IAE V2500 Beamer

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

Training Manual A319/A320/A321

71 Power Plant - V2500A

EASA Part 66 Cat. B1

Drain System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46 Pylon Drains . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

71-00 Introduction Engine Mark Numbers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IAE V2530-A5 Data. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Safety Zones. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Inlet Hazard Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Jet Wake Hazard Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Noise Danger Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2 4 4 6 6 6 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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forward Engine Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AFT Engine Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Change . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Removal / Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nacelle D/O. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table of Contents

32 32 34 36 36 40

72 Engine - V2500A 72-00 Engine Presentation Engine Main Bearings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Front Bearing Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO 4 Bearing Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rear Bearing Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Modules. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Module 31 (Fan Module) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inlet Cone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Front Blade Retaining Ring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Blade Removal / Installation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annulus Fillers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reposition of the Annulus Filler Seals. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72-31-11 Fan Blade Repair. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Blade Inspection / Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Repair of the Fan Disk Rear Ramp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TAP Transient Acoustic Propagation Test . . . . . . . . . . . . . . . . . . . . . . . . . Fan Trim Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . One Shot Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Module 32 Intermediate Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Module 40 HP Compressor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Combustion Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Common Nozzle Assembly (CNA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Angle and Main Gearbox. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drive Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Borescoping . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Borescope Inspection of the HP Comp. . . . . . . . . . . . . . . . . . . . . . . . . . . . Borescope Access. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10 12 14 16 18 20 22 24 26 28 28 30 30 30 35 37 38 38 42 45 47 49 51 53 55 55 57 58

73 Engine Fuel and Control - V2500A

71-00 Power Plant Drains General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

Sep08/Technical Training Copyright by SR Technics

for training purposes only

Contents - I

Training Manual A319/A320/A321 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


Table of Contents EASA Part 66 Cat. B1

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 signal.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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 - II

Training Manual A319/A320/A321 EIU Input Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CFDS System Report/Test EIU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LAST Leg Report . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRU Indentification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ground Scanning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU CFDS Discrete Outputs Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU CFDS Discrete Outputs Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU Discrete Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

84 86 87 88 88 89 90 92 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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starting Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starting Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter Air Control Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Start Air Control Valve Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cranking-Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wet Cranking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Automatic Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EEC Auto Start Abort . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Manual Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Continuous Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

75 Engine Air - V2500A


12 12 14 16 18 22 24 26 26 28 30


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


Contents - III

Training Manual A319/A320/A321 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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Hydraulic Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Manual Deployment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


2 4 6 8 8


Thrust Reverser Independent Locking System . . . . . . . . . . . . . . . . . . . . . . Component Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reverser Hydraulic Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HCU in Forward Thrust Position . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HCU Deploy Sequence Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HCU Stow Sequence Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Command Limitation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flexshaft Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic Actuators Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Nonlocking Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lower Locking Actuators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Manual Deploy / Stow. . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Deactivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC CFDS Reverser Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC T/R Test (Fault Detected). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC T/R Test (NOT O.K.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10 10 12 14 16 18 20 22 24 24 26 28 30 32 34 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 - IV

Training Manual A319/A320/A321 No 4 Bearing Pressure Transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . No 4 Bearing Scavenge Valve Description . . . . . . . . . . . . . . . . . . . . . . . . . No 4 Bearing Scavenge Valve Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System Pressure Sensing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Oil Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Magnetic Chip Detectors (M.C.D.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Master Chip Detector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDG Oil Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79-30 Oil Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ECAM Oil Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Quantity Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Temperature Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Oil Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scav. Filt. Diff. Pressure Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

26 28 28 30 32 32 34 36 38 40 40 40 42 42 42 42 42


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

24 Electrical Power - V2500A

30 Ice and Rain Protection - V2500A

24-22 AC Main Generation

30-00 Eng. Air Intake Ice Protection

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


System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Control Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Anti Ice Duct and Valve. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anti-Ice Valve Deactivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2 2 4 5 5

36 Pneumatics - V2500A 36-10 General Distribution - Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HP Bleed Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure Regulator Valve (PRV). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Temperature Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . System Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


2 2 2 2 3 3

Contents - V



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



Contents - VI


Power Plant V2500A 71-00 Introduction

71 Power Plant - V2500A 71-00 Introduction

The senior partners Rolls Royce and Pratt & Whitney assemble the engines at their respective plants in Derby, UK and Middletown Connecticut, USA

It is produced by International Aero Engines (IAE) corporation.

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 related activity.

On March 11, 1983 five of the world’s leading aerospace manufacturers signed a collaboration agreement to create, for the first 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.

IAE is responsible for the coordination of manufacture and assembly of the engines, sales, marketing, contracting 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%

Sep08/Technical Training Copyright by SRTechnics

Low Pressure Turbine Module


71-00-1



Engine Mark Numbers The V2500 engine has been designated the “V” because IAE was originally a fivenation consortium. The “V” is the Roman numeral for five. For easy identification of the present and all future variants of the V2500, International Aero Engines has introduced a new engine designation system. • All engines will retain V2500 as their generic name. • The first three characters of the full designation are V25, identifying each engine 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 simple way of grouping by name, engines with similar characteristics. The designation V2500 - D collectively describes, irrespective of thrust, all engines for McDonnell Douglas applications and V2500 - A all engines for Airbus Industrie. Similarly, V2500 - 5 describes all engines built to the -5 mechanical standard, irrespective of airframe application. The only engine exempt from this idents is the current service engine, which is already 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



71-00-2



Figure 1: Engine Mark Numbers

V2530-A5 Generic to all V2500 engines

Mechanical Standarts of engine Takeoff thrust in thousands of pounds


Airframe manufacturer - A for Airbus Industrie - D for McDonnell Douglas

MARK NUMBER

TAKEOFF THRUST (LB)

V2522 - A5

22.000

A319

V2500 - A1

25.000

A320 - 200

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


AIRCRAFT

71-00-3



Introduction

IAE V2530-A5 Data

The V2530 - A5 engine is a two spool, axial flow, high bypass ratio turbofan engine.

Fan tip diameter:

63.5 in (161cm)

Bare engine length:

126 in (320 cm)

20% of thrust is produced by the engine core.

Weight:

4942 lbs (2242 KG)

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.

Take - off thrust:

30,000 lb, flat rated to +30 deg. C

Bypass ratio:

5.44 : 1

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.

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

80% of the thrust is produced by the fan.

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.

The IAE V2530-A5 engine is flat rated. The rated thrust can be obtained for a limited 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.



71-00-4



Figure 2: V2500 Propulsion Unit



71-00-5



Safety Zones

Jet Wake Hazard Areas

Engine Inlet Hazard Areas

Warning:

Warning:

During run up operations, extreme care should be exercised when operating the engines.

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 conditions at idle and take-off thrust.

Refer to the diagram showing the jet wake hazard areas for the conditions at idle and take-off thrust.

Figure 3: Engine Inlet Hazard Areas

Noise Danger Areas Warning: Ear protection must be worn by all persons working near the engine while it operates. Loud noise from the engine can cause temporary or permanent damage to the ears.



71-00-6



Figure 4: Jet Wake Hazard Areas



71-00-7


71-00 Nacelle Access Doors & Openings

Power Plant V2500A 71-00 Nacelle Access Doors & Openings

Figure 5: Access Doors & Openings

Nacelle General The nacelle ensures airflow around the engine during its operation and also provides 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)

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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.

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Pressure relief Doors:

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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 pressure relief door in the left fan cowl.

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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 open during flight, they will not latch close again automatically. The door (doors) will be found open during ground inspections.



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71-00-8



Figure 6: Nacelle Access Doors

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71-00-9

Training Manual A319/A320/A321 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)

71-00-10



Figure 7: Power Plant Installation Presentation - General



71-00-11



Figure 8: Nacelle D/O - Air Intake Cowl



71-00-12



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



71-00-13



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.



71-00-14



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



71-00-15



Nacelle D/O Common Nozzle Assembly The Common Nozzle Assembly (CNA) mixes the exhaust gases from the secondary and primary airflows. It is bolted to the rear flange of the turbine exhaust case. 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).



71-00-16



Figure 11: Nacelle D/O - Mixed Exhaust System



71-00-17



Figure 12: Fan Cowls Opening / Closing



71-00-18



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 adjusting nut at the back of the latch keeper Figure 13: Fan Cowl Latch Adjustment



71-00-19


Thrust Reverser Cowl Doors


Figure 14: Thrust Reverser Hydraulic Control Unit (HCU)

T/R Cowling ("C-Duct") Opening / Closing (9$2!5,)# #/.42/,5.)4

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

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71-00-20



Figure 15: C-Duct Opening/Closing

THEFAIRINGMUSTBEREMOVED BEFORETHEREVERSERISDEPLOYED ANDTHEC DUCTOPENED



71-00-21



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)



71-00-22



Figure 16: Thrust Reverser Half Latches

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71-00-23



Latch Access Panel & Take Up Device An access panel, as shown below, is provided to gain access to the three BIFURCATION "C" duct latches and the "C" duct take up device (also called, Auxiliary Latch Assembly). The take up device is 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, inside the L/H "C" duct, when not in use. Red Open Flags, installed on the C-Duct indicate that the Bifurcation latches are open.



71-00-24



Figure 17: Latch Panel & Take Up Device



71-00-25



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.(take up device) If you do not do this, the front latch can get caught between the thrust reverser halves and the auxiliary latch assembly and the hook can get damaged.



71-00-26



Figure 18: Front Latch with Open Indicator



71-00-27



C - Duct Opening / Closing System On each "C" duct a single acting hydraulic actuator is provided for opening. A hydraulic hand pump must be connected to a self sealing /quick release hydraulic connection for opening. The hydraulic fluid used in the system is engine lubricating oil.



71-00-28



Figure 19: "C" Duct Opening/Closing



71-00-29



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 Cduct. 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 position. The ’C’ ducts weigh approx 578 lbs each. Serious injury to personnel working under the ’C’ ducts can occur if the ’C’ duct is suddenly released.



71-00-30



Figure 20: „C“ Duct Hold Open Struts

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71-00-31


Power Plant V2500A 71-00 Engine Mounts

71-00 Engine Mounts

The beam assembly is aligned on the aircraft pylon by two shear pins and attached with five bolts.

General

The thrust of the engine is transmitted through the thrust links, the cross beam assembly and the beam assembly to the aircraft pylon.

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.

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.

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 exhaust 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.



71-00-32



Figure 21: Forward Engine Mount



71-00-33



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.



71-00-34



Figure 22: AFT Engine Mount



71-00-35



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 positioned 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 conditions, 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.



71-00-36



Figure 23: Hold Open Braces and Adjustable Struts

THRUST REVERSER COWL DOOR



ADJUSTABLE STRUT

T/R OPENING ACTUATOR

PYLON

71-00-37



Figure 24: Engine Removal



71-00-38



Figure 25: Bootstrap Equipment



71-00-39

Training Manual A319/A320/A321 Nacelle D/O

71-00 Engine Mounts

FUEL SYSTEM • fuel supply, • fuel return to tank,

Fluid Disconnect Panel The fluid disconnect panel provides the fluid connection between engine and pylon. It is located on the left hand side of the fan case upper part. Fluid connection lines:

Power Plant V2500A

HYDRAULIC SYSTEM • hydraulic pump suction, • hydraulic pump pressure delivery, • hydraulic pump case drain.

Figure 26: Fluid Disconnect Panel



71-00-40



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



71-00-41

Training Manual A319/A320/A321 Core Electrical Junction Box


It is located in the forward mount zone.

The core electrical junction box provides the interface between the core electrical harnesses and the pylon. Figure 28: Core Electrical Junction Box



71-00-42


Power Plant V2500A 71-00 Power Plant Drains

71-00 Power Plant Drains



71-00-43



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.



71-00-44



Figure 29: Drain Mast



71-00-45



Drain System Description Fuel Drain The fuel drain lines come from engine accessories 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 the hydraulic pump. The other drains are for engine oil or accessory lubricant.



71-00-46



Figure 30: Drain System ,%&43)$% /), 4!.+ 3#500%2

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71-00-47



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.



71-00-48



Figure 31: Pylon Drains



71-00-49





71-00-50


Power Plant V2500A 72-00

72 Engine - V2500A



72-00-1


Power Plant V2500A 72-00 Engine Presentation

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 exhausts into the Common Nozzle Assembly (C.N.A.)

By-pass Flow The fan exhaust air (cold stream) entering the by-pass duct passes through the fan outlet guide vanes and flows along the by-pass duct to exhaust into the C.N.A.

Nacelle The nacelle ensures airflow around the engine during its operation and also provides 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.



72-00-2



Figure 1: Engine Components Location (L/H Side)

NOSE CONE

FAN CASE

FUEL COOLED OIL COOLER

HP COMPRESSOR SECTION

REAR ENGINE MOUNT

FUEL FILTER OIL TANK HYDRAULIC PUMP

OIL PUMP GEARBOX


COMBUSTION SECTION FUEL PUMP

COMMON NOZZLE

STAGE 7C BLEED VALVE


72-00-3



Figure 2: Engine Components Location (R/H Side) ELECTRONIC ENGINE CONTROL TURBINE SECTION

STAGE 7 BLEED VALVES

RELAY BOX

AIR COOLED OIL COOLER NO.4 BEARING COMPARTMENT AIR COOLER

LP COMPRESSOR (FAN)

STARTER

INTEGRATED DRIVE GENERATOR

DE-OILER

BLEED VALVE CONTROL VALVES



72-00-4



Figure 3: Propulsion Unit Outline 3 LP COMPRESSOR SPLITTER FAIRING 2 LP COMPRESSOR FAN BLADES

4 LP COMPRESSOR OUTLET GUIDE VANES

5 HP COMPRESSOR WING

1 AIR INLET COWL PYLON

63”

COLD STREAM

V2500-A1

V2500-A1 HOT STREAM

V2500-A5

V2500-A5

63.5”

V2500-D5

V2500-D5

COLD STREAM

6 LP COMPRESSOR STAGE 1.5, 2, 2.3 AND 2.5 BLADES

10 INLET CONE

9 INLET CONE FAIRING

7 LP COMPRESSOR STAGE 1.5 AND 2 VANES

8 LP COMPRESSOR CASE



72-00-5



STAGE NUMBERING V2530-A5 STAGES :

COMPONENT :

STAGE NUMBER :

NOTES :

1

FAN

1

ACOC,ACC,ACAC

1 2 3 4

LOW PRESSURE COMPRESSOR ( BOOSTER )

1,5 2 2,3 2.5

(Booster Stage Bleed Valve = 2.5 Bleed Ring)

3 4 5

VSV ( & IGV ) VSV VSV

7 8 9 10 11 12

CUST. BLEED, A / I, Hdlg. Bleed, Internal Cooling

1 2 3 5 6 7 8 9 10

HIGH PRESSURE COMPRESSOR

1 2 3 4 5

CUST. BLEED Hdlg. Bleed, Buffer Air, 1. HPT & NGV, Muscl Air 20 Fuel Nozzles, 2 Ignitor Plugs

COMBUSTION CHAMBER 1 2

B.S.B.V.

