPC12
air is taken from the engine, it is cooled by an air cooled air cooler (ACAC) and
passed back into the engine.
This cooling and sealing air is called buffer air.
The buffer cooling air supply flows around the outside of the bearing in a cooling
type jacket.
In addition to cooling the buffer air is allowed to pass across the carbon seal and
pressurise the no.4 bearing.
This bearing compartment has the following:
• Oil fed from the pressure pump.
Scavenge oil and vent air recovery by the build up of pressure in the bearing com-
partment forcing the air and oil out. The air and oil passes through the no.4 bearing
scavenge valve and then into the de oiler.

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Figure 4: Centre Bearing Compartment No.4

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Rear Bearing Compartment (Bearing no.5) The following bearing is located in the rear bearing compartment:
• Roller bearing no.5. (LP Radial) Description The rear bearing compartment has one carbon seal. This seal allows HPC8 air to
leak across the seal thus preventing oil loss from the bearing compartment.
This bearing compartment has the following:
• Oil fed from the pressure pump.
• Scavenge oil recovery by the scavenge pumps.
There is no vent outlet.
The vent air is removed from the bearing compartment along with the scavenge
oil.
• Service experience has shown it is very important to clean the number 5 bear-
ing compartment oil feed, scavenge tubes and compartment (oil jet) to prevent
the build up of coke/carbon. Ineterval 6000 FH.
• Blocked oil scavenge tubes cause oil flooding in the number 5 bearing com-
partment, which is characterized by tail pipe smoke, tail pipe fire, high oil con-
sumption and/or oil wetness in the LPT, all of which cause maintenance
disruption. Oil feed tube blockage caus es oil starvation of the number 5 bear-
ing, which can result in bearing damage.

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Figure 5: Rear Bearing Compartment No.5

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79-00 Oil System
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Oil System Components Presentation Oil Tank The tank is located on the top L. H. side of the gearbox.
The normal max-usable oil quantity in the tank is 25 US qts, the maximum oil tank
capacity is 30.5 US qts
Features:
• oil qty. transmitter
• pressure and gravity fill ports
• sight glass for level indication
• internal de-aerator
• tank pressurisation valve (6 psi)
• strainer in tank outlet
• mounting for scavenge filter and master chip detector Engine Oil Servicing Where conditions permit, the oil tank shou ld be checked and oil added, if neces-
sary, within a period of 5 to 20 minutes after engine shutdown. If the engine is
stopped for 10 hours or more, a dry motoring must be performed. This make sure
that the oil level shown in the tank is correct before oil is added. Oil Quantity Transmitter The oil quantity transmitter is located in the oil tank. Power Supply The system is supplied with 28VDC from busbar ENG 1, 101PP (DC BUS 1)
through circuit breaker 1EN1 (2EN1). Description The oil quantity transmitter is a tank prob e with a capacitor (tube portion) and an
electronic module (on the top of the transmitter) for probe energizing and signal
output. Output Voltage 1VDC to 9VDC varying linearly with the usable oil quantity from 0 to 25.8 quarts.
Figure 6: De-arator

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Figure 7: Oil Tank

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Oil Pressure Pump The pressure pump is a one stage gear type pump and supplies oil under pressure
to the engine bearings, gearbox drive and accessory drives. The oil is pumped
through a pressure filter to remove any large debris. It has a cleanable filter ele-
ment. The pressure filter housing is installed at the oil pressure pump.
The pressure filter housing incorporates a pressure priming connection and a an-
tidrain valve to prevent oil loss during removal.
The filter does not have a bypass.
The pressure pump housing incorporates the pressure filter, a cold start pressure
relief valve and a pressure pump flow limiting valve.
The pressure relief valve bypasses the pressure circuit during cold starts. Location The pump is attached to the front face of the external gearbox on the left hand
side, just below the oil tank.

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Figure 8: Pressure Pump & Filter

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Air Cooled Oil Cooler (ACOC) Location The ACOC is mounted on the engine fan case. Operation The ACOC is a additional oil cooler whic h removes heat from the engine lubricat-
ing oil using fan air and maintains the oil temperature within the specified range.
The filtered oil flows through the air cooled oil cooler before being cooled again
through the fuel cooled oil cooler.
The cooling air and the oil flows through the air / oil heat exchanger are shown be-
low. Features • oil bypass valve
• ACOC oil temperature thermocouple (for heat management system)
• modulated air flow as commanded by EEC (heat management system).
air flow regulated by air control valve.
• Fuel pressure operated actuator
• Feedback LVDT
ACOC AIR CONTROL VALVE FAIL SAFE POSITION: ”OPEN” ACOC Oil Temperature Thermocouple (refer to 73-20 Heat Management System)
The ACOC thermocouple is used for the heat management system which is con-
trolled by the EEC.

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Figure 9: ACOC Air Flow

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Fuel Cooled Oil Cooler (FCOC) Location The oil passed through the ACOC flows through the fuel cooled oil cooler (FCOC),
installed on the left hand side of the fan ca sing, before it is sent to the bearing com-
partments and both the angle and main gearboxes. Purpose • The FCOC cools the oil by using low pressure fuel.
• The FCOC also warms the low temperature fuel to the de-icing level.
• The FCOC has 2 bypass valves. Description The FCOC consits of a housing containing a removable core, a header and a fuel
filter cap. The core is composed of va cuum brazed tubes through which fuel pass-
es. Bypass Valves • One is an oil pressure relief bypass valve which diverts the excessive oil pres-
sure during engine cold start.
• The other is a fuel filter bypass valve which ensures fuel flow in the event of
fuel filter clogging.

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Figure 10: Fuel Cooled Oil Cooler

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Scavenge System The scavenge system main components are:
• chip detectors,
• five scavenge pumps with strainers,
• one common scavenge filter.
• a 2-positions scavenge valve. (Bearing No.4) Scavenge Pumps Purpose The scavenge pump returns the oil back to the oil tank. Description The scavenge pump is a five-stage gear type pump on the rear left side of the ge-
abox. Four stages of the scavenge pump are two-gear displacement pumps. The
stage used for the two main gearbox scavenge lines consists of three meshing
gears producing two inlets and outlets on opposite sides. All 6 scavenge pumps
are housed together as a single unit. The pump capacity is determined by the
width of the gears.

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Figure 11: Scavenge Pump Assembly

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Scavenge Oil Components Scavenge Filter The flows from the 6 scavenge pumps are mixed together at the scavenge filter
common filter inlet. Location The filter is mounted to the rear of the oil tank. Features • disposable filter element
• by-pass valve (opens when filter clogs)
• Differential pressure connections
• provides housing for the master magnetic chip detector
• Oil Temperature sensor Scavenge Filter Differential Pressure Switch The scavenge filter differential pressure sw itch is installed on a bracket at the top
left side of the engine fan case, near the FCOC.
Switches the ECAM OIL FILTER CLOG warning when the filter becomes blocked
(+12PSI or - 2 PSI differential press) Engine Oil Temperature The scavenge oil temperature thermocouple is located in the combined scavenge
line between the master magnetic chip detector and the scavenge filter for indica-
tion in the cockpit.
The oil temperature is sensed by a dual resi stor unit. The unit consists of a sealed,
wire-wound resistance element. This element causes a linear change in the DC
resistance when exposed to a temperature change.
Temperature measurement range:
- 60 deg. C to 250 deg. C.
The analog signal from the scavenge oil temperature thermocouple is transmitted
to the EIU. The EIU transforms this signal into a digital signal. This digital signal is
then transmitted to the lower ECAM display unit through the FWCs and the DMC.

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Figure 12: Scavenge Filter

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De-oiler Location
The de-oiler is bolted to the right hand front face of the external gearbox. Purpose • To separate the breather air/oil mixture.
• return the oil to the oil scavenge system via its own scavenge pump.
• vent the air overboard through the R/H fan cowl. Features • provides mounting for the No.4 bearing chamber scavenge valve.
• overboard vent.
• provides location for the No.4 bearing magnetic chip detector housing.
Figure 13: De-Oiler

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Figure 14: De-oiler

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No 4 Bearing Scavenge Valve Location The valve is mounted on the front face of the de-oiler casing. Purpose Maintains No.4 bearing compartment seal differential pressure to reduce over-
board loss of vent air and to prevent deteri ation of the carbon seals by restricting
the venting of the compartment air/oil mixture to the de-oiler. Type of Valve Pneumatically operated two position valve. Features • Position feed back signal to EIU (reed switch)
• uses stage 10 air as servo air
• uses value of pressure of stage 10 air as operating parameter.
• Fully open at low engine speeds (stage 10 air less than 150 PSI)
• Minimum open at high engine speed (stage 10 air more than 200 PSI) No 4 Bearing Pressure Transducer Purpose The purpose of the No.4 bearing indicating system is to monitor the correct oper-
ation of the No.4 bearing 2-position scavenge valve and to detect a No.4 bearing
carbon-seal failure.
The No.4 bearing pressure transducer is installed on the right side of the deoiler
and senses pressure at the No.4 bearing outlet line.
Linear output 1VDC to 9 VDC (0 To 300 PSIG)
Figure 15:

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Figure 16: No 4 Bearing Scavenge Valve

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79-00 Oil System
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No 4 Bearing Scavenge Valve Description Purpose
Maintains the centre bearing compartment (No 4 bearing) seal deferential pres-
sure by controlling the venting of the comp artment air/oil mixture to the de oiler.
Location
The no.4 bearing scavenge valve is located on the front of the de oiler, which is
located on the front face of the external gearbox.
Description
The no.4 bearing scavenge valve has the following features;
• Operational feed back signal to EIU.
• Uses HPC10 air as the servo air for the valve operation.
• Stage 10 air less than 150 psi the valve is at maximum open position.
• Stage 10 air more than 200 psi the valve is at minimum open position.
• Feedback to EIU of valve operation is the valve position indicator; scavenge oil
pressure sensor and Pb indication from the EEC.
The no.4 bearing scavenge valve controls the flow of the scavenge oil and vent air
by varying the size of the orifice of the valve.
This allows the scavenge oil and vent air to enter the de oiler under controlled con-
ditions. High flow When the engine is at low power the valve is at the high flow position.
Therefore the valve is fully open and the pressure differential is maintained across
the carbon seal. Low flow When the engine is at high power the valve is at the low flow position.
Therefore the valve is at the restricted flow condition and the pressure differential
is maintained across the carbon seal.
Note:
High flow at high power will cause a lo wer seal differential pressure. This will lead
to the flow of buffer air across the carbon seal to increase.
The increase flow of buffer air leads to the carbon seal drying out.
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
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 deoiler. The transducer supplies a pressure sig-
nal 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)

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Figure 17: No.4 Bearing Scavenge Valve

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Engine Oil Pressure The Oil pressure is directly linked to t he opening and closing of the No.4 Bearing
Scavenge Valve.
A closing of the valve to minimum flow (at approx. 85% N2) will restrict the return
scavenge flow to the deoiler.
This will result in a pressure drop, becau se the ratio of the pressures will change.
(the oil pressure is the differential pres sure 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.