HIGH PRESSURE TURBINE

1 2

LOW PRESSURE TURBINE

3 4 5 6 7

ACTIVE CLEARANCE CONTROL

ACTIVE CLEARANCE CONTROL

COMMON NOZZLE



72-00-6



Figure 4: Stage Numbering



72-00-7

Training Manual A319/A320/A321 Flowpath aerodynamic stations have been established to facilitate engine performance assessment and monitoring.

Airflow Stations STA 0 1

Ambient Intake Lip of Air Intake

2

Fan Inlet

LP Compressor Exit, (HP) Compressor Inlet

2.5

3

HP Compressor Exit

4 4.5

Combustion Section Exit HP Turbine Exit

4.9

LP Turbine Exit

5 12.5

Turbine Exhaust Case Exit Fan Exit

Legend 1)

2)

Designation

CIP

T2.5

CIT

P3

Pb, CDP

T3 – – P4.9 (P5) T4.9 – P12.5

CDT – –

Parameters used for Engine Engine Trend Control Monitoring 1 ) –

Remarks

– for EPR Calculation 2 )

–

– –

– – for EPR Calculation 2 )

EGT – FEP

– –

–

– Same Sensor used as for Engine Control is also used for Trend Monitoring; – Special Sensor used for Engine Trend Monitoring

EPR / Engine Pressure Ratio =


The manufacture uses numerical station designations. The station numbers are used as subscripts when designating different temperatures and pressures, throughout the engine.

Measured Parameters often used Designation Abbreviations P0 Pamb – – P2 FIP T2 FIT P2.5


P4.9 P2


72-00-8



Figure 5: Engine Stations AERODYNAMIC STATIONS 1

STATION PT PSIA TT °C TT °F

2

12.5

2 14.7 15 59

2.5 26.2 74.1 164.4

2.5

3

12.5 24 64.6 148.3

3 438.9 540 1003.9

4

4.5

4.5 82.9 791.4 1456.5

4.9

4.9 19.9 496.6 925.9

V2500-A1



72-00-9



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 centered by a series of rod springs ("Squirrel Cage")). • Bearing 4 (Single track roller bearing).



72-00-10



Figure 6: Engine Bearings & Compartments



72-00-11



Front Bearing Compartment The bearings No. 1, 2 and 3 are located in the front bearing compartment which is at the center of the intermediate module 32. The compartment is sealed using air supported carbon seals, and oil filled (hydraulic) seal between the two shafts. This seal is supported by 8th stage air. Adequate pressure drops across the seals to ensure satisfactory sealing. This is achieved by venting the compartment, by an external tube, to the de-oiler.

Gearbox Drive The HP stubshaft, which is located axially by No 3 bearing, has at its front end a bevel drive gear which provides the drive for the main accessory gearbox, through the tower shaft. The HP stubshaft separates from the HP compressor module at the curvic coupling and remains as part of the intermediate case module.

Description The drawing below shows details of No 2 and No 3 bearings. A phonic wheel is fitted to the LP stubshaft, this interacts with speed probes to provide LP shaft speed signals (N1) to the EEC and the Engine Vibration Monitoring Unit (EVMU) which is aircraft mounted. The hydraulic seal prevents oil leakage from the compartment passing rearwards between the HP and LP shafts. No 3 bearing is hydraulically damped. The oil flow to the No. 3 bearing damper is maintained at the full oil feed pressure whilst the rest of the flow passes through a restrictor to drop the pressure. This allows larger jet diameters to facilitate flow tolerance control. The outer race is supported by a series of eighteen spring rods which allow some slight radial movement of the bearing. The bearing is centralised by 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.



72-00-12



Figure 7: Front Bearing Compartment



72-00-13



NO 4 Bearing Compartment The No 4 bearing compartment is situated in an inherently hostile, high temperature 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 (called "buffer air") is admitted to the space between the chamber and first heat shield. The 12th stage air is cooled by fan air via the buffer air cooler, located on the rear left hand side of the engine. The buffer air is exhausted from the cooling spaces close to the upstream side of the carbon seals, creating an area of cooler air from which the seal leakage is obtained. This results in an acceptable temperature of the air leaking into the bearing compartment. Buffer air flow rates are controlled by restrictors at the outlet from the cooling passages. The bearing compartment internal pressure level is determined by the area of the variable scavenge valve. (called No 4 bearing scavenge valve and described in the oil system). This valve acts as a variable restrictor in the compartment vent / scavenge line. A drain hole is provided to indicate a possible leckage at the No 4 bearing compartment. It is located in the exhaust at 5 o clock position (aft looking forward) 12th stage air cooler (BUFFER AIR) The No. 4 bearing compartment air cooler is installed on the turbine casing. The exchanger is held by its coolant air duct flanges.



72-00-14



Figure 8: No.4 Bearing Compartment

DIFFUSER CASE REAR INNER FLANGE

FRONT WALL

HEATSHIELD

BEARING SUPPORT ASSEMBLY

COOLING DUCT

REAR WALL 12 TH STAGE P COMPRESSOR AIR

REAR SEAL FRONT SEAL

FRONT SEAL SEAT

No. 4 BEARING RING LOCK AND NUT

REAR SEAL SEAT



72-00-15

Training Manual A319/A320/A321 Rear Bearing Compartment


Figure 9: Bearing No.5 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 bolts. 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.



72-00-16



Figure 10: Rear Bearing Compartment

TEC

STAGE 8 AIR REAR THERMAL BLANKET

COMPARTMENT COVER BLIND CAP

LPT SHAFT


BLIND CAP

PACKING


72-00-17



Engine Modules

High Pressure Compressor

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 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.

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 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 clearances and to provide structural cooling.

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 combustor that provide low velocity shroud air to feed the combustor liners and to minimize performance losses.

High Pressure Turbine

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.

The module numbers refer to the ATA chapter reference for that module.

Accessory Drive Gearbox The accessory drive gearbox provides shaft horse power to drive engine and aircraft accessories. These include fuel, oil and hydraulic pressure pumps and electrical power generators for the EEC and for the aircraft. The gearbox also includes provision for a starter which is used to drive the N2 shaft for engine starting.

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 assembly, high spool stubshaft and the station 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 bearings for the low spool and high spool stubshafts.



72-00-18



Figure 11: Engine Modules



72-00-19

Training Manual A319/A320/A321 Module 31 (Fan Module)


In order to minimize the leakage of air between the fillers and the aerofoils, there is a rubber seal bonded to each side of each filler.

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

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.

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 located by the front and rear blade retaining rings. Each filler is secured to the front blade retaining ring by a bolt.



72-00-20



Figure 12: LP Compressor (Fan)



72-00-21



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-balanced 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 respective weights must be marked before removal, so as to ensure they are refitted to the same position.A special tool is used to remove the Inlet Cone to prevent it from damage as shown below.



72-00-22



Figure 13: Inlet Cone Removal ).,%4#/.% &,!.'% &2/.4",!$%2%4!).).' 2).'&,!.'%

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72-00-23



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 retaining ring and permits the individual securing 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 blade 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 identified. For this reason it is important to identify and make a note of the original blade and annulus filler positions prior to their removal. When the Nose Cone is fitted, it is possible to identify the positions of blades numbers 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 counterclockwise 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



72-00-24



Figure 14: Front Blade Retaining Ring



72-00-25

Training Manual A319/A320/A321 Fan Blade Removal / Installation


The moment weight of the fan blade is written on the root surface

Removal Figure 15: Fan Blade Profile

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 retaining ring and screws into each of the 22 annulus fillers. Both rings of bolts must be removed before attempting to remove 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 purpose.

HONEYCOMB CORE

CONVEX SKIN

CONCAVE SKIN

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 blade 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.



72-00-26



Figure 16: Fan Blade Removal / Installation



72-00-27


Annulus Fillers


Make sure the plastic strip has a smooth surface and edges. If you use a strip with a rough edge surface or edges, damage to the seal can occur.

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

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.

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 filler 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 installed, it is possible that when the Front Retaining Ring is subsequently torque tightened in place onto the Fan Disk, it may result in the deformation and displacement of the Rear Retaining Ring. This could cause it to come into contact 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 displaced 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.



72-00-28



Figure 17: Annulus Filler



72-00-29


72-31-11 Fan Blade Repair Fan Blade Inspection / Repair Before any repair is carried out, reference must be made to the AMM Chapter 7231-11 Page Block 800. Repair Damage on the Low Pressure Compressor (LPC) Fan Blades by Local Material Removal

2. 3. 4. 5.


Wash the repaired area with a cloth soacked in the solution. Use a cloth soaked in clean cold water until the area is fully cleaned. If necessary repeat steps (2) and (3). 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 • • • •

•

•

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 DAMAGE OR WEAR IS MONITORED, TO GET BACK LP COMPRESSOR EFFICIENCY 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 surface 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 removal.

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).


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 minimum 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 completely removed and the surface finish is made the same as the adjacent material. 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.


72-00-30



Figure 18: Fan Blade Repair Limits

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72-00-31



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72-00-32

Training Manual A319/A320/A321 E. Examine the LP Compressor Fan Blades


Figure 20: LP Compressor Fan Blade

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).

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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). !&4%2

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72-00-33



Figure 21: Fan Blade leading and trailing edge limits TRAILING EDGE

AREA F

LEADING EDGE

0.75in. (19.05mm)

Bt 2.00in. (50.80mm)

At

Ct

11.500in. (292.10mm)

ANNULUS

LINE

1.50in.

Br Ar

(38.10mm)

3.00in. (76.20mm)

Cr

3.00in. (76.20mm)

MAXIMUM SERVICEABLE LIMITS FOR SURFACE DAMAGE DEPTH ON CONVEX AND CONCAVE SURFACES.

Ar Br F


0.008in. (0.20mm) 0.008in. (0.20mm) 0.008in. (0.20mm)

At Bt Ct


0.025in. (0.63mm) 0.025in. (0.63mm) 0.008in. (0.20mm)

72-00-34



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 ‘dove-tail’ slot in the fan disk. Great care must be taken to inspect the fan disk and the security of the rear ramps, as they play an important role in providing a firm fixing and support for the individual fan blades. Should it be discovered that a rear ramp has become separated from the disk it must be refitted/replaced and a full description 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. V02019) 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 ensure adequate thickness of adhesive is maintained between the mating surfaces. 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.



72-00-35



Figure 22: Fan Disk Rear Ramp

REAR FACE OF REAR RAMP REAR FACE OF DISK DIRECTION OF INSERTION FOR REAR RAMP

0.020 (0,50) 0.000 (0,00)

DISK

B

REAR RAMP

B

SECTION

A-A

MASKING TAPE

AB FLAT BOTTOM PORTION 0,15 ( 0.006 0.004(0,10))

GLUE LINE THICKNESS TO BE CONTROLLED BY FOUR PIPS ON REAR RAMP

FOUR PIPS VIEW ON

D

SECTION

C-C


BOND REAR RAMP WHERE MARKED ALL DIMENSIONS ARE IN IN. (MM)


72-00-36

Training Manual A319/A320/A321 TAP Transient Acoustic Propagation Test


Figure 23: TAP Test

E.g. after ingestion of birds, foreign objects or slush a TAP test of the LP Compressor 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 pea sized) of ultrasonic couplant (Material No. V06-148) to the lower convex airfoil adjacent to the annulus filler line. 2. Attach the probe to the fan blade. 3. Press the ON switch.

ACOUSTIC EMISSION PROBE

ANNULUS FILLER LINE

•

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 PROTECTION IS INSTALLED WHEN YOU READ THE VALUE. THE PROTECTION 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 approximately four seconds. If the display shows the message COUPLING FAILURE, apply more ultrasonic couplant (Material No. V06-148) and do the inspection again.

ACOUSTIC EMISSION PROBE

EXEC MENU ON OFF

5. If the TAP-test display value is more than 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.



72-00-37


Fan Trim Balance


Fan Trim Balance with the EVMU (One Shot Method)

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).

This procedure can be used for consecutive fan trim balances if necessary. If consecutive fan trim balances with this method do not give significant results, carryout a fan trim balance with the ‘Trial Weight’ method.

The one shot method allows balancing of the fan with fewer engine ground runs required and has proved itself effective in service use.

This information is contained in the EVMU and by accessing the EVMU Engine Unbalance menu, it is possible to establish the necessary adjustments required to eliminate out of balance situations.

If necessary a Vibration survey (Test No 8) may be performed to obtain the vibration characteristics of the engine.

Note:

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 either automatic or for selected speeds and on the ground may be manually selected.

Prior to carrying out any adjustments, 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 physically verify the position and part number of the balance weights already installed onto the front blade-retaining ring. Figure 24: Moment Weights

Best results are obtained from data in the 80-90% N1 speed range with 85% N1 being the best single speed point, for ground running an average of correction.

14 LOCATING HOLE (22 OFF) PUSHER BOLT

One Shot Method The following procedure may be used to trim balance an engine fan whilst mounted 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.