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Figure 18: Oil Pressure Chart

.
-!). /), 02%3352%.O SCAVENGE PRESSURE 03)'
%.').% /), 4%-0%2!452%
-/4 -534 "% # &
/2 -/2%
-!8
-/0
-).
-/0
&/2 -).)-5- )$,% /), 02%3352% #/22%#4)/. 2%&%2 4/ 4(%
%.').% /0%2!4)/.!, ,)-)43

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79-00 Oil System
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Oil System Pressure Sensing General The oil pressure indicating system gives a cockpit indication of the engine oil sys-
tem working pressure.
The indication of this pressure comes elec trically from an oil pressure transmitter
on each engine.
• The oil pressure transmitter is bolted to a bracket on the top left side of the en-
gine fan case.
• The oil pressure transmitter is connected to the engine oil system by two steel
tubes. One tube connects to the oil supply tube (to the engine and gearbox
bearings). The other tube connects to the No. 4 bearing oil scavenge tube (to
the oil scavenge pump).
• Power supply: 28VDC from busbar 101PP (202PP).
• Pressure range: 0 to 400 psid.
• Output voltage: 1VDC to 9VDC varying linearly with pressure from 0 to 400
psid. Low Oil Pressure Switch The low oil pressure switch is installed on a bracket at the top left side of the en-
gine fan case, beside the oil pressure transmitter.
The oil pressure switch is connected between the oil supply tube and the No.4
bearing scavenge tube.
When the oil pressure drops below 60 psi the switch closes and a red warning is
triggert in the cockpit.
The set point range is between 45psi and 75psi.

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Figure 19: LOP Switch and Oil Press. Transmitter

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79-00 Oil System
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Magnetic Chip Detectors (M.C.D.) A total of 7 M.C.D. ‘s are us ed in the oil scavenge system.
Each bearing compartment and gearbox has its own deticated M.C.D. (two in the
case of the main gearbox) although that for the No.4 bearing is located in the de-
oiler scavenge outlet).
Magnetic Chip Detectors Location
The M.C.D. ‘s for:
• No.1, 2 and 3 bearings
• main gearbox / L/H scavenge pick-up
• angle gearbox
are located to the rear of the main gearbox on the L/H side, as shown below.
The M.C.D.‘s for:
• No.5 bearing
• De - oiler (No.4 bearing)
• Main gearbox (R/H scavenge pick up)
are located as shown below.
Do not try to install the M CD if the seal rings are no t installed. A saftey mech-
anism is installed in the MCD housing to prevent installation of the MCD if
the front seal ring is not installed.
If only the front seal ring is installed, failure of this s eal ring could result in an
in-flight shutdown of the engine because of oil leakage.

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Figure 20: Chip Detectors

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Master Chip Detector The master chip detector is located in the combined scavenge return linie, on the
scavenge filter housing.
The Master Chip Detector is accessible through its own access panel in the L/H
fan cowl.
If the master M.C.D. indicates a problem then each of the other M.C.D.‘s is in-
spected to indicate the source of the problem.
Do not try to install the M CD if the seal rings are no t installed. A saftey mech-
anism is installed in the MCD housing to prevent installation of the MCD if
the front seal ring is not installed.
If only the front seal ring is installed, failure of this s eal ring could result in an
in-flight shutdown of the engine because of oil leakage.
Figure 21: Master Chip Detector Post and Pre-Mod
3 MAGNETIC CHIP
DETECTOR HOUSING
2 SEAL RING
1 MAGNETIC CHIP
BAYONET PIN
HOLES
INDICATION
LOCKED
SLOT
HOLE
INDICATION
LOCKED
4 MAGNETIC CHIP
DETECTOR
MARK
INDICATION
LOCKED
OMNI SEAL
5 SEAL RING
MARK
INDICATION
LOCKED
6 MAGNETIC CHIP
DETECTOR
HOUSING
SBE79-0042

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Figure 22: Magntic Chip Detectors

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79-00 Oil System
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IDG Oil Servicing IDG Oil Pressure Fill A quick fill coupling situated on the transmi ssion casing enables pressure filling or
topping up the unit with oil. The oil thus introduced flows to the transmission via
the scavenge filter and external cooler circuit. This ensures:
• the priming of the external circuit
• the filtration of any oil introduced.
An internal standpipe connected to an ov erflow drain ensures a correct quantity of
oil. Oil Filter A clogged filter indication is provided by a local visual pop out indicator. The indi-
cator is installed on the anti drive end of the IDG. Oil Level Check You can read the oil level through two sight glasses located on the IDG.
One sight glass serves for the CFM 56 engi ne, the other one for the V2500 engine.
• The oil level must be at or near the linie between the yellow and green bands.
• If the oil level is not at this position, connect the overflow drain hose and drain
the oil until the correct fillin g level is reached. This will also depressurize the
IDG case.
If the overflow drainage procedure is used it can take up to 20 minutes to
complete.
Failure to observe the overflow time requirements can cause high oil level
condition resulting in elevated oper ating temperatures and damage/ discon-
nect to IDG.

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Figure 23: IDG Oil Servicing

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79-00 Oil System
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79-30 Oil Indicating System General The oil system monitoring is performed by:
• indications:
– oil quantity (quarts)
– oil temperature (degree celsius)
– oil pressure (psi)
• audio and visual warnings:
– oil low pressure (LO PRESS)
– oil filter clogging (OIL FILTER CLOG) ECAM Oil Indications 1. Oil quantity indication fl ashes green (Advisory):
– when QTY <4quarts.
2. Oil pressure indication color turns red (Warning):
– when press <60PSI.
3. Oil temperature indication flashes green (Advisory):
– when TEMP >156 deg.C
– turns amber when oil TEMP < 10 deg C or > 165 deg C.
Oil HI TEMP is displayed:
– when oil TEMP >165 deg C or 156 deg C more than 15 min.
4. Oil filter clog (White & amber) warning appears on the screen when the engine
scavenge filter is clogged.

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Figure 24: ECAM Oil Indication

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79-00 Oil System
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Oil Quantity Indicating The analog signal from the oil quantity transmitter is sent to:
• the SDAC1
• the SDAC2
• the EIU which transforms the analog signal into a digital signal.
The DMC’s process the information received as a priority order from the EIU’s
through FWC 1 and 2, SDAC1, SDAC2.
The oil quantity displayed in green on the ECAM display unit is graduated from:
• 0 to 25.8 qts in analog form (the normal max-usable oil quantity in the tank is
25 US qts, the maximum oil tank capacity is 30.5 US qts)
• 0 to 99.9 in digital form. Oil Temperature Indication The analog signal from the scavenge oil temperature thermocouple is transmitted
to the EIU. The EIU transforms this signal into a digital signal.
This digital signal is then transmitted to the lower ECAM display unit through the
FWCs and the DMC.
The ECAM oil temperature indication scale is graduated from 0 deg.C to 999
deg.C. Oil Pressure Indication The analog signal from the oil pressure transmitter is transmitted to the SDAC 1,
SDAC2 and the EIU. The EIU transforms this signal into a digital signal.
This digital signal is then transmitted to the lower ECAM display unit through the
FWCs and the DMC.
The order of priority has been defined as follows:
SDAC 1
SDAC 2
EIU.
The oil pressure indication scale is graduated from 0 - 400 PSI. Low Oil Pressure Switch The low oil pressure information is send to different aircraft systems.
Low Oil Pressure switching:
• To Steering (ATA 32-51)
• To Door Warning (ATA 52-73)
• To FWC (ATA 31-52)
• To FAC (ATA 22)
• To FMGC (ATA 22-65)
• To IDG System Control (ATA 24-21)
Low Oil Pressure Switching via EIU:
• To CIDS (ATA 23-73)
• To DFDRS INTCOM Monitoring (ATA 31-33)
• To CVR Power Supply (ATA 23-71)
• To WHC (ATA 30-42)
• To PHC (ATA 30-31)
• To FCDC (ATA 27-95)
• To Blue Main Hydraulic PWR (ATA 29-12)
• To Rain RPLNT (ATA 30-45)
Scav. Filt. Diff. Pressure Warning The Scavenge filter diff. pressure warning is send to the SDAC 1, 2 and then to
ECAM. A message will be displayed on the E/WD.

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Figure 25: Basic Schematic

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79-00 Oil System
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24-22-1
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24 Electrical Power - V2500A

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24-22 AC Main Generation
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24-22 AC Main Generation General Each engine drives its associated Integrated Drive Generator (IDG) through the
accessory gearbox. The drive speed varies according to the engine rating.
The IDG is split in two parts: the drive and the generator.
The IDG is cooled and lubricated by the IDG oil system. Generator Drive Using the variable speed input, the gene rator drive produces a constant speed on
the output shaft via a variable ratio differential.
The output constant speed is regulated at 12000 RPM. Speed Control A mechanical governor, acting on a hydraulic trim unit, contro ls the differential
gear in order to maintain the constant output speed.
The differential gear also controls the oil system pumps in order to lubricate and
cool the IDG components. Control and Monitoring AC generation is monitored by the Generator Control Unit (GCU). GEN 1 OR 2
push button Controls generator excitation via its Generator Control Unit.
For safety reasons and IDG protection, an IDG1 (or IDG2) guarded push button
allows manual disconnection of the IDG.
Reset of the system can only be performed on ground, with engines stopped, by
pulling the reset handle mounted on IDG casing. Generator The generator is a conventional 3 co-axial component brushless generator which
consists of:
• a Permanent Magnet Generator,
• a rotating diode pilot exciter,
• the generator itself.
The generator is driven at a constant speed of 12000 RPM and cooled by oil
spraying.
Generator Control Unit Supply The Permanent Magnet Generator supplies the exciter field through the Generator
Control Relay and the Generator Control Unit through a Rectifier Unit.
The Generator Control Unit (GCU) supply from the aircraft network is duplicated
(Back up supply).
The excitation control and regulation modu le keeps the voltage at the nominal val-
ue at the Point Of Regulation (POR). Generator Operation Control The generator is controlled by the corresponding generator push button. When
pressed in, if the generator speed is high enough, the generator is energized.
If the delivered parameters are correct (P ower Ready relay closed) the Generator
Line Contactor (GLC) closes to supply its network. Generator Monitoring The FAULT light comes on when any generator parameter is not correct or when
the Generator Line Contactor is open.
During the AVIONICS SMOKE procedure, the FAULT light does not come on
when the GEN1 LINE push button is set to off.
The generator failure signal is sent to SDAC 1 and 2 through the Electrical Gen-
eration Interface Unit (EGIU). When the engine is shut down, the corresponding
GEN FAULT light is on. Generator 1 To avoid complete loss of fuel pumps during the smoke procedure the GEN 1
LINE push button is released out to open the line contactor.
The generator 1 is still excited and supplies fuel pumps 1 LH and 1 RH. Generator Reset When the GEN push button is released out after a fault detection, the Generator
Control Unit is reset.

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24-22 AC Main Generation
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Figure 1: IDG Location

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Electrical Power
24-22 AC Main Generation
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Integrated Drive Generator The IDG disconnection signal is inhibit ed when the corresponding engine is not
running.
DO NOT PUSH THE IDG DISCONNECT PUSHBUTTON SWITCH FOR MORE
THAN 3 SECONDS.
THERE MUST BE AT LEAST 60 SECONDS BETWEEN TWO OPERATIONS OF
THE SWITCH.
Figure 2: IDG Description
Variable
Input
Speed
Disconnect
Mechanism
4500-
9120 RPM
Engine
A
ccessory
Gear Box
Input
Stepup
Gear
Reset
Handle
Variable
Unit
Hydraulic Trim Unit
Fixed
Unit
Mechanical
Governor
Oil System
Charge Pump
Deaerator
Scavenge
Pumps
Oil
Drive
ELEC Panel
IDG 1
GEN 1
FAULT FAULT
OFF
Generator
Control
Unit
Differential
Gear
Constant
Output Speed
12000 RPM
IDG
Generator
115V
400Hz
Permanent
Magnet
Generator
3 Phase
400 Hz
Generator
P
M
G
S
U
P
P
L
Y
T
O
G
C
U
F
I
E
L
D
E
X
C
I
T
A
T
I
O
N

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24-22 AC Main Generation
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Figure 3: IDG Oil Cooling and Warning
IDG 1
Generator
Differential
Gear
Hydraulic
Trim Unit
Governor
Drive
Oil
System
Oil in
Temp
Sensor
Fuel/Oil
Heat
Exchanger
Oil out
Temp
Sensor
Charge
Pressure
Switch
Input Speed Sensor
Generator
GCU 1
Oil Inlet Temp
T° Rise
Indication
Oil Outlet
Temp
Overheat
Temp
>185°C
Low
Pressure
Low
Speed
IDG 1
FAULT
G
P
C
U
C
F
D
I
U
E
G
I
U
S
D
A
C
1
S
D
A
C
2
E/W
Display
Master
Caut
SC
System
Display
U
S
E
R
S
Charge
PumpDe-Activator
IDG
Charge
Pressure
Switch
Cooler
Bypass
Valve
Oil in
Temp
Sensor
Oil out
Te m p
Sensor
Oil Filter
Relief Valve
Scavenge Pump
Oil Sump
Pressure
Fill PortClogging
Indicator
Fuel/Oil
Heat Exchanger
Fuel System
Disconnect Solenoid

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Figure 4: IDG Oil Level and Differential Pressure Indication DO NOT OPERATE THE IDG:
IF IT CONTAINS TOO MUCH OIL
IF IT DOES NOT CONTAIN ENOUGH OIL
IF YOU DO, YOU CAN CAUSE DAMAGE TO THE IDG.
THE OIL OVERFLOW DRAINAGE PROCEDURE CAN TAKE UP TO 20 MIN-
UTES TO COMPLETE. FAILURE TO OBSERVE THE OVERFLOW TIME RE-
QUIREMENTS CAN CAUSE HIGH IDG OIL LEVEL CONDITION RESULTING IN
ELEVATED OPERATING TEMPERATURES AND DAMAGE TO THE IDG.
Figure 5: IDG Front View
Normal
(Reset)
ΔP Indicator Butto
n
(Silver End, Red
Cylindrical Side)
Extended
OVER
FULL
ADD
OIL
ADD
OIL
1
2
3
Red
Y
ellow
Green
Red
CFM-66
A320
A
A
B
B

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24-22 AC Main Generation
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Servicing of IDG 1. If the oil level is above the line between the green and the yellow band (IDG
cold) or above the yellow band (IDG hot), oil servicing is required.
2. If the oil level is within the green band (IDG cold) or within green or yellow
bands (IDG hot), oil servicing is not required.
3. If the oil level is below the gree n band, oil servicing is required.
The yellow band corresponds to the oil thermal expansion margin.
Figure 6: Servicing of IDG
Filter Clogging
Indicator
Oil Filter
Vent Valve
(Vacuum)
Electrical
C
onnectors
Oil Level
Indicator
Oil Out
Port
Oil IN
Port
Overflow
Drain Port
Pressure
Fill Port
Case Drain
Plug
Disconnect
Reset Handle
1
2
3
Red Band
Yellow Band
Green Band
Red Band
OVER
FULL
ADDADD
OILOIL
DPI RESETS
DPI RESET LABEL
REFER TO APPROPRIATE
DOCUMENTATION FOR DETAILS OF
THE ALTERNATE DPI PROCEDURE 123 4
REMOVE
IDG
ΔP INDICATOR
BUTTON
NORMAL
(RESET)
EXTENDED

24-22-8
Training Manual
A319/A320/A321
Electrical Power
24-22 AC Main Generation
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 7: Servicing of IDG
Overflow
Drain Hose
Pressure
Fill Hose
Overflow
Drain Hose
Overflow
Drain Hose
Dust
CAP
Dus
t
CAP
Step One
Step Two
Step Three
Attach overflow drain and pressure
fill hoses.
Some oil may come out of the overflow
drain hose when it is connected.
Pump filtered oil into the IDG until
at least 1 more quart of oil comes out
the overflow drain hose.
Remove pressure fill hose only.
Install dust cap.
Allow to drain the overflow-
drain about 20 minutes!
Remove overflow drain hose when
drainage slows to drops.
Install dust cap.

24-22-9
Electrical Power
24-22 AC Main Generation
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 8: IDG Oil Filter / IDG Installation
Tension
Bolt
O-Rings
QAD
Ring
Lockwire
Bracket
Bracket
Tension
Bolt
4
3
2
1
Phase Lead Installation
Alternate
Configuration
Terminal
Block Stud
Generator
Terminal Lea
d
Assembly
Square
Washer
Terminal
Block

24-22-10
Training Manual
A319/A320/A321
Electrical Power
24-22 AC Main Generation
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
AC Main System The two engine generators provide the AC main generation. The AC main gener-
ation supplies the whole aircraft in normal flight configuration. The transfer circuit
supplies either one or the two distribution networks from any generation source:
•main,
• auxiliary,
• or ground. System Description When the two engines run in normal conditions, generator 1 and generator 2 sup-
ply their own network. Generator 1 supplies network 1, including:
• AC BUS 1,
• AC ESSENTIAL BUS,
• AC SHEDABLE ESSENTIAL BUS.
Generator 2 supplies network 2, corresponding to AC BUS 2.
Networks 1 and 2 are supplied in priority order:
• by their generator,
• by the electrical ground power unit,
• by the auxiliary generator,
• or by the other generator.
GEN1 and GEN2 push button switches, on the panel 35VU on the overhead pan-
el, control the generators 1 and 2 respectively via the GCU.
Figure 9: Main AC Distribution System

26-12-1
Power Plant V2500A
26-12
Training Manual
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Copyright by SRTechnics
26 Fire Protection - V2500A

26-12-2
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
26-12 Engine Fire and Overheat Detection Fire Detectors Each engine fire detection system consis ts of two independent loops A and B con-
nected in parallel to the Fi re Detection Unit (FDU).
Each loop comprises three fire detectors connected in parallel.
Loops A and B are connected in parallel to the Fire Detection Unit (FDU).
Each loop comprises:
• Fan fire detector
• Pylon fire detector
• Core fire detector. Fire Detection Unit (FDU) One Fire Detection Unit is provided for each engine.
The Fire Detection Unit (FDU) processes signals received from the fire detectors. Warnings The Fire Detection Unit generates signals for ECAM display, Centralized Fault
Display System utilization and cockpit local warnings.
• Fire warning signals are sent to ECAM and engine fire and start control panels.
• Loop failure warnings are sent to ECAM and Centralized Fault Display System
(CFDS). Test P/B On the engine fire panel, the TEST pushbutton permits the fire detection and the
extinguishing systems to be checked.
During the test, the SQUIB lights come on if the continuity of the squib circuit is
correct. The DISCH lights are also activated but as a lamp test.
The TEST pb checks simultaneously the integrity of the:
• Fire detection loops A and B, FDU, indications and warnings.
• Squib circuit continuity.