11 BOLT (22 OFF)

2 LOCATING PIN (22 OFF)

3 SHOULDER HEADLESS PIN (3 OFF) 13 THREADED HOLE (6 OFF) 10 BOLT (36 OFF) 12 LOCATING HOLE (3 OFF) 4 FAN DISK PULLER BOLT 5 BOLT

9 BALANCE WEIGHT

Mass Coefficient (K) value by which the amplitude must be multiplied to give correction mass required or a given speed 6 FRONT BLADE RETAINING RING

8 NUT

7 BALANCE WEIGHT



72-00-38



Figure 25: Moment Weights



72-00-39



Figure 26: Trim Balance CFDS Procedure

SYSTEM REPORT / TEST ENG EIU 2>

FADEC 1B>

FADEC 2B>

CFDS

EVMU ENG1 TRIM BALANCE METHOD SELECTION

SYSTEM REPORT / TEST

EVMU ENGINE UNBALANCE

EVMU ENG1 CURRENT VIB DATA

F/CTL>

FUEL>

FLIGHT DATA

ENG 2>

ICE & RAIN>

LOAD

ENG 2>

INST>

L/G> NAV>

ENG> TOILETS>

TRIM

ENG 2> PRINT*

EVMU
N1 RPM 3041


PRINT*

PRINT*


PRINT*

EVMU ENG1 FLIGHT DATA ACCLRM A 160408 MIL DEG D/M 0.2 +0 03/01 NO ACQUISITION 0.5 +230 03/01 0.5 +236 03/01 0.6 +189 03/01 CONT>

PRINT*

72-00-40



Figure 27:

EVMU ENG1 ONE SHOT TRIM BAL INFLUENCE COEFF SELECT

EVMU ENG1 INST CURRENT WGHTS 36 BOLT FLANGE

<33 / 5AXXXX

34 / 5AXXXX>

<35 / N-A

36 / 5AXXXX>

CONT> PRINT*

EVMU ENG1 INST CURRENT WGHTS

<36 BOLT FLANGE

<24 BOLT FLANGE

<24 BOLT FLANGE

BALANCE SOLUTION CONT> PRINT*


PRINT*

EVMU ENG1 WEIGHTS TO CHANGE SOL XX XXQZIN / XXXDEG

BALANCE SOLUTION CONT>
PRINT*


02 / 5AXXXX>

<19 / 5AXXXX

20 / 5AXXXX>

<03 / N-A

04 / 5AXXXX>

<21 / N-A

22 / 5AXXXX>

<05 / 5AXXXX

06 / 5AXXXX>

<23 / N-A

24 / 5AXXXX>

<07 / 5AXXXX

08 / 5AXXXX> PRINT*

BALANCE SOLUTION CONT>
36 BOLT FLANGE 24 BOLT 02 / 5A0127 01 / 5A0103 03 / 5A0127 03 / REMOVE 10 / 5A0107 04 / 5A0127

<01 / 5AXXXX


PRINT*


<36 BOLT FLANGE

<36 BOLT FLANGE <24 BOLT FLANGE


CONFIG UPDATE>
PRINT*

If RETURN is pressed, this will quit the menu without storing the new bolt configuration.

If CONFIG UPDATE is pressed, the EVMU automatically updates the current weights configuration, relying on the assumption that the proposed weights have been installed.

CONT>
PRINT*


72-00-41



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.



72-00-42



Figure 28: Booster Stage Bleed Valve

AA

AA MID ARM OPEN FAN FRAME

CLOSE BLEED VALVE

BLEED VALVE

OPEN

CLOSE

UPPER POWER ARM

ACTUATING ROD LOWER POWER ARM



72-00-43



Figure 29: Fan Case Section

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72-00-44

Training Manual A319/A320/A321 Module 40 HP Compressor


Figure 30: HP Compressor

The HP compressor has 10 stages. It utilises variable inlet guide vanes at the inlet to stage 3 and variable stator vanes at stages 3, 4 and 5 The front casing, which houses stages 3 to 6, is made in two halves which bolt together along horizontal flanges. It is bolted to the intermediate casing (module 32) at the front and to the outer casing at the rear. The rear compressor casing has inner and outer casings as shown. Flanges on the inner case form annular manifolds which provide 7 and 10 stage air offtakes. On the V2500-A1 the Inlet Guide Vanes and stages 3, 4, 5 & 6 are variable.



72-00-45



Figure 31: HP Compressor



72-00-46

Training Manual A319/A320/A321 Combustion Section


Figure 32: Combustor Cut

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 nozzles. 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 casing. The inner combustion liner is attached to the turbine nozzle guide vane assembly. The inner and outer liners are manufactured from sheet metal with 100 separate liner segments attached to the inner surface. The segments can be replaced independently.



72-00-47



Figure 33: Combustion Section



72-00-48



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 exhaust 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 reverser and the pylon. The cold airflow exhaust is part of the thrust reverser system described in 78-3000. When the thrust reverser operates, the cold and hot outflows divide, and go in different directions.



72-00-49



Figure 34: Common Nozzle Assembly

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72-00-50



Angle and Main Gearbox The cast aluminium gearbox assembly transmits power from the engine to provide drives for the accessories mounted on the gearbox front and rear faces. During engine starting the gearbox also transmits power from the pneumatic starter motor to the engine. The gearbox also provides a hand cranking for the HP rotor (N2) for maintenance operations. The gearbox is mounted by 4 flexible links to the bottom of the fan case. main gearbox 3 links angle gearbox 1 link

Features Front Face • 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.)



72-00-51



Figure 35: Angle and Main Gearbox -!).'%!2"/8

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72-00-52



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!



72-00-53



Figure 36: Drive Seals



72-00-54



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).



72-00-55



Figure 37: Manual Handcranking 3 STARTER IDLER GEAR

4 WASHER (2 OFF)

5 NUT (2 OFF)

2 ADAPTOR

1 EXTERNAL GEARBOX MODULE

1 PACKING 2 CRANK COVER 3 WASHER

4 NUT 5 STARTER IDLER GEAR

6 EXTERNAL GEARBOX MODULE



72-00-56

Training Manual A319/A320/A321 Borescope Inspection of the HP Comp.


Figure 38: Borescope Inspection Equipment

Borescope ports are provided to give access for visual inspection of the compressor and the turbine. For further information and limits refer to AMM 72-00-00.

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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 the access port of the stage of the compressor 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

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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. !.',%+./" &/22)'(4!.$,%&4

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72-00-57

Training Manual A319/A320/A321 Borescope Access


Figure 39: Borescope Access Booster

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 damage 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 accessed 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.



72-00-58



Figure 40: HP Compressor Borescope Access



72-00-59





72-00-60



73 Engine Fuel and Control - V2500A



73-00-1


73-00 Fuel System Presentation

Power Plant V2500A 73-00 Fuel System Presentation

Controlling

General The fuel system enables delivery of a fuel flow corresponding to the power required 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,

The Fuel Authority Digital Electronic Control (FADEC) system provides full range control of the engine to achieve steady state and transient performance when operated in combination with aircraft subsystems. The FADEC is a dual channel EEC with crosstalk and failure detection capability. In case of specific failure detection, the FADEC switches from one channel to the other. The FADEC System operates compatibly with applicable aircraft systems to perform the following: • Control of fuel flow, stator vanes and bleeds to automatically maintain forward and reverse thrust settings and to provide 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 clearance 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.

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 controls the fuel flow supplied to the fuel nozzles (through the fuel flow meter and the fuel distribution valve). The FMU also provides hydraulic pressure to all hydraulic system external actuators. These include the Booster Stage Bleed Valve actuators, Stator Vane Actuator, 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 valve enables the selection of four basic configurations between which the flow paths of the fuel in the engine are varied to maintain the critical IDG oil, engine oil and fuel temperatures within specified limits. The transfer between configurations is determined by a software logic contained in the EEC.



73-00-2



Figure 1: Fuel System Schematic



73-00-3


73-10 Fuel Distribution Components


Description

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 fuel filter element is a low pressure filter which removes all contamination from fuel to go through it.

The FRV operates to control fuel flow which goes back to the aircraft fuel tank acting as a fuel cooler.

Fuel Filter

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 pressure switch is de - energized which causes the caution to go off. The differential pressure switch signal 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 management system (HMS) of the engine. The FD & RV has two valves in one body. They are a fuel diverter valve (FDV) and a fuel return valve (FRV).



73-00-4



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



73-00-5

Training Manual A319/A320/A321 Fuel Distribution Valve


Figure 3: FDV Location

General The fuel distribution valve (FDV) subdivides scheduled engine fuel flow from the fuel metering unit (FMU) equally to ten fuel manifolds, each of which in turn feeds two nozzles.

Description The fuel distribution valve is installed at the 4:00 o’clock location, at the front flange of the diffuser case. The fuel distribution valve receives fuel through a fuel line from the fuel metering unit. The fuel goes through a 200 micron strainer, and then into ten internal discharge ports. The ten discharge ports are connected to the ten fuel manifolds. Eight of the ten internal discharge ports in the valve are connected after an engine shutdown. Eight of the fuel manifolds are drained into the engine through the lowest fuel nozzle. The two fuel manifolds which remain full help supply fuel for the next engine start.



73-00-6



Figure 4: Fuel Distribution Valve



73-00-7

Training Manual A319/A320/A321 Fuel Manifold and Tubes


Figure 5: Fuel Nozzle

Description The fuel manifold and fuel tubes consist of several single wall tubes which carry fuel between components in the fuel system. Fuel supplied to the fuel nozzles is carried by a large tube from the fuel metering unit to the fuel distribution valve. At the fuel distribution valve the fuel supply is split and carried to twenty fuel nozzles by ten manifolds. Each fuel manifold feeds two fuel nozzles. Fuel pressure for actuating various valves is supplied by small tubes from the fuel metering unit mounted on the fuel pump. All the brackets and tubings are fire proof.

Fuel Nozzle General The fuel nozzles receive fuel from the fuel manifolds. The fuel nozzles mix the fuel with air, and send the mixture into the combustion chamber in a controlled pattern.

Description/Operation There are 20 fuel nozzles equally spaced around the diffuser case assembly. The fuel nozzles are installed through the wall of the case, and each nozzle is held in position by three bolts. The fuel nozzles carry the fuel through a single orifice. The fuel is vaporized by high-velocity air as it enters the combustion chamber. The fuel nozzle forms the atomized mixture of fuel and air into the correct pattern for satisfactory combustion.

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The design of the fuel nozzle results in fast vaporization of the fuel through the full range of operation. This results in decreased emissions, high combustion efficiency, and good start quality.

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The high-velocity flow of fuel prevents formation of coke on areas where fuel touches metal. Heatshields installed also prevent formation of coke.

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73-00-8



Figure 6: Fuel Distribution Tubes



73-00-9



Fuel Pump General The LP / HP fuel pumps are housed in a single pump unit which is driven by a common gearbox output shaft. A low pressure (LP) stage and a high pressure (HP) stage provide fuel at the flows and pressures required for operation of hydromechanical components and for combustion in the burner. The unit consists of a LP centrifugal boost stage which feeds an HP single stage, two gear pump. The housing has provision for mounting the fuel metering unit (FMU). The LP stage receives fuel from aircraft tanks through the aircraft pumps. The LP pump is designed to provide fuel to the HP gear stage with the aircraft pumps inoperative. After passing through the LP boost stage, fuel proceeds through the fuel filter to the HP gear stage. A coarse mesh strainer is provided at the inlet to the HP gear stage. This stage is protected from overpressure by a relief valve. Exceeding flow from the gearstage pump is recirculated through the FMU bypass loop to the low pressure side of the pump.

Fuel Metering Unit The FMU is the interface between the EEC and the fuel system. It is located on the dual fuel pumps unit, on the rear of the main gearbox, and is retained by four bolts as shown below. All the fuel delivered by the HP fuel pumps - which is much more than the engine requires - passes to the F.M.U. The FMU, under the control of the EEC meters the fuel supply to the spray nozzles. It also supplies HP fuel for the operation (muscle) of a number of actuators. Any fuel supplied by the HP pumps which is not needed for these two uses is returned, from the FMU to the LP side of the fuel system. In addition to the fuel metering function the FMU also houses the: • Overspeed Valve • Pressure Raising and Shut Off Valve 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.



73-00-10



Figure 7: Fuel Metering Unit



73-00-11

Training Manual A319/A320/A321 Fuel Metering Unit

Overspeed Valve

Fuel metering is achieved by the Fuel Metering Valve and the Pressure Drop Regulator and Spill Valve, which act together in the following sequence:

Operation

Signals from the EEC cause the torque motor to change position, which directs fuel servo pressure to re-position the Fuel Metering Valve. This changes the size of the metering orifice through which the fuel passes which in turn changes the pressure drop across the metering valve. The change in the pressure drop is sensed by the Pressure Drop Regulator which will re-position the spill valve and so increase/decrease the fuel flow through the fuel metering valve until the pressure drop is restored to its datum value. The increase/decrease in fuel flow causes the engine to accelerate/decelerate until the actual EPR is that demanded by the EEC signal. Movement of the Fuel Metering Valve is transmitted through a rack and pinion mechanism to drive a dual output position resolver. The resolver output is fed back to the EEC. The EEC automatically corrects changes in fuel density. Bi-metallic washers located in the pressure drop governor and spill valve assembly provide automatic compensation 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.


The overspeed valve is spring loaded to the closed position, it is opened by increasing fuel pressure during engine start and during normal engine operation is always fully open. In the event of an overspeed (109,1% N1, 105,4% N2) the EEC sends 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 around the overspeed valve to prevent engine flame out. The overspeed valve is hydraulically latched in the closed position, thus preventing the engine from being reaccelerated The recommended procedure is for the flight crew to shut down the engine. To shut down the engine is the only way to release the hydraulic latching. 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.

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.

It is commanded open by the MASTER SWITCH in the cockpit during MANUAL starts. The PRSOV can be commanded closed by the EEC during AUTO start sequences if the sequence has to be stopped for any reason. The EEC’s ability to close the shut off valve is inhibited above 43% N2. Above 43% N2, and in flight, the PRSOV can only be closed by the master switch in the cockpit. FAIL SAFE POSITION OF THE PRSOV: " LAST COMMANDED POSITION "

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 "



73-00-12



Figure 8: Fuel Metering Unit Schematic



73-00-13



Figure 9: FMU - Engine Shut Down



73-00-14

Training Manual A319/A320/A321 Figure 10: FMU - Engine Running



Figure 11: FMU - Engine Overspeed


73-00-15



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.

control this 90 deg. movement and set the electrical circuit for the next operation. One of the two motors can open or close the valve if the other motor does not operate.

Low Pressure Fuel Shut Off Valve

The actuator drive shaft has a see/feel indicator where it goes through the actuator body. The see/feel indicator gives an indication of the valve position without removal of the fuel LP fuel valve.

The LP fuel - valve 12QM (13QM) is in the fuel supply line to its related engine. The LP fuel - valve is usually open and in this configuration 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 sources: • 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 necessary to make sure that the valve can still operate. Thus the electrical supply to each motor goes through a different routing. The routing for motor 1 is along the front spar. The routing for motor 2 is along the rear spar and then forward through the flap track fairing at RIB 6. The actuators send position data to the System Data - 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 electrical motors which drive the same differential - gear to turn the ball valve through 90 deg. The limit switches in the actuator



73-00-16



Figure 12: LP Fuel Shut-Off Valve



73-00-17

Training Manual A319/A320/A321 HP & LP Fuel SOV Control

•

The HP fuel shut off valve control is fully electrical.


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.

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.

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 overrides 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).

Closure of the HP fuel PRSOV: it is controlled directly from the MASTER 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 independent 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 - energizes 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 actuator



73-00-18



Figure 13: HP and LP Fuel Shutoff Valve (SOV)

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73-00-19



73-20 Heat Management System

Fuel Temp. Thermocouple

Presentation

The Fuel Temperature is measured by the thermocouple at the fuel exit of the FCOC (Fuel Cooled Oil Cooler).

General

The thermocouple is composed of stainless steel sheathed sensing portion, stainless steel installing flange with seal spigot and electrical connector.

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.

The control of fuel temperature is done by the fuel diverter valve which is installed upstream of the FCOC.

FUEL TEMPERATURE:

IDG Oil Cooler Temp. Thermocouple

The fuel temperature is measured at the exit of the filter.

IDG Fuel Cooled Oil Cooler oil temperature is measured at the IDG Oil Cooler Exit by a thermocouple.

OIL TEMPERATURTE: The engine oil temperature is measured upstream of the ACOC.

The termocouple gives an electrical output in relation to the temperature of the oil in the fuel cooled IDG oil cooler.

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 minimising 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

ACOC Oil Temp. Thermocouple The oil temperature is measured at the ACOC inlet by a thermocouple. The thermocouple is composed of stainless steel sheathed sensing portion, stainless steel installing flange with seal spigot and electrical connector.

There are four basic configurations between which the flow paths of fuel in the engine L.P. fuel system are varied. Within each configuration the cooling capacity may be varied by control valves which form the Fuel Diverter and Back to Tank Valve. The transfer between modes of operation is determined by software logic contained in the EEC. The logic is generated around the limiting temperatures of the fuel and oil within the system together with the signal from the aircraft which permits/inhibits fuel spill to aircraft tanks.

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 return fuel from control actuator are plumbed to the diverter valve which normally turns the flow to the FCOC exit.


This temperature information is send to the EEC and is used for the heat management system.

The temperature is transmitted to the EEC (Electronic Engine Control). In response 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.


73-00-20



Figure 14: HMS Main System Components



73-00-21



Fuel Diverter & Return Valve

" FDV SOLENOID DE - ENERGIZED " (MODE 4 or 5)

General

Return to Tank Modes

The FDRV configuration allows four modes of operation according to electrical signals from the EEC (based on fuel and oil temperature measurements transmitted by thermocouples).

HMS Mode 1 (Normal Mode)

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 changes with the position of the valves.

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 tanks 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.

Fuel Return Valve

For example; • Low engine speeds. • High HP fuel pump inlet temperature.

The EEC operates the dual-wound torque motor to control the servo pressure.

In this mode the fuel/oil heat exchanger is operating as a fuel cooler.

This servo fuel pushes the main valve.

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.

Thus, the fuel flow can be controlled through the FDRV and the EEC.

The pressure balance between two sides of the main valve (Valves C and D) gives the direction and the speed of the valve movement.

The ACOC modulating valve is fully open.

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:



73-00-22

Training Manual A319/A320/A321 Figure 15: Return to Tank Mode 1



Figure 16: Return to Tank Mode 4


73-00-23



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 passing 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 permitted because IDG oil temperature is excessive or return to tank is not permissible due to the high return fuel temperature. The ACOC valve is fully open. This mode is adopted if the conditions exist. In case the oil temperature cannot be kept within the limits the FADEC system will increase the engine speed (FAIL SAFE MODE OF OPERATION. .



73-00-24

Training Manual A319/A320/A321 Figure 17: NO Return to Tank Mode 3



Figure 18: NO Return to Tank Mode 5


73-00-25



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.



73-00-26



Figure 19: Air Modulating Valve



73-00-27



IDG Fuel Cooled Oil Cooler The IDG oil cooler is installed at the left hand side on the fan case, near the FCOC. The IDG oil cooler has two sets of inlet and outlet ports. One set of ports is used for the flow of the fuel to or from the fuel diverter and return valve. The other set of ports is used for the flow of oil from and to the IDG. The hot scavenge oil which has been used to lubricate and cool the IDG, flows from the IDG to the oil cooler. As the oil goes through the oil cooler, the heat in the oil is transmitted to the fuel. The cooled oil then returns to the IDG. Two drain plugs are also installed in the oil cooler, one for the fuel and one for the oil. 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.



73-00-28



Figure 20: IDG FCOC Oil Cooler



73-00-29



Indicating

•

General

When the pressure loss in the filter decreases between 0 and -1.5 psid from the filter clog energizing pressure, the pressure switch is de-energized which causes the caution to go off.

Indicating

The associated caution message to come on the upper ECAM DU.

The differential pressure switch signal is fed directly to the SDAC through the hardware.

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 associated with the MASTER CAUT light and the aural warning (singlechime).

Fuel Flow Indication, Fuel Used The Fuel Flow Transmitter is installed near the FMU. The signals are routed to the EEC and via the DMCs to the ECAM. The Fuel Used-is calculated in the DMCs. The fuel flow transmitter signal is fed to the FADEC which processes it and transmits the information to the ECAM system for display.

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.



73-00-30



Figure 21: Fuel System Indication

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73-00-31



FADEC Presentation

In case of an overspeed, an incorporated valve reduces the fuel flow.

FADEC = Full Authority Digital Engine Control

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.

General

The functions of the FADEC are also reset when the ENG MASTER lever is set to OFF.

The Full Authority Digital Engine Control system consists of an Electronic Engine Control plus a Fuel Metering Unit, sensors and peripheral components.

Compressor Airflow and Turbine Clearance Control

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 command channel alternates each engine start.

Interfaces

The EEC controls the compressor airflow and the turbine clearance through separated 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.

The EEC receives air data parameters from the Air Data Inertional Reference System (ADIRS), and operational commands from the Engine Interface Unit (EIU).

Turbine clearance control: • HP and LP Turbine Active Clearance Control (ACC) valves. • 10th stage make-up air valve.

It also provides the data outputs nescessary for the Flight Management and Guidance Computers (FMGCs), and the fault message to the EIU for aircraft maintenance data system.

Engine oil cooling: • Air Cooled Oil Cooler (ACOC)servo valve.

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 correct fuel flow, overspeed protection and Engine Shut Down.



73-00-32



Figure 22: FADEC Presentation IAE V2500



73-00-33

Training Manual A319/A320/A321 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 directional valve included in the HCU to provide deployment and stowage of the thrust reverser translating sleeves.

•


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 parameters are optionally available. These are P12.5, P2.5 and T2.

FADEC System Maintenance Fault Detection The FADEC maintenance is facilitated by internal extensive Built in Test Equipment (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.

Start and Ignition Control

Non Volatile Memory

Each channel can control the starter valve operation, the fuel Pressure Raising Shut - Off Valve opening and the ignition during the engine start sequence.

In flight fault data is stored in FADEC non volatile memory and, when requested, is available on an aircraft centralized maintenance display unit.

Fuel Diverter and Return Valve

Communication with CFDS

The EEC manages the thermal exchange between the engine oil, IDG oil and engine fuel system by means of a Fuel Diverter and Return Valve.

Ground test of electrical and electronic parts is possible from cockpit, with engines not running, through the CFDS.

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 FADEC provides engine control system self testing to detect problems at LRU level.

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 engine oil temperature (T OIL).

FADEC is such that no engine ground run for trim purposes is necessary after component replacement.

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.



73-00-34



Figure 23: FADEC Presentation V2500



73-00-35

Training Manual A319/A320/A321 FADEC Functions The FADEC system operates compatibly with applicable aircraft systems to perform 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.

• • • • • •


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.

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)

7 Engine condition monitoring parameters transmission.

3 Power management • Automatic engine thrust rating control • Thrust parameter limit computation • manual power management through constant ratings versus throttle lever relationship – take-off / go-around at full forward throttle lever position – flex take-off at constant intermediate 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.

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.

8 Detection, isolation, accommodation and memorization of its internal system failures. 9 Fuel return & diverter valve control

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



73-00-36



Figure 24: FADEC Architecture



73-00-37



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



73-00-38



Figure 25: Engine Control P/B‘s and Switches

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ON O N O FF OFF

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73-00-39



Figure 26: Engine Circuit Breakers



73-00-40



Figure 27: Engine Circuit Breakers



73-00-41



Failures and Redundancy

Engine Surge

Improved reliability is achieved by utilising dual sensors dual feedback. • Dual sensors are used to supply all EEC inputs exept pressures, (single pressure 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 input signal and output controls. • Each channel normally uses its own input signals but each channel can also use input signals from the other channel if required i. e. if it recognises faulty or suspect, inputs. • An output fault in one channel will cause switchover to control from the other channel. • In the event of faults in both channels a pre-determined hierarchy decides 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 surge is detected by a rapid decrease in burner pressure or the value of rate of change of burner pressure, which indicates that surge varies with engine power level. Once detected, the EEC will reset the stator vanes by several degrees in the closed direction, open the booster 7th and 10th stage bleeds, and lower the maximum Wf/Pb schedule. Recovery of burner pressure to its steady state level or the elapse of a timer will release the resets on the schedules and allow the bleeds to close. Figure 28: Stall and Surge

Engine Limits Protection General The FADEC prevents inadvertent overboosting of the expected rating (EPR limit and EPR target) during power setting. It also prevents exceedance of rotor speeds (N1 and N2) and burner pressure limits. In addition, the FADEC unit monitors EGT and sends an appropriate indication to the cockpit display in case of exceedance of the limit. The FADEC unit also provides surge recovery.

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. Supplementary electronic circuitry for overspeed protection is also incorporated in the EEC. Trip signals for hardware and software are combined to activate a torque motor which drives a separate overspeed valve in the fuel metering unit to reduce fuel flow to a minimum value. The engine can be shut down to reset the overspeed system to allow a restart if desired.



73-00-42



Figure 29: FADEC Processing and Fault Logic



73-00-43



Power Management

Autothrust Activation / Deactivation

Autothrust Mode

The autothrust function (ATHR) can be engaged or active.

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 autothrust function will be active 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 manually or autothrust is reset with the autothrust 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 lever 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 levers is activated, • Or autothrust signalis lost from EIU.

Manual Mode

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 pushbutton on the throttle. The ATHR function is engaged automatically in the FMGC by auto pilot mode demand and manually by action on the ATHR pushbutton located on the Flight Control Unit (FCU). The ATHR de-activation and ATHR disengagement are achieved by action on the disconnect pushbutton located on the throttle 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 regardless 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 actual 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.

This mode is entered any time the conditions for autothrust or memo modes are not present. In this mode, thrust lever sets an EPR value proportional to the thrust lever position up to maximum take-off thrust.

AUTOTHRUST IS ONLY ACTIVE IN EPR MODE. IN RATED & UNRATED N1 MODE AUTOTHRUST IS LOST.

Flexible Take-Off Rating FLEXIBLE TAKE-OFF rating is set by the assumed temperature method with the possibility to insert an assumed temperature value higher than the maximum one certified for engine operation. (30 deg C.)



73-00-44



Figure 30: Auto Thrust Definition

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73-00-45



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 current thrust setting mode. If the fault is no longer present, the EEC will switch to the primary control mode (EPR). If the fault later reoccurs, reversion back to N1 mode (rated or unrated) will result.



73-00-46



Figure 31: Power Setting Requirements Schematic



73-00-47



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 utilizes 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.



73-00-48



Figure 32: Rated N1 Mode



73-00-49



Unrated N1 Setting Requirements Unrated N1 The loss of the T2 signal will cause automatic 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 rated thrust is not recommended and will result in reduced engine life. The maximum N1 must therefore be determined from charts in the Flight Crew Operating 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

Complete Output Signal Failure

General

In case of complete output failure there will be no current flow through torque motors or solenoids. The associated component will be the " FAIL-SAFE " position.

The Electronic Engine control (EEC) system is dual, the two channels are equal. Failures are classified as class 1, 2, 3.

If the EEC power supply is lost, the components will go into"FAILE-SAFE" position.

According to the failure class, the system can use data from the other channel, or switch to the other channel. Faults are memorized in the system BITE as they occur.

Input Fault Strategy All sensors and feedback signals are dual. Each parameter sensor as well as feedback sensors used by each channel come from two different 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 automatic switchover to the standby active channel.

T/S Action One Channel - most likely LRU failure.



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Figure 34: FADEC Single Input Signal Failure



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Component Fail Safe States

COMPONENTS:

FAIL SAFE STATE:

METERING VALVE

MIN FLOW

VARIABLE ST ATOR 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 PARTIAL LY 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

* If there is a failure of the thrust reverser hydraulic control unit directional valve while the reverser is deployed, the reverser will remain deployed.



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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 IN VALI D,POWER IS SET TO IDLE. ON GROUND: SET IDLE IN FLIGHT: AT TAKE OFF FREEZE LAST VALID TLA,THEN SELECT MCT AT SLAT RETRACTION AUTOTHRUST CAPABILITY.

ONE TLA:

THE EEC USES THE REDUNDANT SENSOR.

BOTH 115V AC:

NO IGNITION NO P2/T2 PROBE HEATING

BOTH 28V DC:

NO START RUN ON ALTERNATOR AB OVE 10% N2

DISAGREEMENT BETWEEN TRA :

ON GROUND: IN FLIGHT: ON REVERSE:



SET FORWARD IDLE SELECT LARGER VALUE BUT LIMIT THIS TO MCT 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 sufficiently 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 reverse idle detent position. Bleed Idle = Bleed demand. Bleed Idle command will set the fuel flow requested for ensuring correct aircraft ECS system pressurization, wing 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 THRUST LEVERS

EIU

Reverse Idle

EIU

Approach Idle

TLA (REV. IDLE) LANDING GEARS

SLAT / FLAP LEVER

WOW (GRD)

LGCIU 1/2

AIR 0

0

1

1

2

2

3

3

FULL

SFCC 1/2

LEVER NOT ZERO

EIU FAULT

FULL

WING ANTI ICE

N2 Idle

Min Idle

Setting

ENG ANTI ICE

ECS DEMAND

ZONE CONT.

EIU

Bleed Idle

ENGINE FUEL TEMPERATURE

PACKs

HMS

PACK CONT. 1/2

EEC



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N1 Speed Table


20-

.


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Figure 36: Ground Idle Speed Diagram N2

V2530-A5 SLS / STD GROUND IDLE ( NO OFFTAKES )

N2 ROTOR SPEED ( RPM / % )

9600 64,2%

9200 61,5%

8800 58.8% 57,5% 8400 56,1%

8000 53,5%

7600 50,8% –80

–60

–40

–20

0

+10

+15

+20

+30

+40

+50

AMBIENT TEMPERATURE ( DEG. C. )



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FADEC Power Supply

FADEC Ground Power Panel

EIU Power Supply

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.

The EIU is powered from the aircraft electrical 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%.

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.

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 directly 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.



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Figure 37: FADEC Power Supply



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FADEC LRU‘S

Data Entry Plug Modification

Electronic Engine Control (EEC)

Description

Data Entry Plug

The DEP links the coded data inputs through the EEC by the use of shorting jumper leads which are used to select the plug pins in a unique combination.