26-12-3
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 1: Engine Fire Warning/Extinguishing Figure 2: Fire Panel/Engine Panel

26-12-4
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 3: Fire Detector - Pylon
09,/. &)2% $%4%#4/23
7$ 7$
"! "
7$
7$
!
!
:
3%.3).' %,%-%.4 2%30/.$%2
&)2% $%4%#4/2
3%.3).' %,%-%.4 !44!#(-%.43
:
%.' 7$ %.' 7$

26-12-5
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 4: Fire Detector - Engine
#/2% :/.% &)2%
$%4%#4/23
7$
2%30/.$%2 !33%-",9
3500/24 45"%
2%#%04!#,% %,%#42)#!,
%,%#42)#!,
2%#%04!#,%
#,)0
#,)0
3%.3/2 45"%
3%.3/2 45"%
2%30/.$%2
!33%-",9
!)2 34!24%2
)$'
(9$2!5,)#
05-0
!##%33/29 :/.% &)2%
$%4%#4/23
7$
/), 4!.+
!
"
!
"
&7$

26-12-6
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Engine Fire Detection Logic General The engine Fire Detection Unit (FDU) has two channels capable of detecting any
case of engine fire and loop failure. Each channel does the same detection logic
depending on the loop A and loop B status. Fire Warning In case of a fire detected on both loops or on one loop with the other faulty, the
channels provide a fire warning to the ENGINE FIRE panel, ENGINE START con-
trol panel and ECAM displays. The FDU generates a fire warning signal if any of
the following conditions are met:
• fire on loop A and fire on loop B
• fire on loop A and fault on loop B
• fault on loop A and fire on loop B
• fault on loop A and fault on loop B within 5 seconds (both loops broken due to
a torching flame). Loop Fault Warning In case of a loop failure the FDU supplie s a loop fault warning signal to the ECAM
and Centralized Fault Display Interface Unit (CFDIU).
The FDU generates an inoperative signal if any of the following conditions are met:
• electrical failure,
• integrity failure,
• detection of a single loop FIRE during more than 16 s while the other loop is in
normal condition. Detection Fault Warning The detection fault logic is based on a dual loop failure. It agrees with a total loss
of the detection system. When the FDU generates two inoperative signals related
to loop A and loop B fault logic, the Flight Warning Computer (FWC) elaborates
the fault warning.

26-12-7
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 5: General and Fire Warning

26-12-8
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 6: Loop Fault Warning

26-12-9
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 7: Detection Fault Warning

26-12-10
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 8: Fire Detector - Schematic

26-12-11
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 9: Fire Detector - Alarm and Fault States

26-12-12
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Electrical Circuits Fire Detection Circuit The engine fire and overheat detection syst em is supplied by the electrical power
from the DC system. Each engine has two continuous loops connected in parallel
to a Fire Detection Unit (FDU). Each FDU has two identical channels: channel A
and B. Each one has its own power supply and is connected to one fire detection
loop. Fire Extinguishing Circuit The engine fire extinguishing system is supplied by electrical power from the DC
system. For each engine the system comprises one ENGINE FIREP/BSW, two
SQUIB/DISCHARGE P/BSWs and one TEST P/B located on the ENG and APU
FIRE panel; and two fire extinguishing bottles located in the engine pylon. Test Pushbutton The operational test lets th e pilot monitor and activate the fire protection system.
It can be done on the ground or in flight. Each engine has its TEST P/B which must
be pushed and held when doing the test. When the TEST P/B is pressed, the fire
warning indications are triggered on the related engine section on the ENG and
APU FIRE panel. Each TEST P/B lets the crew check the condition of the fire de-
tectors, the FDU, and the indication warn ings. It initiates the loops and squib tests.

26-12-13
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 10: Fire Detection Circuit

26-12-14
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 11: Fire Extinguishing Circuit

26-12-15
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Engine Fire Pushbutton Pushing the ENG FIRE P/BSW arms the engine fire extinguishing bottles firing
system. The SQUIB lights come on to indicate that fire bottles can be used. At the
same time the engine is isolated from the rest of the aircraft systems. Disch Pushbutton When the AGENT P/BSW is selected the extinguishing agent flows in the rigid
pipes and is immediately sprayed in the engine zones. The amber DISCH legend
comes on when the fire extinguisher bottle is completely discharged. The DISCH
alarm module sends the related signal to the Flight Warning Computer (FWC). Engine Fire Pushbutton Interfaces General During the engine fire procedure, the ENGINE FIRE P/BSW is manually released
out. This triggers several automatic sequences simplifying further crew actions
and system monitoring. Monitoring Interfaces Releasing the fire P/BSW out cancels the Continuous Repetitive Chime (CRC),
signals the action to the Flight Warning Computer (FWC) for further management
of other warnings and messages and the SQUIB light comes on, on the engine fire
control panel. Supply Interfaces Quick isolation of all systems on the rela ted engine, which could be the origin of
the fire or feed the fire, is achieved as soon as the fire P/BSW is released out.
These systems are:
•fuel,
•air,
• electric power,
• hydraulic power.
The electric supply to the Engine Interface Unit (EIU) is also disconnected.

26-12-16
Training Manual
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Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 12: General ... Supply Interfaces

26-12-17
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 13: Engine Fire Extinguisher Bottle

26-12-18
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 14: Squib and Low Pressure Switch

26-12-19
Fire Protection
26-12 Engine Fire and Overheat Detection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 15: Distribution Lines Installations

26-12-20
Training Manual
A319/A320/A321
Fire Protection
26-12 Engine Fire and Overheat Detection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics

26-99-1
Fire Protection
26-99 CFDS System Report / Test
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
26-99 CFDS System Report / Test

26-99-2
Training Manual
A319/A320/A321
Fire Protection
26-99 CFDS System Report / Test
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 1: Fire Protection-System Report/Test Figure 2: Engine or APU FDU-System Report/Test

26-99-3
Fire Protection
26-99 CFDS System Report / Test
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
FDU - Bite The tests performed by each FDU are:
• Power up test
• MCDU test
• In Operation test
The MCDU test is identical to the power up test.
Figure 3: FDU-Bite
The power up test is performed automatically as soon as the Fire Detection Unit
is electrically supplied and only if the aircraft is on ground.
At power up test, internal functions of the FDU are tested and all the detectors are
isolated from the FDU because detector conditions are simulated by the BITE.
The power up test will be initiated if the computer power supply has been de-en-
ergized for more than 200 ms. The duration of the power up test is 57 seconds.
The MCDU test is performed by maintenance crew from the MCDU with the air-
craft on ground.
The In Operation test is divided into a cyclic test and a permanent test.
The In Operation test includes:
• a cyclic test automatically performed and pr ovided that the aircraft is in flight.
During this test, the FDU internal functions are tested as well as the loop B
power supply (for engines and APU), discrepancies between LGCIU1 and
LGCIU2 inputs and the pin programming.
• a permanent test, automatically performed when the system operates. During
this test the FDU receives and analyses both detection loop signals. The FDU
continuously monitors the circuits and is capable of detecting one or more fail-
ures in both loop detection circuits.

26-99-4
Training Manual
A319/A320/A321
Fire Protection
26-99 CFDS System Report / Test
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 4: FDU Fault Messages-Examples Figure 5: MCDU Messages from FDU

30-00-1
Power Plant V2500A
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Training Manual
A319/A320/A321
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30 Ice and Rain Protection - V2500A

30-00-2
Training Manual
A319/A320/A321
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
30-00 Eng. Air Intake Ice Protection System Description Engine Air Intake Anti-Ice Air Source The air bled from the 7th stage of th e high compressor is the heat source.
A solenoid-operated shutoff valve (which is designed to fail to the open position)
provides the on-off control. The piccolo tube distributes the air within the leading
edge of the intake cowl. The spent air exhau sts via a flush duct in the aft cavity of
the intake cowl. Valve For each Engine, hot bleed air is ducted via an ”ON/OFF” valve.
The valve is pneumatically operated, electrically controlled and spring loaded
closed.
Upon energization of the solenoid, the valve will close.
In case of loss of electrical power supp ly and pneumatic air supply available, the
valve will open.
• It has a “Manual Override and Lock”. It can be blocked in the OPEN or in the
CLOSED position. Control For each engine, the ”ON/OFF” valve is controlled by a pushbutton.
Continuous ignition (A/B) is automatically activated on both engines when the
valve is opened.
The ”FAULT” light comes on during transit or in case of abnormal operation.
When the anti-ice valve is open, the zo ne controller determines the bleed air de-
mand for the Full Authority Digital Engine Control (FADEC) system.
ECAM Page
If at least one of the two engine air inta ke anti-ice systems is selected ”ON”, a mes-
sage appears in GREEN on the ”ECAM MEMO” display.
System Control ON - (PB-Switch In, Blue) The ON light comes on in blue. (valve solenoid deenergized).
ENG ANTI ICE ON is indicated on the ECAM MEMO page.
When the anti ice valve is open (valve position sw. NOT CLOSED), the zone con-
troller sends a signal to t he FADEC (ECS signal), this will:
• Modulate the Idle speed to Min. PS3 Schedule Demand for both engines.
• Switch the Cont. Ignition- ON (via EIU/EEC). OFF - (PB-Switch Out) Anti ice system is OFF (valve solenoid energized). FAULT - (PB Switch In, Amber) Fault light illuminates amber when valve not fully open. FAULT - (PB-Switch Out, Amber) Fault light illuminates amber.
The ECAM is activated
• Single chime sounds
• MASTER CAUT light ”ON”
• Warning message:
– ANTI ICE ENG 1 (2) VALVE CLSD
– ANTI ICE ENG 1 (2) VALVE OPEN.

30-00-3
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 1: Engine Nacelle A/I Architecture
12
7
OPEN POSITION
SIGNAL
CABIN ZONE
CONTROLLER
FADEC

30-00-4
Training Manual
A319/A320/A321
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
System Control Schematic Figure 2: Control Schematic

30-00-5
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Engine Anti Ice Duct and Valve
Anti-Ice Valve Deactivation refer to MEL. ATA 30. Procedure • Lock the intake anti-ice valve (1) in the open or the closed position
• Remove the lock-pin (4) from the transportation hole (5) in the valve (1).
• Use an applicable wrench on the nut (2) and move the valve to the necessary
position (open or closed).
• Hold the valve in the necessary position and install the lock-pin (4) in to the
valve locking hole (3).

30-00-6
Training Manual
A319/A320/A321
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 3: Engine Anti-Ice Duct and Valve

30-00-7
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics

30-00-8
Training Manual
A319/A320/A321
Power Plant V2500A
30-00 Eng. Air Intake Ice Protection
for training purposes only Sep08/Technical Training
Copyright by SRTechnics

36-10-1
Power Plant V2500A
36-10
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
36 Pneumatics - V2500A

36-10-2
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
36-10 General Distribution - Description and Operation General Both engine bleed air system are similar, but engine 1 only has a direct HP supply
line to the hydraulic tank pressurization system. Each system is designed to:
• Select the air source compressor stage (IP = 7th stage or HP = 10th stage air)
• Regulate the bleed air pressure to 44 PSI
• Regulate the bleed air temperature to 200° C +/- 15°C
Air is generally bled from an Intermediate Pressure (IP= 7th) stage of the engine
High Pressure (HP= 10th) compressor to minimize engine pressure losses. This
is the normal engine air-bleed configuration.
The IP stage is the 7th HP compressor stage. At low engine speeds, when the
pressure from the IP stage is insufficient, air is automatically bled from a higher
compressor stage (HP stage). This happens especially at some aircraft holding
points and during descent, with engines at idle.
The HP stage is the 10th HP compressor stage. Transfer of air bleed is achieved
by means of a pneumatically-operated butterfly valve, designated HP bleed valve
(4000HA).
When the HP bleed valve is closed, air is directly bled from the IP stage through
an IP bleed check valve (7110HM), fitted with two flappers.
When the HP bleed valve is open, the HP stage pressure is admitted into the
pneumatic ducting and closes the IP bleed check valve. Air is then bled from the
HP stage only.
HP Bleed Valve The HP bleed valve operates pneumatically and is connected by a sense line to a
pressure regulator valve (PRV 4001HA).
It is springloaded closed and starts to op en at 8 psi HP stage air pressure. It reg-
ulates the downstream pressure to 36 psi when open.
It pneumatically closes if:
• the HP stage air is above 100 psi.
• the downstream pressure from the IP stage is above 36 psi.
• the pressure regulting valve (PRV) is closed.
• the HP bleed override solenoid (4029KS) is energized (IAE-V2500 only)
(During cruise with normal bleed condition, the solenoid (4029KS) is ener-
gized. This causes the solenoid opens to ambient the HPV PRV coupling
sense line which lets the HPV close pneumatically. It avoid a permanent HP
bleed due to low IP engine pressures.)
When the HP bleed valve is closed, air is directly bled from the IP stage through
an IP bleed check valve (7110HM), fitted with two flappers.
When the HP bleed valve is open, the HP stage pressure is admitted into the
pneumatic ducting and closes the IP bleed check valve. Air is then bled from the
HP stage only. Pressure Regulator Valve (PRV) The bleed pressure regulator valve (PRV) regulates the downstream pressure to
44 psi.It is installed in the duct downstream of the IP bleed check valve and the
HP bleed valve. The bleed pressure regulator valve also operates pneumatically
but opening and closing can be controlled by the temperature limitation thermostat
(CTS 10HA) via sense line.A bleed pressure regulator valve control solenoid
(10HA). The CTS is installed in the duct downstream of the bleed air precooler ex-
changer (7150HM). The CTS controls the bleed pressure regulator valve which
controls the HP bleed valve at the same time.
The CTS reduces the PRV outlet pressure if the precooler outlet temperature ex-
ceeds 235°C. The PRV is springloaded closed and stars to open at 8 psi upstream
pressure. The PRV is pneumatically controlled to close via the CTS sense line if:
• the precooler outlet temperature is above 245°C
• a reverse flow condition exists
• the control solenoid on the CTS is energized

36-10-3
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Overpressure Safety Device The overpressure valve (OPV) protects the system in case of PRV failure. It is fully
pneumatically operated and springloaded open.The OPV starts to close >75 psi
and is fully closed > 85 psi and re-opens < 35 psi. Bleed Temperature Control The bleed temperature is regulated with a precooler. The cooling air flow from the
engine fan is controlled by the fan air va lve (FAV). The fan air valve is springload-
ed closed and pneumatically controlled by the temperature control thermostat
(CT) via sense line. The CT modulates the fan air valve to control the bleed air
temperature at 200°C +/- 15°C.
The regulated pressure transducer (Pr) sends the PRV downstream pressure sig-
nal to both BMCs (Bleed Monitoring Computer) for pressure monitoring and indi-
cation.
The transferred pressure transducer (Pt) sends the PRV upstream pressure signal
to the respective BMC for LRU failure monitoring via CFDS.
The control temperature sensor (CTS) is a dual sensor and sends the bleed air
temperature to both BMCs for monitoring and indication. System Installation This system is installed in the nacelle and pylon of each engine and includes:
• an Intermediate Pressure Bleed Check (IPC) Valve,
• a High Pressure Bleed Valve (HPV),
• a Pressure Regulator Bleed Valve (PRV) which permits or stops the bleed air
supply. It also keeps the downstream pressure to a specified limit with a Bleed
Pressure Regulated Valve Control Solenoid,
• a HP solenoid valve (4029KS) which allows the air in the sense line between
the PRV and the HPV to vent to the atmosphere. This causes the HPV to close.
• an Over-Pressure Valve (OPV) which protects the downstream pneumatic sys-
tem if the PRV does not operate,
• a bleed air precooler exchanger (air-to- air) which controls the air temperature
downstream of the system. The engine fan supplies cooling air through a Fan
Air Valve (FAV) to the precooler. A Fan-Air Valve Control Thermostat installed
downstream of the precooler controls the butterfly plate of the FAV,
• an Exchanger Outlet Temperature Sensor which monitors the temperature in
the ducts,
• two pressure transducers which monitor the pressure in the ducts,
• two Bleed Monitoring Computers (BMC1 and BMC2) which receive information
from the sensors. They monitor the system and control its operation,
• several temperature sensors (provided for regulation) which detect overtem-
perature in ducts and give temperature indication.
This system is installed in the MI D and AFT fuselage and contains:
• a crossbleed valve which isolates or connects the right and left bleed air and
distribution systems,
• an APU bleed load valve which is a part of the APU. This valve controls the
bleed air flow from the compressor of the APU when the supply system of the
engine is off or does not operate,
• an APU bleed check valve in the APU duct which protects the APU against
bleed air from the engine(s).
The Ground Supply of Compressed Air (Ref. 36-13-00) This system is installed in the lower MID fuselage on the left side and includes a
ground connector behind panel 191DB. A check valve is installed inside the
ground connector. This stops the loss of air when the ground supply unit is not
connected. The HP air is supplied to th e distribution systems through the ground
connector. Protection of the Pylon and the Nacelles This system is installed between the engine pylons and the fuselage. It has a pro-
tection function of the wing leading edge and the nacelle. If there is a major leak
in the pneumatic system, a door opens and the pressure is released.

36-10-4
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 1: Bleed Air System
TO LH
BLEED AIR
SYSTEM
F
A
N
IP
CHECK
VALVE
(7110HM)
IP
7
th
STAGE HP
10
th
STAGE
HIGH PRESSURE
BLEED VALVE
(4000HM1)
OVER
BOARD
REG
BLEED
PRESSURE
REGULATOR
VALVE
(4001HA1)
FAN AIR
VALVE
(9HA1)
CT
S
BLEED PRESSURE
REGULATED VALVE
CONTROL SOLEN OID
(10HA1)
FAN AIR VALVE
CONTROL
THERMOSTAT
(7170HM1)
TEMPERATURE LIMITING THERMOSTA T
- LIMITS TEMP TO 235 - 245 C
BY REDUCING PRV OUTLET
PRESSURE TO 17,5 PSI
- CLOSES PRV AND HPV IF
SOLENOID ENERGIZED
- PREVENTS REVERSE FLOW
BY CLOSING OF PRV AND HPV
TEMP CONTROL
T0 200 C
PRECOOLER
(7150HM)
OVER
PRESSURE
VALVE
(5HA1)
VALVE POS.
TO BMC 1
VALVE POS.
TO BMC 1
VALVE POS.
TO BMC 1+2
VALVE POS.
TO BMC 1+2
TP
TP
TP
TO HYDRAULIC
RESERVOIR
(ENG 1 ONLY)
TP
TP
VENTED IF
PRV IS CLOSED
VENT
T
T
REGULATED
PRESSURE
TRANSDUCER
(8HA1)
TRANSFER
PRESSURE
TRANSDUCER
(7HA1)
TO BMC 2
TO BMC 1
TO BMC 1
Pr
Pt
FAV
OPV
PRV
HPV
REG
T
CONTROL
TEMPERATURE
SENSOR
(6HA1)
TO BMC 2
TO BMC 1
TP
(REG 44 PSI)
HIGH PRESSURE BLEED V ALVE (HPV)
- REGULATES 36 PSI AND CLOSES IF:
- 10 t h STAGE PRESS >100PSI OR
- 7t h STAGE PRESS > 36 PSI OR
- COUPLING SENSE LINE IS VENTED
(PRV CLOSE OR HP OVERRIDE
SOLENOID ENERGIZED)
HP BLEED OVERRIDE SOLENOID
- ENERGIZED TO OPEN IF:
- WING ANTI-ICE NOT SELECTED AND
- BOTH BLEED AIR SYSTEM IN USE AND
- 9th STAGE PRESSURE (PS3) AND
> 80 PSI AND
- FLIGHT ALTITUDE > 15'000 ft
SERVO PRESSURE
TO ENGINE NACELLE
ANTI-ICE VALVE
ENG
START
SYS.
S
HP BLEED OVERRIDE
SOLENOID ENG 1
(4029KS)
VENT
CTS
COUPLING SENSE LINE