The Data Entry Plug (DEP) provides discrete inputs to the EEC. Located to the Junction 6 of the EEC it provides unique engine data to channel A and B. 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 revision from EPR mode to unrated N1 mode occurs.

During a life of an engine, it may be necessary to change the DEP configuration, either during incorporation of Service Bulletins or after engine overhaul, when the EPR modifier code may need to be changed. This is 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 connectors 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 connector 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 assemblymust 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)

Electrical Connections

Harness (Electrical) and Pressure Connections

Front Face

Two identical, but separate electrical harnesses provide the input/output circuits between the E.E.C. and the relevant sensor/control actuator, and the aircraft interface.

J1

E.B.U. 4000 KSA

The harness connectors are ’keyed’ to prevent misconnection.

J2

Engine D202P

J3

Engine D203P

J4

Engine D204P

J11

Engine D211P

Single pressure signals are directed to pressure transducers • located within the E.E.C. • the pressure transducers then supply digital electronic signals to channels A and B.

Harness Connector Plug Identification

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)


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)



73-00-65



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



73-00-67

Training Manual A319/A320/A321 FADEC LRU‘S Sensors


Figure 41:

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 temperature, 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 steady state operation of the engine.



73-00-68



Figure 42: P3/T3 Sensor .54

#(2/-%,

.54

!,5-%, .54 .54

0302%3352% 45"%

043%.3/2 #2

#2

!,

!,

02%3352% 0/24

"/,4

3#(%-!4)#

'!3+%4 $)&&53%2 #!3% !33%-",9

"/33



73-00-69



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 alternator(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.



73-00-70



Figure 43: P2.5 / T2.5 Sensors



73-00-71



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 - SYSTEM 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 EEC compares feedback position against commanded position.

The trouble shooting menu has 4 submenus: • FLIGHT DATA • GROUND DATA • AIRCRAFT DATA • EEC CONFIGURATION

If failed in one channel: • EEC switches to the other channel (the ability to switch is based on relative helth of the other channel) If failed in both channels: • Healthiest channel continues to command actuator. T/S ACTION:

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 identification which CFDS FAULT message belongs to which troubleshooting data (eg. Ground Scanning menu.)

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 from the LRU LVDT‘s or microswitches.

In the example a OSPXCF (OVERSPEED CROSS CHECK FAILURE) is indicated.

The EEC compares channel A against Channel B.

Ground Data

Failure of Reverser: EEC will select most stowed and will not allow a deploy.

This menu gives additional failure data (temperatures, pressures, RPM, etc.) when a fault occured on ground. This data is saved in a CELL.

Failure of Temperature sensors: EEC will use fail safe value.

The cell allows a identification which CFDS FAULT message belongs to which troubleshooting data (eg. Ground Scanning menu.)

Most likely a LRU problem, next check harness then EEC

Failure of TRA: EEC has specific fault accomodation based on previous value.

T/S ACTION:

Input Latched Failed (ILF)

FADEC Failure Types Definition

(Single Input Signal Failure) There is no channel changeover for input signal failure, as long as the Cross Channel Data Link is operativ.

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 helth of the other channel)

Faults are not latched. Thus automatic recovery is possible.

If failed in both channels: • specific output is depowered (exception - solenoids are depowered in groups) T/S ACTION: Most likley a loose connector or chaffed harness next LRU and finally EEC.

Track-Check Failures (TKF) Failure of the system to follow the commands of the EEC.



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Figure 45: Trouble Shooting Report



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Figure 46: Flight Data / Ground Data

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 )

FADEC 1B FLIGHT DATA

Page one of the Cell 31

PG:01 Fault Code

CELL: 31 FAULT: WOFWAF RPM: N1 = 5326 N2 = 14392 DEG C: T5 = 554.0 T2 = 26.0 FLTPH = 3 TCJC = 42.0 PSIA: PB = 458.5 P2 = 14.62

Mach Number

N2 RPM T2 Temperature ( Eng. Inlet ) Flight Phase Total Air Pressure ( Eng. Station 2 ) EEC Operating Hours

MN = .117 HOURS = 571.0

Note: The Abbreviations used in the GROUND DATA are the same.



73-00-76



Figure 47: Flight Data / Ground Data

FADEC 1B FLIGHT DATA

Page two of the Cell 31

PG:02

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

CELL: 31 FAULT: WOFWAF ALT: = 336.0 FT EPRI = 1.562 SVA : = 1.906 INCH INCOM = 1 FF = 11162 PPH BACKUP = 0 B 25 = 1.218 INCH LEG = 398.0

EPR ( indicated ) Channel in Control 1 = Yes , 0 = No N1 Mode 1 = Yes 0 = No ( EPR Mode ) Flight Legs

WOW = 1

Note: The Abbreviations used in the GROUND DATA are the same.



73-00-77

Training Manual A319/A320/A321 FADEC System Test The FADEC SELF TEST should really be known as the FADEC SYSTEM TEST.


12. Pressure Sensor(s) Disagree - Indicates that the static pressure sensor test ran and any two pressure sensors were not within the specified tolerances.

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 tested 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 that the activity monitor circuit test failed or the local channel was unable to provide power to any Output or there were interface or input fault. 8. Input / lnternal Test Passed - Indicates 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 - Indicates 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.



73-00-78



Figure 48: FADEC Self Test



73-00-79

Training Manual A319/A320/A321 FADEC Ground Scanning


Figure 49: 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.)



73-00-80



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



73-00-81

Training Manual A319/A320/A321 Scheduled Maintenance Report


Figure 51: SMR Menu

At the aircraft level : • Level “A” faults are considered class ” 1” faults with the associated specific information to the flight crew (ECAM warnings, advisory information...). • Level “B” faults are regrouped under the generic ECAM warning “ENG X MINOR 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 Industrie and the engine manufacturers in order to fit the dispatch time limitation associated 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 manufacturers 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 introduce 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.



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Figure 52: Scheduled Maintenance Report

V2500-A1 EEC Standards

EIS V2500-A1

“SMR“ Menu

“CLASS 3 Faults” Menu Unlimited Class 3 Faults

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

SCN 12C

and subsequent

V2500-A5 EEC Standards

V2500-A5

“CLASS 3 FAULTS” Menu Unlimited Class 3 Faults

“SMR” Menu


+

Long Time Dispatch Faults SCN 10A

“CLASS 3 Faults” Menu Unlimited Class 3 Faults

and subsequent


73-00-83


Engine Interface Unit


Figure 53: EIU Location

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 between the airframe and the corresponding FADEC located on the engine, thus reducing 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 channel 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.



73-00-84



Figure 54: EIU Schematic



73-00-85



EIU Interfaces SIGNALS WING ANTI-ICE SWITCH ENGINE FIRE P/B SIGNAL LOW OIL PRESSURE SWITCH (AND GROUND)

FADEC GROUND POWER P/B LGCIU 1 AND 2 (GROUND SIGNAL) SFCC 1 AND 2 SEC 1 ,2 AND 3 FLSCU 1 AND 2 ENGINE SELECTED OIL PRESSURE,OIL QUANTITY AND OIL TEMPERA TURE NACELLE TEMPERA TURE START VALVE POSITION (FROM EEC) N2 GREATER THAN MINIMUM IDLE (FROM EEC) ENGINE START FAULT SIGNAL APU BOOST DEMAND SIGNAL (EIU) TLA IN TAKE-OFF POSITION (MIN. T/O N2, FROM EEC)

THRUST REVERSER (FROM SEC 1,2 AND 3 )


PURPOSE ENGINE BLEED COMPUTATION LOCIG FADEC ENGINE SHUTDOWN LOGIC -COCKPIT WARNING SIGNALS -HYDRAULIC MONITORING -WINDOW AND PROBE HEATING SYSTEM -AVIONIC VENTILATION SYSTEM -RAIN REPELLENT SYSTEM -CIDS,CVR,DFDR FADEC POWER SUPPLY LOGIC THRUST REVERSER AND IDLE LOGIG ENGINE FLIGHT IDLE COMPUTATION LOGIC THRUST REVERSER INHIBITION CONTROL HEAT MANAGEMENT SYSTEM FUEL RETURN V ALVE CONTROL ENGINE 1 OR 2 INDENTIFICATION INDICATION ECAM INDICATING (ECAM) ECS FOR AUTOMATIC PACK VALVE CLOSURE, DURING ENGINE START FUNCTIONAL TEST INHIBITION OF THE RADIO AL TIMETER TRANSCEIVER -BLUE HYDRAULIC SYSTEM PUMP CONTROL ILLUMINA TION OF FAULT LIGHT ON THE ENGINE START PANEL MAIN ENGINE START MODE TO THE APU ELECTRONIC CONTROL BOX PACK CONTROLLER FOR INLET FLAP CLOSURE -AVIONIC EQUIPMENT VENTILATION CONTROLLER ( CLOSED CIRCUIT CONFIGURATION ) -CABIN PRESSURIZATION COMPUTER PRE-PRESSURIZATION MODE THRUST REVERSER INHIBITION RELAY


73-00-86



CFDS System Report/Test EIU

The EIU is a Type 1 System.

This Page shows the menu of the Engine Interface Unit (EIU).

The EIU is availlable in CFDS back up Mode.

Figure 55: EIU Menu



73-00-87

Training Manual A319/A320/A321 LAST Leg Report

LRU Indentification

Last leg Report

Shows the EIU part number.


Here are Displayed the Internal EIU Faillures that Occured during Last Flights. Figure 56: Last Leg Rep./ LRU Indentification



73-00-88

Training Manual A319/A320/A321 Ground Scanning

•


RTOK means Re - Test Ok, you can ignore this Fault

This Page gives the EIU Faillures still presend on Ground. Figure 57: Ground Scanning



73-00-89



EIU CFDS Discrete Outputs Simulation

Simulation: "LOP GND 1 "

The Purpose of this Menu is to Simulate some Engine Interface Unit (EIU)

Use this key to simulate OIL LOW PRESS & GND for these systems through the MCDU:

Discrete Outputs by Setting their Status to 0 or1. The DISCRETE OUTPUT SIMULATION can operate systems and components without special indication on the 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 "

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 architecture as long as you work on the EIU DISCRETE OUTPUTS menu.

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.

Simulation: "LOP GND 2 "

APU BOOST 1 simulates a not closed starter air valve. • The APU is boosted (if running).

Use this key to simulate OIL LOW PRESS & GND for these systems through the MCDU: PHC1, PHC3, WHC1, AEVC, DFDR and CVR.

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 STRICTLY REQUIRED FOR TROUBLE SHOOTING PURPOSE THUS TO AVOID PREMATURE CORE PERFORMANCE DETERIORATION. IF REALLY REQUIRED, DO NOT OPERATE THE APU FOR MORE THAN ONE (1) MINUTE IN BOOST MODE CONDITION. NOTE THAT FOR TROUBLE SHOOTING AN ECS SYSTEM MALFUNCTION, THE ENGINE BLEED SHOULD BE PREFERRED.

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 extraction is low in the avionics compartment.

Simulation: " T/R INHIB "

Simulation: " FAULT "

To simulate the authorization of closure of the thrust reverser directional control valve solenoid (through the relay 14KS1(2)) through the MCDU.

Use this key to simulate a disagree between the position and the command of the HP fuel valve 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

FAULT discrete output status becomes "1" and FAULT legend of the 5KS1(2) annunciator light comes on.



73-00-90



Figure 58: Discrete Outputs Simulation



73-00-91

Training Manual A319/A320/A321 EIU CFDS Discrete Outputs Simulation


Simulation of " TLA > MCT " To simulate "TLA > MCT" for the following systems:

Simulation: " HP FUEL PN "

AEVC, PACK CONTROLLERS, CABIN PRESSURE CONTROLLERS.

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.

Push the line key adjacent to TLA "TLA > MCT" discrete output status > MCT becomes "1" On the ECAM PRESS page check that the inlet and extract skin air valves close.

The zone controller uses the HP fuel valve position to make the bleed status on label 061. Then it sends it to the EEC through the EIU (label 030). This input can change the bleed status only if the PRV opens (engine in operation).

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 OUTPUTSmenu.



73-00-92



Figure 59: Discrete Outputs Simulation



73-00-93



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)



73-00-94



Figure 60: EIU Discrete Outputs



73-00-95





73-00-96



74 Ignition - V2500A



74-00-1


Power Plant V2500A 74-00 Ignition System Presentation

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 supplied from the AC 115V normal and standby bus bars to the relay box. The 115V relays which are used to connect / isolate the supplies are located in the relay box and are controlled by signals from the EEC. The same relay box also houses the relays which control the 115V AC supplies for P2/T2 probe heating. According to M.E.L. the IGN. system A is required as minimum!



74-00-2



Figure 1: Ignition System Component



74-00-3



Ignition Starting - Operation

Continuous Ignition Selection

Description

Manual Selection

The ignition circuit is supplied with 115VAC - 400Hz. The electrical power is supplied via the EEC and EIU which controls the ignition of the igniter plugs.

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.

A dormant failure of an ignition exciter is not possible for more than one flight because: • the two ignition systems are independent • the EEC selects alternately ignition system A or B.

Automatic selection

FAIL SAFE POSITION: "IGN RELAYS, IGN ON" Ignition during Automatic Start Sequence When an automatic start sequence has been activated by the EEC (ENG/ MODE selector switch in IGN/START position and MASTER control switch to ON), the EEC energizes automatically the appropriate ignition exciter when N2 reaches between 10%-16% depending on TAT and keeps it energized until N2 reaches 43%. For inflight restart the EEC selects simultaneously both ignition exciters On the ground, after engine start, the selector must be placed in NORM position, then back to IGN/START to select continuous ignition. (both ignitors) In flight after engine restart, if the selector is maintained in IGN/START position, the EEC selects the continuous ignition on the corresponding engine In case of a fault during an automatic starting on the ground, the EEC aborts automatically the sequence by closing the starter shut-off valve and the HP fuel shutoff valve and deenergizing the ignitors.

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 running, through the maintenance MENU mode of the FADEC or manually (Manual Start without air)

Ignition System Circuit Breakers There are 5 ignition CB’s installed in the cockpit. 49VU and 121VU

Ignition during Alternate Start Sequence (Manual Start Procedure) When a manual start sequence has been activated by the EEC (ENG/MODE selector switch in IGN/START position and the ENG/MAN START pushbutton switch selected to ON) the EEC energizes both ignition exciters. The deenergization of the ignition exciters is automatically commanded by the EEC when engine N2 speed reaches 43%. (Starter cut-out) Positioning of the MASTER control switch to OFF, during that starting sequence, results in ignition exciter deenergization.