36-10-5
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 2: Bleed Air System Layout

36-10-6
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 3: Component Location, Engine and Pylon
Z400
Z400
7150HM
Bleed Air Precooler
Exchanger
9HA
5HA
Fan Air Valve
(FAV)
Overpressure
Valve (OPV)
A
A
to Starter Valve Bleed Pressure
Regulator
Valve (PRV)
Bleed Pressure Regulator
Valve Control Solenoid
4001HA
10HA
Fan Air Valve Contr
ol
Thermostat
7170HM
Wing Anti Ice
7HA
Bleed Transfer Pressure
Transducer
8HA
Bleed Regulated Pressure
Transducer
Fan Air Inlet
(Air from Fan)
4000HA
HP Bleed
Valve (HP)
7110HM
IP Bleed
Check Valve (IPC)
(CTS )
(CT)
(Pr)
(Pt)
6HA(T)
Heat Exchanger Outlet Sensor

36-10-7
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
BMC Bleed Monitoring Computer The two BMCs monitor the operation of the HP bleed valve (close/open micros-
witch signals, and transfer pressure leve l). They receive and process the signals
and transmit the information per data bus by the System Data Acquisition Concen-
trator (SDAC) to the ECAM system which generates the system display.
The indications are: pressure, temperatur e and position of the main valves (PRV,
HP Bleed Valve, CROSSBLEED and APU BLEED valve). The two BMCs signal
directly to the AIR COND overhead control panel the ENG 1 (2) BLEED FAULT
signal.
Additionally, they transmit the information to the Centralized Fault Display Inter-
face Unit (CFDIU). The CFDIU generates maintenance information which is dis-
played on the Multi Function Control Display Unit (MCDU) if the MCDU MENU is
selected. Fault Detection and Monitoring of the System The monitoring system detects failures and abnormal operation of the engine
bleed air supply system. It warns the crew and transmits the relevant information
to the upper and lower ECAM display units. Additionally the MASTER CAUT light
comes on and a single chime sounds. The system also enables abnormal opera-
tion and failure to be detect ed during flight in order to facilitate replacement on the
ground of faulty components (Line Replaceable Units, LRU).
Valves are fitted with position microswitches for monitoring.
An exchanger outlet temperature sensor monitors the precooler outlet tempera-
ture.
Two pressure transducers monitor the air pressure available in circuit.
The two BMCs monitor the electrical signals from the microswitches of the valves,
the temperature at the precooler outlet, the transferred and the regulated pres-
sures. Additionally, they monitor ambient overheat in pylons, wings and the fuse-
lage.
The two BMCs trigger a warning in case of:
• overpressure (>57 psi TD 15sec.)
• overtemperature (>257°C TD 55sec.)
• ambient overheat (Wing, Pylon or APU duct leak)
• APU air supply and PRV not closed (TD 8sec.)
The two BMCs control the closure of the PRV (during warning, engine start, APU
bleed) automatic mode of CROSSBLEED valve and APU bleed valve opening
availability.
The two BMCs monitor the correct operation of the whole system and detect ab-
normal function of an item. They send this data to the Centralized Fault Display
System (CFDS) (Maintenance Computer).
If both BMC are failed, the following messages are displayed:
On ECAM W/D: Bleed Monitoring Fault
On ECAM S/D: xx are displayed in place of temperature, pressure indication and
valve position.

36-10-8
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Figure 4: BMC Interfaces

36-10-9
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
High Pressure Bleed Valve The HP Bleed Valve is a 4 in. dia. butterf ly-type valve which operates as a shut-
off and pressure regulating valve. The HP bleed valve is normally spring-loaded
closed in the absence of upstream pressure. A minimum pressure of 8 psig is nec-
essary to open the valve.
The HP bleed valve pneumatically limits the downstream static pressure to 36 plus
or minus 3 psig. It closes fully pneumatically when the upstream static pressure
reaches 100 plus or minus 5 psig. A pneumatic sense line connects the HP bleed
valve with the bleed pressure regulator valve (PRV) in order to make sure that the
HP bleed valve will close when bleed pressure regulator valve is controlled closed.
A319 only: A solenoid is installed on a brac ket in each pylon. It is connected by a
sense line to the HPV-PRV coupling sense line. When the engine is used with the
old engine bleed air design, the Thrust Specific Fuel Consumption (TSFC) in-
creases. This is because of low IP engine pressures give permanent HP bleed.
To avoid this, during cruise with normal bleed condition:
• Wing Anti-Icing (WAI) not selected ON,
• Normal bleed configuration (2 bleeds, 2 packs),
• Ps3 more than or equal to 80 psig,
• Altitude over 15000ft,
The solenoid is energized by the Bleed monitoring computer (BMC). It opens to
ambient the HPV-PRV coupling sense line which lets HPV controlled close pneu-
matically. Regulation The HP bleed valve upstream pressure supplies chamber (1) of the regulator
through a jet to control the position of the clapper (2) and maintain constant air
pressure in the HP bleed valve actuator opening chamber.
The test intake is used for checking corr ect valve operation on the ground by di-
rectly supplying the regulator.
The HP bleed valve downstream pressure supplies the HP bleed valve actuator
closing chamber through distribution clapp er (3). Indeed when downstream pres-
sure reaches the value determined by spring preloading. Opening/Closing reduced pressure air supplies chamber (4) of the opening/closing sub assembly
though a jet to control the position of cl apper (5) against its lower seat position and
allow the HP bleed valve actuator openi ng chamber supply with reduced pressure
air.
When chamber (4) is vented to ambient the clapper (5) leaves its lower seat posi-
tion and reduced pressure air is allowed to supply the HP bleed valve actuator
closing chamber (by unseating the springloaded ball).
Figure 5: HP Bleed Valve
to PRV
4
5
3
Bleed
Air Flow
Butterfly
2
1
Test Intake
6
Opening
Chamber
Closing
Chamber
Safety
Valve
Microswitc
h

36-10-10
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Bleed Pressure Regulator Valve (PRV) (1)The PRV is a 4 in. dia. butterfly-type valve, normally spring-loaded closed in ab-
sence of upstream pressure. A minimum upstream pressure of 8 psig is necessary
to open the valve.
The PRV pneumatically regulates the downstream pressure to 44 plus or minus 3
psig.
It closes automatically in the following cases:
• overtemperature downstream of the precooler exchanger (257 +/- 3) deg.C
(60 sec. delay),
• overpressure downstream of the PRV (57 +/- 3) psig (15 sec. delay),
• ambient overheat in pylon/wing/fuselage ducts surrounding areas,
• APU bleed valve not closed,
• corresponding starter valve not closed.
It is controlled in closed position by crew action on:
• ENG FIRE pushbutton switch
• ENG BLEED pushbutton switch.
The PRV closes pneumatically in case of impending reverse flow to the engine.
The Overpressure Valve (OPV) installed downstream of the PRV protects the sys-
tem against damage if overpressure occurs.
A sense line (1/4 in. dia.) connects the PRV to the HP Bleed Valve in order to close
the HP Bleed Valve if the PRV is closed or controlled to close. The thermal fuse
installed in the valve body causes the valv e to close at 450 plus or minus 25 deg.C. Regulation The upstream pressure supplies chamber (1) of the regulator through a jet to con-
trol the position of the clapper (2) and main tain constant air pressure in the actu-
ator opening chamber.
The regulator calibration can be modified by the secondary stage of the regulator
which is pneumatically connected to the Bleed Pressure regulator valve Control
Solenoid according to the air temperature sensed downstream to the PCE. The air
pressure in chamber (3) can vary according to an air leakage controlled by the
Bleed Pressure Regulator Valve Control Solenoid. As clapper (2) remains in con-
tact with its seat (4), down stream pressure still supplies the actuator closing cham-
ber despite a reduced pressure air value lower than the nominal regulation
threshold.
The test intake is used for checking corr ect valve operation on the ground by di-
rectly supplying the regulator.
The downstream pressure supplies the actuator closing chamber through distribu-
tion clapper (4). Indeed when downstream pressure reaches the value determined
by spring preloading.
Figure 6: Bleed Pressure Regulator Valve
f
rom HP
Bleed
Valve
5
67
to Solenoid Thermostat
Closing
Chamber
Test
Intake
Butterfly
Microswitc
h
Bleed Air
Flow
13
4
2

36-10-11
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Overpressure Valve (OPV) The OPV is a 4 in. dia. butterfly-type valv e, whose operation is fully pneumatic. In
normal conditions the valve is spring-loaded open. Regulation When the upstream pressure increases and reaches 75 psig, the OPV starts to
close (pressure on the piston overcomes the spring force). This decreases the air
flow and so reduces the downstream pressure. At 85 psig upstream pressure the
OPV is fully closed, it opens again when the upstream pressure has decreased to
less than or equal to 35 psig. Safety Devices and Indications The OPV is equipped with a test port which serves to perform an "in situ" test.
A microswitch in the OPV signals the extreme open position.
Controls and Indicating
OPV operation is fully pneumatic. It ca nnot be controlled from the cockpit.
Position of the overpressure valve can be seen on the BMC current data label 066
bit 11. (Status 0 = fully open)
Figure 7: Overpressure Valve
Butterfly
Bleed Air
Pressure
T
est Port
Regulator
Assembly
Closing
Chamber
Microswitch
Pneumat
ic
Actuato
r
A

36-10-12
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Fan Air Valve (FAV) The FAV is a 5.5 in. dia. butterfly-type valv e, normally spring - loaded closed in the
absence of pressure. A minimum upstream pressure of 8 psig is necessary to
open the valve. The FAV regulates the dowstream precooler exchanger tempera-
ture to 200 plus or minus 15 deg.C (27 deg.F). Regulation A thermostat installed downstream of the precooler exchanger senses the hot air
temperature and sends to the valve a pressure signal corresponding to precooler
cooling air demand. The FAV butterfly takes a position from fully closed to fully
open to maintain the temperature value of air bled within limits. Safety Devices and Indications The FAV is equipped with a test port which serves to perform an "in situ" test.
A manual override serves to close the valve mechanically on the ground.
Two microswitches in the valve signal the full open and full closed positions of the
butterfly. A thermal fuse installed on the valve body closes the valve if the nacelle
temperature reaches 450 plus or minus 25 deg.C (45 deg.F).
Position of the fan air valve can be seen on the BMC current data label 066 bit 12.
(Status 0 = fully open)
Figure 8: Fan Air Valve
B
B
A
A
Position
Indicator
Microswitc
h
Electrica
l
Connecto
r
Thermal Fuse
Vent Screw
Test Intake
Pressure Tapping
(Motive Pressure)

36-10-13
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Fan Air Valve Control Thermostat CT (1) The fan air valve control thermostat is installed downstream of the bleed air
precooler exchanger. It controls, through the fan air valve, the engine fan cooling
airflow in order to maintain the bleed air temperature to 200 deg.C (392 deg.F)
plus or minus 15 deg.C (27 deg.F).
(2) Detailed Description
The fan air valve control thermostat contains two mains parts:
• a temperature sensing element
• a pressure regulator. Regulation When the temperature downstream of the precooler exchanger is below the re-
quired value:
• the INVAR rod valve remains on its seat
• no air flows through the pressure regulator
• the FAV remains closed.
When the temperature is over the required value differential expansion between
the INVAR rod and the stainless steel sensing tube opens the rod valve causing
the venting of the chamber A and thus allowing a pressure signal through the ther-
mostat to the opening chamber of the FAV.
Between both values the FAV butterfly has an intermediate position.
Figure 9: CT
Pressure
Reducing Valv
e
Air Venting
Regulating Probe
Filter
Precooled Air Outlet
to the Opening
Chamber of
Fan Air Valve
Chamber B
Chamber A
Clapper
CT (7170HM)

36-10-14
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Temperature Limitation CTS When the temperature downstream of bleed air precooler exchanger increases
and reaches 235 deg.C (455 deg.F), the INVAR rod in the sensing tube starts to
open the rod valve by differential dilatati on. This cause a modification of the but-
terfly position of the bleed pressure regulator valve which tends to close to reduce
the downstream pressure. If the temperature increases up to 245 deg.C (473
deg.F) the rod valve will be fu ll open and the bleed pressure limited to 17.5 psig. Closure of bleed pressure regulator valve When the solenoid is energized, its valv e moves away from its seal and vents the
bleed pressure regulator valve which closes. When the solenoid is not energized,
the solenoid valve is spring-loaded closed.
The Bleed Pressure Regulator Valve Control Solenoid has no direct effect on
the HP Bleed Valve (HPV) operation.
Figure 10: CTS Location
Figure 11: CTS
CTS (10HA
)
A
A
Filter
Solenoid
Assembly
Non Return
Assembly
Attachment
Plate
Thermostat
Assembly
N
ote: LH Side Shown
RH Side is Symmetrical
Electrical Connector
to Pressure Regulat
or
Valve (PRV)
from Precooler
Upstream
Solenoid Assembly
Solenoid
Non Return Assembly
U
pstream
P
recooler
Pressure
R
egulator
A
ssembly
Downstream
Precooler
Pressure
Sensing
Tube
Invar Rod
Air Vent
Thermostat
Assembly
to PRV
Solenoid
Valve
Plunger
Electrica
l
Connect
or

36-10-15
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Bleed Transfer Regulated Pressure Transducers Pt The pressure transducer is a piezo-resistive type cell. It senses the bleed transfer/
regulated pressure and transforms it into a proportional current voltage.
Each pressure transducer consists of:
• a measuring electronic cell
• an electrical connector
• a pressure port.
Operation The pressure to be measured is ducted to the transducer via a sense line. It acts
on the integrated strain gage of the pi ezo-resistive cell to generate an electrical
signal proportional to the pressure variatio n. The signal is transmitted to the bleed
monitoring computer.
Z
420/480
Z410/470
A
Pylon Loop
TCT
TLT
A
TPT
CTS
Wing Anti
Ice Duct
Tube
RPT
Precooler
7159HM
Transfered Pressure
Transducer
7HA1 (7HA2)
Regulated Pressure
Transducer
8HA1 (8HA2)
Electrical
Connector
Electrical
Connector
Label
Housing
Control Temp. Sens
or
6HA1 (6HA2)

36-10-16
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
Temperature Control Description and Operation To improve the operation of the engine bl eed air, the temperature limitation func-
tion of the CTS has been deleted. When 235°C (455°F) is reached, the Pressure
Regulating Valve (PRV) no longer decreases the pressure to reduce the temper-
ature downstream. To replace the CTS function, the thermal efficiency of the PCE
has been improved and the high outlet temperature threshold has been reduced.
If the PCE outlet temperature reaches 240°C (464°F), the Bleed Monitoring Com-
puter (BMC) generates a class 2 maintenance message ''AIR BLEED'' on the
ECAM STATUS page. An associated maintenance message ''THRM (Thermo-
stat), FAV (Fan Air Valve) or sense line'' can be seen on the MCDU.
Figure 12: Temperature Control

36-10-17
Pneumatic
36-10 General
Training Manual
A319/A320/A321
for training purposes only Sep08/Technical Training
Copyright by SRTechnics
CFDS MCDU Pages On the SYSTEM REPORT/TEST menu page:
AIR BLEED key is replaced by PNEU key.
On the SYSTEM REPORT/TEST AIR BLEED page:
AIR BLEED title is replaced by PEU,
PRINT key is deleted.
Figure 13:
SYSTEM REPORT / TEST
AIR BLEED
<RETURN
SYSTEM REPORT / TEST
Classic Enhanced
A319/320/321 A318
<AIR BLEED
<APU
<BMC1
<BMC2
ENG><PNEU
<APU
ENG>
TOILET>
<RETURN
SYSTEM REPORT / TEST
<BMC1
<BMC2
SYSTEM REPORT / TEST
PNEU
<RETURN
<RETURN PRINT*

36-10-18
Training Manual
A319/A320/A321
Pneumatic
36-10 General
for training purposes only Sep08/Technical Training
Copyright by SRTechnics

Study Questions IAE V2500-1
Sep08/Technical Training
Copyright by SR Technics
Study Questions
IAE V2500
Training Manual
A319/A320/A321
For training purposes only
IAE V2500-Study Questions 71-00 General 1. How is the thrust rating defined on the V2500-A1?
a) By an engine data programming plug.
b) By a cockpit selection via MCDU.
c) Automatically by FADEC control.
2. Which condition is sufficient to close the entry corridor?
a) Engine running.
b) Engine running above 80% of N2.
c) Engine running above minimum idle.
3. With the engine running, what is dangerous.
a) The inlet suction.
b) The inlet suction and the jet wake.
c) The inlet suction, the jet wake and the noise.
4. An oil access door is provided for servicing:
a) On the right side fan cowl door.
b) On the left side fan cowl door.
c) On the left side thrust reverser "C" duct door.
5. A starter valve access panel is provided for manual operation:
a) On the right side fan cowl door.
b) On the left side fan cowl door.
c) On the left side thrust reverser "C" duct door.
6. The exhaust directs rearward:
a) The fan air discharge.
b) The engine core exhaust.
c) Both the engine core exhaust and the fan air discharge
7. Where are the drains located?
a) At the rear part of the pylon.
b) At the bottom of the engine.
c) Both at the rear part of the pylo n and at the bottom of the engine.
8. Which fluids can be discharged through the drain?
a) Water, hydraulic and fuel.
b) Water and fuel.
c) Hydraulic and fuel.
72-00 Modules 9. How is the main gearbox driven?
a) By the HP rotor,
b) By the LP rotor
c) By the HP rotor via the angle gearbox.
10.The compressor stage Nº12 is:
a) The 12th stage of the HP compressor,
b) The 10th stage of the HP compressor,
c) The 7th stage of the HP compressor.
11.What bearings support the HP rotor?
a) Nº3 and Nº4,
b) Nº3, Nº4 and Nº5,
c) Nº2, Nº3 and Nº4. 73-00 Engine Fuel 12.What directly provides the closure of the PRSOV?
a) The FMU