74-00-4



Figure 2: Ignition and Starting System Eng. 1

CHA

IGNITER PLUG A

CHB 401XP 115VAC ESS BUS

1JH ENG 1&2/IGN/SYS A 49VU

EXCITER A

5A

901XP 115VAC STAT INV BUS BAR

2JH1 3A 103XP 115VAC BUS 1

401PP 28VDC ESS BUS

FILTER

CHB

3JH1 FILTER

ENG/1GN 1/SYS B 121VU 3A

FIRE P/BSW 1WD

4100 KS RELAY BOX

2KS1

A

FILTER

A

3A

STARTER VALVE

4KS1 301PP 28VDC BAT BUS

IGNITER PLUG B

EXCITER B

CHA

B

FILTER CRANK NORM IGN/ START

3A

3KC

ON OFF

B

ON OFF MAN START 11 12 13 14 15 16 LABEL 031

B A

FILTER

PERMANENT

ECAM

CRANK NORM IGN/START

MAGNETIC A1

6KS

A2

ALTERNATOR A

B1 ENGINE 1 ON


9KS1 MAN START 1

B2 1KS1 EIU2 86VU


4000KS1 EEC1

74-00-5



Ignition System Test Igniter Plug Test The operation of the igniter plugs can be checked on the ground, engine not running, through the maintenance MENU mode of the FADEC. The test will be performed by selecting the 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.



74-00-6



Figure 3: FADEC Ignition Test



74-00-7



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.



74-00-8



Figure 4: FADEC Ignition Test Cont.



74-00-9

Training Manual A319/A320/A321 Ignition Test without CFDS

MAKE SURE THAT THERE IS ZERO PSI AT THE STARTER VALVE INLET BEFORE YOU PUSH THE MAN START P/B.

For the test procedure, refer to AMM TASK74-00-00-710-041-01

READ THE PRESSURE ON THE ECAM START PAGE.

During the test, an aural check of the ignitor plug operation has to be done.




74-00-10



Figure 5: Ignition Test without CFDS



74-00-11


74-00 Starting 80-00

•

Power Plant V2500A 74-00 Starting 80-00

the MASTER control lever controls the HP fuel shut-off valve. No start abort by the FADEC in case of failure.

General Starting Schematic The starting system of the engine utilizes pressurized 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 starter shut-off valve (SOV) pneumatically operated 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 disengages 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 fuel 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 position 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 pushbutton switch command the starter shut-off valve,



74-00-12



Figure 6: Starting System Schematic /4(%2 !# 3934%-3

"-#

/6%2"/!2$

02%#//,%2

4,4 &)2%

4#4

7!,, 0

0

&!6

/06

026

(06

4/(9$2!5,)#2%3%26/)202%3352):!4)/. %.').%/.,9

)0#

34!24%2

(0 )0

6!,6%

%.').%

34!24%2


4/.!#%,,% !.4) )#%6!,6%


74-00-13



Starting Components

Starter Air Control Valve

Starter Motor

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 pneumatic starter motor is mounted on the forward face of the external gearbox and provides the drive to rotate the H.P. compressor to a speed at which light up can occur. 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 lubricated by an integral lubrication system.

The start valve controls the air flow from the starter air duct to the starter motor. The start valve basically comprises a butterfly type valve housed in a cylindrical valve body with in-line flanged end connectors, an actuator, a solenoid valve and a pressure controller. A micro switch provides valve position feed back information to the FADEC.

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 turbine unit that accelerates the H.P. rotor to the required speed for engine starting. The unit is mounted on the front face of the external gearbox. The starter, shown below, comprises a single stage turbine, a reduction gear train, a clutch and an output drive shaft - all housed within a case incorporating an air inlet and exhaust. Compressed air enters the starter, impinges on the turbine blades to rotate the turbine, and leaves through the air exhaust. The reduction gear train converts the high speed, low torque rotation of the turbine to low speed, high torque rotation of the gear train hub. The ratchet teeth of the gear hub engage the pawls of the output drive shaft to transmit drive to the external gearbox, which in turn accelerates the engine H.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 output drive shaft continue to rotate. When the starter output drive shaft rotational speed increases above a predetermined r.p.m., centrifugal force overcomes the tension of the clutch leaf springs, allowing the pawls to be pulled clear of the gear hub ratchet teeth to disengage the output drive shaft from the turbine.



74-00-14



Figure 7: Starting Components



74-00-15



Starter Air Control Valve Description The start air control valve is a pneumatically operated, electrically controlled shutoff valve positioned on the lower right hand side of the L.P. compressor (fan) case.

Manual Operation The starter air valve can be opened/ closed 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"



74-00-16



Figure 8: Starter Air Control Valve



74-00-17



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.



74-00-18



Figure 9: Starter Valve Test via CFDS



74-00-19



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



74-00-20



Figure 10: Starter Valve Test via CFDS



74-00-21



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 closed 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.



74-00-22



Figure 11: Dry Cranking Procedure



74-00-23



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 breakers) 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 the OFF position to shut-off the fuel, also the starter valve closes. The EEC automatically reengages the starter at 10% N2 and the engine should be motored for at least 60 seconds to eliminate 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 continuous cranking, stop for a cooling period of 30 minutes.



74-00-24



Figure 12: Wet Cranking Procedure



74-00-25

Training Manual A319/A320/A321 Automatic Start The automatic start mode gives the EEC full control to automatically sequence the starter air valve, ignition relays and the fuel 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 numbers where windmilling conditions are insufficient for engine starting. (The EEC has input data necessary to activate starter assist function where necessary.)

• •


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, ignition, and starter air and provides the appropriate fault indication to the cockpit. (Auto Start Fault) Autostart fault messages will be displayed until approximately idle speed. The EEC’s ability to shut off fuel is inhibited 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 command indicated on the ECAM page ("Master lever OFF"). The operator then has to decide to perform a new engine start or troubleshoot the system.

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 manual 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



74-00-26



Figure 13: Automatic Start Procedure



74-00-27

Training Manual A319/A320/A321 Manual Start


Figure 14: ECAM Start Pages

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 sequence the starter, ignition and 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 set 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 ignitors 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 the ignitors off and the EEC is resetted.



74-00-28



Figure 15: Manual Start Procedure



74-00-29



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



74-00-30



Figure 17: Continuous Ignition Logic 401XP.B 115VAC ESS BUS 24-58-11

2

3

4 31-54-06

D

AB 14B 14D UNSD AB

D

WIRING DIAGRAM

2 7 6 7

1JH C/B ENGINE/1 AND 2/ ING/SYS A 49VU210

901XP.A 115VAC STAT INV BUS 24-58-14

103XP.A 115VAC BUS 1 24-58-02

SUPPLY SWITCHING

AB 14F

1

6 7

1

2

3

4 31-54-06

2JH1 (2JH2) C/B ENGINE/ING/ENG1 (ENG2)/ SYS A/BAT 121VU212

204XP.A 115VAC BUS 2 24-58-02

1 3

14H AB AC 5

4 5

4 6

2 4 31-54-06

AC

1

AC

C

UNSD AC

C

A

J3

A

J3

E

J9

E D

J9

74-31-01 74-31-02

N M

CH.A

N M

CH.B A

D

1 UNSD 3 A

2 7 9 8

3JH1 (3JH2) C/B ENGINE/ING/ ENG1 (ENG2)/SYS B 121VU212

9 8 4 5

A

J3

A

J3

E

J9

E C

J9

R
CH.A

R
CH.B

C

4100KS RELAY BOX 436 (446) 73-25

28VDC

AA 5B

SCH02

A

LIGHT "FAULT" ON

AB A 21(3) 9A A

A 10(1) A

3A 2A

5A 4E

OFF ON

SCH13 (SCH21) 8LP (19LP) BOARD ANN LT TEST & INTFC 70VU126 33-14

SCH01 3KC (2KC) CTL SWENG/MASTER 1(2) 115VU210 76-12

A

7 4

BUS A

A B

5A 5C AB

A B

AA 14K 15J

CRANK AUTO IGNITION EEC1 (EEC2) INPUT B1

A

A C3 A

1F AA

J1 J7

A B

EIU-1 (EIU-2) BUS A INPUT

CH.B

P N

A B

EIU-1 (EIU-2) BUS A INPUT

CH.A

A B

73-25 SCH10 EEC B OUTPUT 1

J7 J1

L K

A B

EEC B OUTPUT 2

L K

A B

73-25 SCH10 EEC A OUTPUT 2

A B

EEC A OUTPUT 1

MAN ENG START EEC1 (EEC2) INPUT A1

A B

15A 15C AA

4000JH2 EXCITERIGNITION, B 454 (464)

P N

SCH08

A UNSD C2 C1 A

26-12 SCH01

5 6 UNSD A

SCH09

J1

TO EEC ARINC INPUTS & OUTPUTS 73-25 SCH10

4001JH2 PLUG-IGNITER, B 454 (464)

UNSD 3 A

5KS1 (5KS2) ANNENG/1 (2) FIRE/FAULT 115VU210 73-25

J7

1C 1A 1B

9 8

FIRE R FAULT A

J7

1(21) 2(22) 3(23) 4(24)

2 1

A 45(2)

1WD ENG/APU FIRE PNL 210 26-21

6KS SEL SWENG/MODE/CRANK AUTO IGN/IGN 115VU210 73-25

4000JH1 EXCITERIGNITION, A 454 (464)

SCH18

A(C) P(G) UNSD S(J) R(H) A(C)

1A

4001JH1 PLUG-IGNITER, A 454 (464)

2

J1

CH.B

CH.A

J8
A

A

7 8

DISCRETE

3 4

SOLENOID

5 6

DISCRETE

1 2

SOLENOID

CHANNEL B

SCH08 9KS1 (9KS2) P/BSWENG/MAN START/1(2) 22VU212 73-25


CHANNEL A

SCH04

SCH03 1KS1 EIU-1 85VU127 73-25

1KS2 EIU-2 86VU128

4000KS EEC 436 (446) 73-25


4005KS VALVEPNEUMATIC STARTER 436 (446)

74-00-31



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

Oil Pressure Oil temperature


-40 deg.c

-10 deg.c

50 deg.c

156 deg.c amber


Minimum

Maximum

60 psi

ISA dependant

165 deg.c red

74-00-32



75 Engine Air - V2500A



75-00-1


75-00 System Presentation

Power Plant V2500A 75-00 System Presentation

Compressor Control General

General • • • • •

The booster stage bleed valve, the variable stator vane and HP compressor bleed valves systems are controlled by the EEC. The booster stage bleed valve controls the LP compressor airflow. The variable stator vane and the 7th and 10th stage bleed valves control the HP compressor airflow.

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)

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.

75-30 Compressor Control • LP Compressor Airflow Control System • HP Compressor Airflow Control System

At low LP spool speeds the booster provides more air than the core engine can utilize. To match the booster discharge airflow to the core engine requirements at low speed, excess air is bled off through booster stage bleed valves (BSBV) into the fan discharge air stream. At higher engine speeds the BSBV are closed so that all the booster discharge (primary air flow) enters the core engine.

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.

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 internal air system consits of: • Propulsion airflow (secondary & primary flows) • Bearing compartments pressurizing air • Cooling air

The VSV system maintains a satisfactory compressor performance over a wide range of operating conditions. The system varies the angle of the inlet guide vanes and stator vanes to aerodynamically match the low pressure stages of compression with the high pressure stages. This variation of vane position changes the effective angle at which the air flows across the compressor blades and vanes. The VSV angle determines the compression characteristics (direction and velocity) for any particular stage at compression.

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.

HP Compressor Bleed Valves The 7th and 10th stages bleed valves maintain a more stable operation of the compressor.

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.



75-00-2



Figure 1: Compressor Control Schematic



75-00-3



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 operates to control the air bled from the LP compressor. The two actuators are mechanically attached to each actuating rod and, the bleed - valve and actuating mechanism. The two actuators are connected hydraulically and operate together by command and feedback signals from/ to the EEC. FAIL SAFE POSITION: "BSBV OPEN" In case of a malfunction "ENG 1 (2) COMPRESSOR VANE" is displayed on the ECAM E / WD.



75-00-4



Figure 2: Booster Stage Bleed Valve System



75-00-5



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.



75-00-6



Figure 3: BSBV and Actuating Mechanism



75-00-7


75-32 HP Comp. Air Flow Sys.


Operation of the VSV Actuator

VSV System Components The four stages of variable incidence stators comprise inlet guide vanes to stage 3 and stages 3, 4 and 5 stator vanes.

Dual wound torque motors convert electrically isolated drive signals from each channel of the Electronics Engine Control (EEC) into hydraulic drive signals to position the actuator piston. If power to the stator vane actuator torque motor is lost, the stator vane actuator will go to the full open position.

Variable Stator Vane Actuation Mechanism

General The purpose of this system is to position the Inlet Guide Vanes (IGV) and stator vanes, using a fuel driven hydraulic actuator, in response to electrical signals provided by the EEC.

Variable Stator Vane (VSV) Control The VSV position is controlled by the EEC as a function of N2 / square root of theta T 2.6 (synteziesed value). The EEC uses the VSV feedback signal from the LVDT‘s to adjust the actual VSV position.

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:

Description

"VANES OPEN"

Variable Stator Vane Actuator

In case of a malfunction "ENG 1 (2) COMPRESSOR VANE" is displayed on the ECAM E / WD.

The stator vane actuator accurately controls vane movement with respect to a torque motor current supplied by the EEC. Operation of the stator vanes in regulated by accurate control of high pressure fuel flow to one or other side of a differential area piston. The piston has an externally adjustable low speed stop at the extended end of its travel. The high speed stop is formed by a collar which limits piston retraction. Provision is made to lock the piston with a rigging pin for setting purposes.

Linear Variable Differential Transformer (LVDT) A Dual Wound Linear Variable Differential Transformer (LVDT) is located in the center of the actuator piston rod. The LVDT completes the electronic control loop by providing a signal of actuator position to the Engine Electronic Control.

Engine Linkage with the VSV Actuator The engine IGV and Stator Vane linkage is connected to a fork end on the piston rod of the VSVA unit. The securing pin of link on to fork end.



75-00-8



Figure 4: VSV System Components



75-00-9



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 crankshaft and the actuator, as shown below. After the fuel supply and return tubes have been disconnected the crankshaft should be rotated to align the rig pin holes in the input lever and the front bearing housing. Spanner (Wrench) flats are provided on the crankshaft for this purpose. Installing the rig pin locks the crankshaft assembly with the actuator and vanes in the high speed position (actuator fully retracted).