Sep08/Technical Training
Copyright by SR Technics
Training Manual
A319/A320/A321
Study Questions
IAE V2500
Study Questions IAE V2500-2
For training purposes only
b) The EEC only
c) The MASTER lever to OFF position
13.Is same EEC channel always in command?
a) command altenates between channels on consecutive flight
b) both channels are in command
c) Always channel "A" in command and channel "B" can be selected by the
crew
14.What provides EEC with primary N2 signal?
a) frequency of the dedicated single phase winding inside the PMA
b) dedicated sensor installed on the LP stub shaft
c) Hall effect sensor installe d at the LP compressor case
15.Which sensors are located at the diffuser case?
Answer:
16.What is the primary mode setting?
a) N1
b) EGT
c) EPR
17.How is the FF signal transmitted from the EEC to the Fuel Metering Valve
(FMV)
Answer:
18.How is the maximum power selected?
a) move the thrust levers to the FLEX / MCT detent
b) automatically by the Autothrust function on Ground only
c) move the thrust levers to the maximum travel stop TOGA
19.With both engines running the Autothrust system is available?
a) Thrust levers between Idle and FLEX / MCT
b) Thrust levers between Idle and Climb
c) Thrust levers in any position
20.What power level is set if A/THR is lost
Answer:
21.The actual Fuel Flow signal to the EEC control circuit originates from?
a) FF transmitter installed downstream of the fuel metering unit
b) Calculated value between actual FF and Fuel used
c) Dual position resolver driven by the the fuel metering unit
22.In A/THR mode, will changes in the engine power cause movement of the
thrust lever?
Answer:
23.What type is the LP pump
Answer:
24.HP fuel system is protected against excessive pressure by?
Answer:
25.How is icing / waxing of the fuel prevented at low temperatures
a) heat from the engine oil transferred from the FCOC
b) low voltage heating element around the fuel pipe
c) by the Fuel Return Valve (FRV)
26.How many torque motors are installed at the FMU and name them
Answer:

Study Questions IAE V2500-3
Sep08/Technical Training
Copyright by SR Technics
Study Questions
IAE V2500
Training Manual
A319/A320/A321
For training purposes only
27.What positions the spill valve
Answer:
28.Servo fuel for the actuators is supplied from?
a) from the LP Pump
b) directly from the fuel control unit
c) dedicated pipe taken away after the fuel flow transmitter
29.The overspeed and PRSOV are opened by?
a) electrically by actuators
b) pneumatically
c) metered fuel pressure
30.Fuel samples can be taken from?
Answer:
74-00 Ignition & Starting 31.When does the PRSOV open during manual engine start?
a) when the master lever is switched on and the metered fuel pressure over-
comes the closing spring on the PRSOV.
b) as soon as the master lever switch is switches to the on position
c) simultaniously with the LP fuel valve
32. How does the starter operate?
a) Electrically.
b) Pneumatically.
c) Hydraulically.
33.In which case can the EEC abort the start?
a) Automatic start and manual start.
b) Automatic start only.
c) Wet motoring.
34.Of which type is the ignition system?
a) Low voltage, low energy.
b) Low voltage, high energy.
c) High voltage, high energy.
35.From where can igniters A and B be selected?
a) The engine panel.
b) The ECAM control panel.
c) The EEC.
36. What causes an automatic continuous relight selection?
a) IGN START selected.
b) Flame-out detected.
c) EEC failure.
37. What happens in manual mode when N2 reaches 43%?
a) MAN START pushbutton ON light goes off.
b) The EEC provides the ECAM with a message indicating that the start valve
must be closed.
c) The start valve automatically closes.
38.Where is electrical power for the Ignition coming from?
Answer:
75-00 Airflow 1. What is the copressor airflow control system used for?

Sep08/Technical Training
Copyright by SR Technics
Training Manual
A319/A320/A321
Study Questions
IAE V2500
Study Questions IAE V2500-4
For training purposes only
2. Which compressor valve is fully modulating?
Answer:
3. Do the VIGV and the VSV have a seperate actuator?
Answer:
4. VIGV and VSV are fitted at which stages of the HP compressor on A1 and A5?
Answer:
5. Where is the servo pressure for stage 7 and 10 handling bleed valve taken
from?
a) 10th stage
b) P3
c) P2.5
6. Do all stage 7 bleed valve operate in transient condition?
Answer:¨
7. Which handling bleed valve open, should a compressor surge occur
Answer:7A, 7C and 10th
8. What is the fail safe position of a ha ndling bleed valve, should the electrical
control solenoid fail?
Answer:
9. What is the VSV actuator fail safe position?
a) VSV actuator fully extended
b) VSV guid vanes fully closed
c) VSV actuator fully retracted
10.Is the fuel diverter valve a modulating valve?
Answer:
11.Is the fuel return to ta nk flow valve modulating?
Answer:
12.What flight condition inhibit the fuel return to tank flow?
Answer:
13.Air is supplied to the active clearance control (ACC) system by?
a) stage 10 air
b) stage 12 air
c) fan air
14.Make-up air valve is closed at which conditions?
Answer:

Study Questions IAE V2500-5
Sep08/Technical Training
Copyright by SR Technics
Study Questions
IAE V2500
Training Manual
A319/A320/A321
For training purposes only
15.Stage 10 HPC make-up air provides?
a) extra cooling air for stage 2 HPT and blades
b) servo pressure for handling bleed solenoid valves
c) cooling of bearing no.4 compartment only
16.What air is used for cooling the HPC 12 air in the ACAC
Answer:
17.What is the cooled HP 12 air used for?
Answer:
18.What is the fail safe position of the engine anti-ice valve?
a) closed
b) last commanded position
c) open
19.When and how is the P2/T2 probe heated?
Answer:
77-00 Indication 1. N1 speed signal is sensed from?
a) a dedicated wiring inside the PMA
b) speed probes interacting a phonic wheel in the front bearing compartment
c) speed probes sensing the rpm of the HP turbine shaft
2. Which engine parameters are diplayed on the E/WD?
Answer:
3. Which engine parameters are diplayed on engine system display?
Answer: Answer:
4. Which speed signal is supplied to the EVMU?
a) N1 only
b) N1 and N2
c) N2 only
5. How many probes interact with the phonic wheel?
Answer:
6. How many thermocouples measure the EGT
a) one
b) three
c) four
7. EGT is measured at which engine station?
Answer:

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Copyright by SR Technics
Training Manual
A319/A320/A321
Study Questions
IAE V2500
Study Questions IAE V2500-6
For training purposes only
8. Where is the P2T2 probe located?
Answer:
9. Which shaft does the air starter rotate
Answer:
10. How is the selection of the accelerometer sensors performed?
a) The EEC selects the sensor to be used.
b) The EVMU selects automatically a sensor for each flight.
c) The MCDU is used to select the sensor to be used for the next flight.
78-00 Exhaust 11.Where is the Hydraulic Isolation Valve located?
a) Inside the Hydraulic Control Unit (HCU)
b) Inside the Spoiler Elevator Computer (SEC)
c) in the Hydraulic “Tee” connection
12.When is it possible to operate the thrust reverser system?
a) On ground only,
b) On ground and in flight,
c) In flight only.
13.In flight what position and power state is the DCV (Directional Control Valve)?
14.What senses the movement of the translating cowls
Answer:
15.How is the movement of the translating cowls synchronised?
a) By the Engine Control Computer (EEC)
b) By a flexible drive cable between the actuators
c) The two halfs are not synchronized to each other
16.What is the logic of the reverse operation?
a) TLA position and aircraft on ground only.
b) TLA position, aircraft on ground and Reverse indication green.
c) TLA position, aircraft on ground, engine running (N2 condition) and both
SECs and EIU signals.
17.What does the CFDIU simulate during the MCDU reverse test?
a) The TLA.
b) The ground logic.
c) The N2 condition.
18.What is the purpose of the auto-restow system?
a) To restow the thrust reverser immediately after detection of any uncom-
manded movement.
b) To restow the thrust reverser afte r detection of any uncommanded move-
ment of at least 10% of travel.
c) To automatically restow the thrust reve rser on cancellation of a deploy com-
mand.

Study Questions IAE V2500-7
Sep08/Technical Training
Copyright by SR Technics
Study Questions
IAE V2500
Training Manual
A319/A320/A321
For training purposes only
19.Which important safety procedure should be performed prior any maintenance
work on or under the c-duct?
Answer:
79-00 Engine Oil 1. How is the oil system protected from high pump delivery pressure?
Answer:
2. Has the pressure filter a bypass valve fitted?
Answer:
3. Where does the filter clog message on the ECAM come from?
Answer:
4. How is the main cooling of the oil system achieved?
a) FCOC LP fuel system
b) ACOC
c) IDG FCOC
5. How is the oil removed from Bearing No.4 compartment?
a) by a scavenge pump located in the gearbox
b) by air pressure entering the compartment through the carbon seals
c) by suction generated of the de-oiler
6. What is the pupose of the No.4 bearing scavenge valve?
Answer:
7. Name the two temp sensors in the oil system
Answer:
8. The ECAM oil pressure indication is?
a) Pressure pump delivery minus No.4 bearing scavenge pressure
b) Pressure pump delivery pressure
c) Scavenge pump pressure
9. What happens if the oil scavenge filter is clogged?
a) Oil filter clog warning is activated in the cockpit.
b) visual clogging indicator on the filter becomes red.
c) Oil low press switch closes.
10.What protects the scavenge pumps from metallic particles?
a) A Master Chip Detector.
b) The scavenge filter.
c) Six magnetic Chip Detectors with strainers.
11.Which of the following is not directly scavenged?
a) 1, 2, 3, Bearing compartment.
b) No 4 Bearing compartment.
c) No 5 Bearing compartment.

Sep08/Technical Training
Copyright by SR Technics
Training Manual
A319/A320/A321
Study Questions
IAE V2500
Study Questions IAE V2500-8
For training purposes only

IAE-V2500 Engine for Airbus Family 319/320

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IAE-V2500 Engine for Airbus Family 319/320

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