75-00-10



Figure 5: VSV Actuator Rig



75-00-11



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 conditions. A total of four bleed valves are used, three on stage 7 and one on stage 10. The handling bleed valves are ‘two position’ only - fully open or fully closed, and are operated pneumatically by their respective solenoid control valve. The solenoid control valves are scheduled by the EEC 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 always be in the correct position (open) for starting.



75-00-12



Figure 6: HP Compressor Bleed Valves



75-00-13



Handling Bleed Valves Function

During an engine deceleration the reverse operation occurs and the bleed valve opens.

Description

Handling bleed valves (surge bleed)

The bleed valve is a two position valve and is either fully open or fully closed. The bleed valve is spring loaded to the open position and so all the bleed valves will be in the correct position - open - for the engine start. When the engine is started the bleed air from the engine will try to close the valve. The valve is kept in the open position by servo air (P3) supplied from the solenoid control valve (solenoid de-energised). The bleed valves will be closed at the correct time during an engine acceleration by the EEC energising the solenoid control valve vents the P3 servo air from the opening chamber of the bleed valve, and the bleed valve will move to the closed position.

The bleed valves and the solenoid control valves all operate in the same manner. FAIL SAFE POSITION: "7th and 10th OPEN".

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 commanded closed at stabilised min idle, 8600 N2, and will not be opened again in Steady state.

Transient The valve will be commanded open during engine acceleration whenever N2 is below the transient closing speed. Thus during an acceleration from min "idle to max" speed the valve will be opened and will remain open until the speed passes the transient closing speed. If the acceleration is to a speed below the transient closing speed the valve will remain open until the acceleration timer expires (30 seconds). During decelerations the valve will be commanded open whenever N2 is below the transient opening speed. The valve remains open until the deceleration ceases and a deceleration time, 2 seconds, expires. The transient regime is slightly modified for operation above 15000 ft but operates in the same way.

Surge / Reverse If the engine is operating in reverse thrust operation is the same as Transient but different speeds apply. In the event of an engine surge the valve will be commanded open, if the speed is below the open speed, and will remain open until the engine restabilises.



75-00-14


Condition

N2C26


7A status

7B status

7C status

10 status

SS ONLY

SS ONLY

SS & TR

SS ONLY

open - closes on reaching idle closed* opens on detection of acceleration, then closes at mid-power

open - closes before reaching idle closed

starting

<8623

open

idle/taxi

8623

open

open - closes before reaching idle closed

8623
open>closed

closed

closed

closed

closed*

closed

take off acceleration

closed

Begin T/O 12100, 90% derated 12044, 80% derated 11965 Begin 11869, Mid 12142, End 12294 ~12100

closed

closed

closed*

closed

closed

closed

closed*

closed

end of cruise deceleration

12000
closed>open

closed

opens on detection of deceleration, then closes

closed

top of descent mid descent end of descent approach touchdown

10819 10211 8509 9085
open open open open open open (if N2C26 below certain threshold) open open (if N2C26 below certain threshold)

closed closed closed closed closed

take off (including derates) climb cruise (mid)

reverse

12135

idle/taxi

<8623

surge recovery

NA

closed closed closed

closed* closed* closed* closed*,*** closed* open (if N2C26 below certain threshold) closed* open (if N2C26 below certain threshold)

closed closed closed closed closed closed closed 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, landing runway altitude, airport traffic, typical mission etc. the EEC 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



75-00-15



Figure 7: HBV OPEN/CLOSED Schematic



75-00-16



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



75-00-17



Figure 8: Bleed Control Valve Solenoids



75-00-18

Training Manual A319/A320/A321 Handling Bleed Valve Malfunctions


Figure 9: HDLG Bleed Valves Malfunction Tables

A engineering order (010169) is released to cover this problems. 7TH / 10TH STAGE HANDLING BLEED VALVES STICKING Hung starts or starting stalls experienced due to 7th and 10th stage handling bleed valves failing to open or close. The consequences of the malfunction of one or more handling bleed 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.



75-00-19


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.

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 air valve open at all engine operating phases except cruise. The valve incorporates 2 micro switches for transmitting valve position to the E.E.C channel A & B.

Purpose The make up air discharges into the area around No 4 bearing housing and supplements the normal airflows in this area and increases the cooling flow passing to the H.P. turbine, stage 2.


The "fail safe" position is valve open, solenoid de-energised. Figure 10: HPC 10 Cooling Tubes 4 Off

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 supplements the normal airflows in this area and increases the cooling flow passing to the H.P. turbine, stage 2. The cooling air used is taken from the 10th stage manifold, and is controlled by a two position pneumatically operated valve. The valve position is controlled by the E.E.C. as a function of corrected N2 and altitude.

Operation Signals from the E.E.C. will energise / deenergise the solenoid control valve. This directs pneumatic servo supplies to position the 10th stage air valve to the open / close position.



75-00-20



Figure 11: HP Turbine Cooling Air Schematic



75-00-21



Figure 12: Stage10 to HPT Air Control Valve



75-00-22



Figure 13: Air Systems Schematic



75-00-23

Training Manual A319/A320/A321 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.

Position C

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.


LPT ACC is closed.

This position represents a typical take off condition. This position is altitude dependent. • 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 position 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 piston moves to point A. LP valve will be partially open (-44 deg) The actuator piston remains at this point at all defective conditions. (HP valve closed)

The 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 effect 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.



75-00-24



Figure 14: Turbine Cooling Control Schematic



75-00-25

Training Manual A319/A320/A321 HPT / LPT Active Clearance Cont. Sys.


Figure 15: ACC Valve

The HP / LP Turbine Active Clearance Control (ACC) system uses fan air to cool the HP and LP cases for blade tip clearance control in order to improve engine performance and maximize the turbine cases life time. Fan air is drawn from a common HP / LP turbine ACC air scoop in the fan duct. This air is divided into HP and LP cooling air and passes through individual short ducts to the Active Clearance Control Valves which direct air for both HP and LP turbine case cooling. The HP Turbine Clearance Control Valve is equipped with 4 plugs in the valve vane. This plugs can be removed according to a service bulletin to allow a permanent cooling of the HP turbine. In case of a valve removal / installation the same configuration must be provided on the new valve. If the plugs must be removed, there is a storage bracket provided on the actuator rod. Do not throw the plugs away !



75-00-26



Figure 16: LPT / HPT Active Clearance Control Valve



75-00-27



HPT / LPT Cooling Manifolds HP Turbine Manifold The assembly consists of a left and right hand tube assemblies which are a simple push fit into the manifold. Air outlet holes on the inner face of the tubes direct the air onto the HP turbine casings.

LP Turbine Manifold The assembly consists of a upper and lower tube assemblies with integral manifolds, both ends of the cooling tubes are sealed. Air outlet holes on the inner surfaces direct the air onto the LP turbine cases.



75-00-28



Figure 17: HPT / LPT Cooling Manifolds



75-00-29



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 exhaust located an the bottom rear centre line of the fan cowl doors.

Zone 2 Ventilation The ventilation of Zone 2 is provided by air exhausting from the active clearance control (A.C.C.) system around the turbine area. The air circulates through the core compartment and exits through the lower bifurcation of the "C" ducts.

Ventilation during Ground Running During ground running local pockets of natural convection exist providing some ventilation of the fan case - Zone 2. Zone 2 ventilation is still effected in the same way as when the engine is running.



75-00-30



Figure 18: Nacelle Ventilation



75-00-31



Figure 19: Classic Lower ECAM Indication

75-41 Nacelle Temperature Nacelle Temperature General

14

The Nacelle Temperature Sensor has a Measurement Range of -54°C to 330°C

/),&),4%2

This Signal is fed to the EIU which Transforms 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 engine starting mode and one of the two temperatures reaches the advisory threshold.

03)

#,/'

#,/'

&&),4%2

#

.!# #

A advisory indication will be created on the engine system page when the temperature reaches approx. 300 - 320°C. On enhanced aircraft the nacelle temperature indication is permanently displayed.

#,/'

#,/'

Figure 20: Enhanced Lower ECAM Indication

03)

#

6)" .

.

#,/'

.!# #

4!4 3!4 )3!



# # #

',/!$

'7

+'

(

75-00-32



Figure 21: Nacelle Temperature Sensor



75-00-33





75-00-34



76 Engine Controls - V2500A



76-00-1


76-00 Engine Controls

Power Plant V2500A 76-00 Engine Controls

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:

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, each of which sends the thrust lever position to the Electronic Engine Control. The extraction current for the resolvers 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 disconnect 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 revers thrust laching lever must be lifted.



76-00-2



Figure 1: Engine Thrust Lever Control

ENGINE THRUST LEVER CONTROL AUTOTHRUST DISCONNECT PB REVERSE THRUST LATCHING LEVER

THRUST LEVER

REVERSE THRUST LATCHING LEVER

MECHANICAL BOX

THRUST CONTROL UNIT

FMU – FUEL METERING VALVE


CHANNEL A

EEC CHANNEL B

RESOLVER 1 RESOLVER 2


76-00-3

Training Manual A319/A320/A321 Bump Rating Push Button(A1 Engined Aircraft only)


Figure 2: Bump Push Bottons

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 generate 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 designated operator missions. Use of the bump rating must be recorded. This information is for tracking by maintenance personnel.



76-00-4



Figure 3: Flat Rated Thrust Control and Modification Inputs



76-00-5



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 engine 1(2) throttle control lever and to the engine 1(2) throttle control unit by means of rods. The artificial feel unit is a friction system which provides a load feedback to the throttle control lever. This artificial 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 transmitted to the to the artificial feel unit and to the throttle control unit. The linear movement of the throttle levers is transformed into a rotary movement at the bellcrank which turns about the friction brake assembly shaft. This movement rotates a toothed quadrant integral with the shaft. This toothed quadrant causes inverse rotation of a gear equipped with a disk which has four detent notches. Each notch corresponds to a throttle lever setting and is felt as a friction point at the throttle levers.



76-00-6



Figure 4: Mechanical Boxes

MECHANICAL BOX(ES)

FRICTION BRAKE ASSEMBLY

An adjustment screw is provided at the lower part of each mechanical box to adjust the artificial feel.

GEAR ASSEMBLY

MECHANICAL BOXES

CASING RIGGING POINT

ADJUSTMENT SCREW BELLCRANK ASSEMBLY

FRICTION ADJUSTMENT SCREW


LEVER ASSEMBLY

DETENT FORCE ADJUSTMENT


76-00-7



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 control 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 assembled 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 engine operation after a failure or a complete loss of the throttle resolver position signal.



76-00-8



Figure 5: Thrust Control Units

CROSS SECTION

3 COUPLED POTENTIOMETERS ELECTRICAL CONNECTORS

C

C

CONTROL LEVER

C

RESOLVER

TOOTHED SEGMENTS

RIGGING POINT

THRUST CONTROL UNIT(S)

3 COUPLED POTENTIOMETERS

– 2 units Each unit consists of : – 2 resolvers – 6 potentiometers.


C

ONE RESOLVER

CONNECTORS


76-00-9



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



76-00-10



Figure 6: Thrust Control System Rigging



76-00-11



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



76-00-12



77 Indicating - V2500A



77-00-1


77-00 Engine Indicating Presentation

• •

Indication General

Power Plant V2500A 77-00 Engine Indicating Presentation

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

Primary Engine Display

• •

The primary engine parameters listed below are permanently displayed on the Engine and Warning display (E/WD): • Engine Pressure Ratio (EPR) • Exhaust Gas Temperature (EGT) • N1 (low rotor speed) • N2 (high rotor speed) • FF (fuel flow)

Some engine parameters also displayed on the CRUISE page

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 becomes amber. Only in case of certain system faults and flight phases a warning message appears on the Engine Warning Display.

Secondary Engine Display The lower display shows the secondary engine parameters listed below. The engine page is available for display by command, manually or automatically during engine start or in case of system fault: • 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



77-00-2



Figure 1: Engine ECAM Indications

Upper E/WD

1.4

1.2

1.6

1. 296

1

4

8

464

EPR

1.6

1. 296

1

10

4

1.4

1.2

4

EGT C

82. 6 89. 0

N1 % N2 %

10

1

ENG

FADEC GND PWR

ON


S

82. 6

0

FLAP

F

89. 0

0

MASTER 1

ENG

ON

ON

25 0

VIB (N2) 0. 3 0. 3

14 .5 300

PSI

166

166

0 C

140

PSI 35

Attention Gatters

MASTER CAUT

VIB (N1) 1. 6 1. 6

215

AB

TAT +10 SAT +10

ENG Control Panel

2

14 .5

140

T.O INHIBIT IGNITION LDG LT

MASTER WARN

QT

300

2

ENGINE

OIL

25

3040 6700 KG

FOB:

F.USED KG

207

56 C

KG/H

3040

8

T.O AUTO BRK SIGNS ON SPLRS ARM FLAPS T.O T.O CONFIG NORMAL

Overhead Maint. Panel

1. 296

FLX

FF

464 4

Lower ENG System Display

ENG 1

OFF

CRANK

MODE NORM

0 PSI C C

MASTER 2

ENG 2

FIRE

FIRE

FAULT

FAULT

1

2


H 25

GW

54700 KG

RH lower Overhead Panel

ENG

O ON OFF OFF F

IGN START

03

1

ON

MAN START

ENG

2

1

ON

ON

N1 MODE

2

ON

77-00-3


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 channels performs this operation independently. 1. Actual EPR Actual EPR is green. 2. Cyan EPR command arc (transient) from current EPR pointer to EPR command value. is only displayed with A / THR engaged. 3. EPR TLA (white circle) Predicted EPR corresponding to the thrust lever position. 4. EPR max (thicker amber mark) It is the limit value of EPR corresponding to the full forward thrust lever position. 5. REV indication Appears in amber when one reverser is unstowed or unlocked or 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 associated EPR rating is displayed in blue. In MREV no EPR value is displayed.

•


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 amber 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 displayed whatever the thrust lever position between IDLE and FLX / MCT.

If FLX mode is selected, the flexible take off temperature in C, selected through the FMS MCDU’ s, is displayed. For FLX mode indication the ADIRU‘s must be switched on. The temperature value is displayed in green and the C is displayed in blue. If a failure occurs on any indication displayed, the analog indication is replaced by amber crosses, the analog indicator and the marks on the circle disappear, the circle becomes amber.

Thrust limit mode is displayed in digital 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 receiving the most advanced thrust lever position).



77-00-4



Figure 2: EPR Indication - Upper ECAM Display Unit

3

4

2 5

1.2

1.6

REV

1. 296

1

4

8

464 4

6

1.4

10

82. 6 89. 0

EPR

1.4

1.2

1.6

1. 152

1

1

4

EGT C N1 % N2 %

TOGA

8

1. 520

MCT

82. 6

x

T.O AUTO BRK SIGNS ON SPLRS ARM FLAPS T.O T.O CONFIG NORMAL


35 C

OR

10

89. 0

1. 503 OR

464 4

FLX


CL MREV

OR OR

T.O INHIBIT IGNITION LDG LT

77-00-5



EPR System Components P2 / T2 Sensor The P2 / T2 sensor is located near the 12 o’clock position of the inlet cowl. It measures total pressure and temperature in the inlet air stream of the engine forward of the engine front flange. The dual output total temperature measurement is accomplished by two resistance-sensing elements housed in the P2/T2 sensor body. Each channel of the Electronic Engine Control (EEC) monitors one of these resistance elements and converts the resistance measurement to a temperature equivalent. The total air pressure is carried via pressure tubing to the pressure sensor located in channel A of the EEC. The P2 / T2 sensor has an anti-icing function accomplished by a single heating element internally bonded to the sensor. The heater is a hermetically sealed, coaxial resistance element brazed internally to the sensor casting. Aircraft power, which is used for the heater, is switched on and off by the EEC depending on TAT (< 7,2 °C heater "ON"), via the relay box. In case of loss of P2 / T2 heating, an automatic reversion from EPR mode to unrated N1 mode occurs.

P4.9 Sensors THE P4.9 SENSOR AND MANIFOLD HAS THREE PROBES WHICH MEASURE THE TOTAL PRESSURE OF THE EXHAUST GAS STREAM. Struts 4, 7 and 10 contain the pressure sensing ports. Each sensing point contains eight radial pressure sensing ports which are combined to yield an average pressure. The resulting average radial pressure value from each strut is then plumbed into a manifold which provides an overall turbine exhaust pressure average (P4.9). A tube from this manifold is connected to the Electronic Engine Control (EEC channel A). A pressure transducer located within the EEC converts the average pressure at station 4.9 into a useable electronic signal (proportional to pressure) that can be processed and used by the EEC to control the engine.



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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

.%840!'%



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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 through corresponding DMC’ s – MCDU action or by the next T/ O.

Figure 6: EGT Indication



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Training Manual A319/A320/A321 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 channel, from where two indepent signals are send to the EEC. Each signal is an average of the four probes.


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Figure 7: EGT System



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77-10 Power

ENG

N1 and N2 Indication 1

N1 Indication The low pressure rotor speed signal is used in the EEC for engine control computation 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) appears when in rated N1 mode. N1 rated MODE can activated automatically or by switching the N1 MODE switch at the overhead panel (close to the ENG MAN START switches). Both engine 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 displayed 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.

ON

MAN START

ENG

2

1

ON

ON

N1 MODE

2

ON

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 during engine start sequence up to 43 % (starter cut out). Turns red if N2 exceeds 100 % and a red "X" appears. The red "X" will disappear through corresponding DMC’s - MCDU action or by the next T/O. General: A failure title will be displayed on E / WD on the MEMO display. If a failure occurs on any indication displayed, the analog indication is replaced by amber crosses, the analog indicator and the marks on the circle disappear, the circle becomes amber.



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Figure 8: N 1 and N2 Speed Indication



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31 Indicating

Should an exceedance occur, the DMC memorizes in its BITE memory the maximum value reached during the Last Flight Leg.

Max Pointer Reset (N1, N2 & EGT)

The values of the N1, N2, EGT red lines and transitory overlimit values are stored in 2 independent tables, one per engine.

Monitoring of the Relevant Display of the Engine Parameters N1, N2, EGT, and FF indications of both engines are monitored internally and externally. The DMC compares the N1 signal received from the EEC 1 with the feedback signal which reflects the displayed position of the N1 needle -

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.

In order to grant dissimilarity with the engine 2 monitoring process the DMC compares the N1 signal from the EEC 2 with the feedback signal representing the N1 digital value.

N1 Red Line Exceedance

The same applies to the EGT parameters indications, but with the displayed position of the engine 2 EGT needle and the engine 1 EGT digital feedback value.

The N1 red line is represented by an arc shaped red ribbon situated at the end of the scale.

As for the N2 and FF parameters, the DMC compares the direct signal from the EEC with the displayed digital value.

If the N1 actual value exceeds the N1 red 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.

In case of detected discrepancy, a CHECK amber message is displayed just below the relevant parameter indication. In addition the FWC’s perform an external monitoring between the feedback signals (that correspond to the displayed values and the signets that are directly received by the FWC’s from the EEC‘s Should a discrepancy occur, for one or more parameters, a CHECK amber message 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 INSTRUMENTS. The menu for the EIS 1,2,3, (DMC 1,2,3) must be selected.

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 same 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.

The memory cells which store the possible exceedance are reset either by pressing 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 indications of both engines.



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Figure 9: Max Pointer Reset



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77-10 Power


Figure 10: N1 Speed Sensor

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 below. • One sensor provides the Engine Vibration Monitoring Unit with N1 analog signals (trim balance sensor). • The N1 electrical harness tube goes through the inner strut of the no. 3 strut of the intermediate structure and to the terminal block. The electrical leads from each sensor goes through the N1 tube and is connected to the terminal block. • For the fan speed sensors, one turn on the LP shaft causes 60 teeth on the phonic wheel to pass its sensor. For the trimbalance sensor, one slot in the phonic wheel passes the sensor one time for one turn. • The EEC speed sensors have two pole pieces compared to the trimbalance sensor 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.



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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 reliability by the elimination of splines, bearings or similar parts which can deteriorate or fail. The rotor is mounted directly on the gearbox output shaft and the stator is bolted to the gearbox housing. The alternator provides two identical and independent power outputs, one for each channel of the EEC. • It comprises two stators (one power and one speed) and a rotor. • Is driven from the main accessory gearbox • Consists of a magnetic rotor running in a stator. The stator has four independing windings, two of which provide three phase frequency AC electric power to respectively channel "A" and "B". The third winding provides a single phase AC analog signal proportional to N2 for the Engine Vibration Monitoring System. The forth winging provides a dedicated N2 signal to Channel "A" of the EEC. • The N2 windings gives an analog signal through the cockpit for ECAM indication. The stator and rotor are sealed from the gearbox by a shaft seal. If a shaft seal failure occurs and the alternator fills with engine oil, the alternator will continue to function normally. To maintain the temperature of the dedicated alternator at an acceptable level the alternator incorporate an integral cooling air manifold using fan air.



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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 changeover 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 and right engines are displayed on the engine and cruise pages.



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Figure 13: Vibration Indication

VIBRATION indications: THE VIBRATION INDICATIONS OF THE LP AND HP ROTORS ARE DISPLAYED IN GREEN. PULSING ADVISORY ABOVE 5 PULSING ADVISORY ABOVE 5

VIB 0.8 VIB 1.2

0.8 0.8

N1 0.9 N2 1.3

1.2 1.2 140

160

80

80

Powersupply 115V AC

VIB SENSOR A

SDAC1 VIB SENSOR B

SDAC2 CFDIU Ded. Gen.



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Engine Vibration Monitoring Unit (EVMU)

Power Supply

Description

The EVMU is supplied with 115V/400Hz by the busbar 101XPA, through the circuit breaker 1EV.

The signal conditioner is composed of: • 2 channel modules • 1 balancing module • 1 data processing module • 1 power supply module.

Built in test equipment (BITE) maintenance and fault information The equipment contains a BITE system to detect internal and external failure.

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 component 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.

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 system are checked: • N1 and N2 NB velocity • unbalance data • N1 and N2 tacho frequencies • accelerometer signals. Any detected failure is stored in the non-volatile memory with GMT, the date and other reference parameters.

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 other modules with the necessary voltages.



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Figure 14: EVMU Schematic



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Components The vibration transducer including two independent 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 detailed 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 check 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 balancing 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 information, 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 shall 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 3132-00.



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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 aircraft 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 automatically 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 analysis 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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Power Plant V2500A 78-00 Reverser System

78-00 Reverser System Introduction Description The thrust reverser comprises a fixed inner and a movable outer (translating) assembly. 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

Inadvertent Stowage/Deployment

General

In either case the LVDT sensors would detect a movement the EEC would execute auto-restow or auto-redeploy.

The thrust reverser is actuated in response to signals from the Engine Electronic Control (EEC). Selection of either stow or deploy from the cockpit generates a signal to the engine EEC which in turn, supplies signals to the thrust reverser hydraulic control unit.

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.

Thrust Reverser Deployment

Pressure is returned to the system and with the directional control valve in its stow position the reverser is returned to its stowed condition.

Thrust reverser deployment is initiated by rearward movement of the reverser lever which inputs a signal, via a dual resolver, to the EEC.

Following auto-restow the isolation valve would remain energized for the remainder of the flight.

The EEC supplies a 28 volt signal to the isolation valve and directional control valve solenoids mounted in the HCU.

If the reverser travel exceeds 15% of its travel from the fully stowed position then the EEC will command idle.

The supply of the signal to the directional control valve solenoid is also dependent if aircraft is on ground (weight on wheels) and upon the closure of the aircraft permission switch (T/R inhibition relay) in that line. This switch is closed by the Throttle 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.

Following restow, full power is again obtainable.

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 rearward 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 active 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 signal to the directional control valve solenoid 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 isolation 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 ensure full lock engagement and completion of the stow cycle.


When auto redeploy is initiated to counteract inadvertent stow, the EEC will command 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 aerodynamic loads on the translating sleeves which will normally be sufficient to deploy the reverser. If the reverser travel exceeds 22% of its travel 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 consists 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

A/C on GND from EIU >50%

AND



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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 actuators. 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



78-00-9

Training Manual A319/A320/A321 Thrust Reverser Independent Locking System


Figure 5: Shut-Off Valve

**ON A/C 116-199,

General An independent locking system is designed to isolate the thrust reverser from the aircraft hydraulic system. This system consists 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 aircraft hydraulic system. The filter is a flowthrough 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 necessary to replace the filter element. Two spring-loaded magnetic pistons keep the pop out indicator in retracted position. The lower magnetic piston monitors the differential 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 approximately 2 mm, the upper magnetic piston spring overcomes the magnetic force and drives the pop-out indicator from its 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.



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



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Training Manual A319/A320/A321 Reverser Hydraulic Control Unit


Pressure Switch

Reverser Hydraulic Control Unit (HCU) General The hydraulic control unit controls hydraulic fluid flow to the thrust reverser actuators. 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).

The pressure switch provides signals to the EEC to indicate when there is hydraulic 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 filter the fluid supply from the aircraft hydraulic system. The filter is a flow through cartridge type filter. The clogging indicator 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 possible to shut the hydraulic supply to the reverser 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!

Isolation Valve The solenoid operated isolation valve isolates the thrust reverser actuation systems from the remaining hydraulic network on the engine. The isolation valve solenoid is a dual coil valve solenoid connected to both channels of the EEC. The isolation valve is in the closed position 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 valve directs high pressure hydraulic fluid 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.



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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 the cockpit, the EEC ensures that deployment is permitted. In that case, the electrical power (28VDC) is sent to the isolation 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 actuators, and then extending the actuators. 3. As soon as both lock sensors indicate unlocked for more than 0.2 seconds (indicating that translating sleeves are “unlocked sleeves” signal is sent by these sensors to the EEC. In the cockpit an amber REV indication is displayed in the middle of the EPR dial or the ECAM display unit. 4. Each translating sleeve arriving at 95 percent 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 remains energized to maintain the hydraulic pressure in the actuators to prevent 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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Figure 9: HCU Deploy Sequence



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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 than 95% deployed, REV indication changes to amber. 2. Hydraulic pressure is supplied to the rod end of the actuator, the head is connected 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) which send the "stowed sleeves" information to the EEC. The REV indication disappears. 4. The actuators move until stowing is complete and the lower actuator locks are engaged after which the isolation valve solenoid is de-energized and the reverser is locked in the forward thrust mode position.



78-00-18



Figure 10: HCU Stow Sequence



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Command Limitation If the Linear Variable Differential Transducers (LVDTs) sense an uncommanded movement of the thrust reverser: • From the stowed position, the EEC commands an automatic stowage • From the full deployed position, the EEC commands an automatic deployment.

Auto-restow In FWD thrust, if the EEC detects any 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 HCU remains energized for the rest of the flight. In FWD thrust, if the EEC detects 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 counteract inadvertent stow, the EEC will command the isolation valve to close and maintain it closed until FWD thrust has been reselected. The air aerodynamic load on the translating sleeves will normally be sufficient to redeploy the thrust reverser. In reverse thrust, if the EEC detects any un commanded movement greater than 22% from full deploy, it commands engine idle power.



78-00-20



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 pressure 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 actuators 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 sleeve 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.



78-00-22



Figure 12: Flexible Drive Shafts



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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 Transformer (LVDT) to indicate piston position, and thus translating cowl position, to the EEC. All actuators use hydraulic snubbing at the end of the deploy stroke to slow down the actuators over the final part of the deploy stroke. All actuators also incorporate the necessary deploy stroke mechanical stops.

Upper Nonlocking Actuator The two upper actuators are identical and in conjunction with the two lower locking actuators, control movement of the fan reverser translating elements in response to hydraulic inputs from the HCU.



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Figure 13: Upper Nonlocking Actuator



78-00-25

Training Manual A319/A320/A321 Lower Locking Actuators The two lower looking actuators are identical and in conjunction with the two upper actuators, control movement of the fan reverser translating elements in response to hydraulic inputs from the hydraulic control unit (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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Figure 15: Lower Locking Actuator



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Thrust Reverser Manual Deploy / Stow Manual Deploy/Stow The thrust reverser may be deployed/stowed manually for maintenance - troubleshooting 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 deactivated 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 bypass 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.



78-00-28



Figure 16: Reverser Manual Operation



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Training Manual A319/A320/A321 Thrust Reverser Deactivation


Figure 17: Deactivation Pin

Deactivation The procedure is summarised below, the full procedure is described in the MM 7830-00 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".



78-00-30



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 system will be tested. (Green ENG 1, Yellow END 2). All the test steps are written 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. Returning the throttle to the FWD idle position will restow 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.



78-00-32



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



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FADEC T/R Test (Fault Detected)



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Training Manual A319/A320/A321 FADEC T/R Test (NOT O.K.)


Figure 20: 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 Throttle Lever to reverse) the test will be interrupted and a new Test must be initiated. The duration of the complete T/R operational Test (opening & closing) is limited to 60 seconds. If this time is exceeded the test will be interrupted and a new Test must be initiated.



78-00-35





78-00-36



79 Oil - V2500A



79-00-1


79-00 Oil System

Power Plant V2500A 79-00 Oil System

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 compartment to protect the seals by maintaining low pressure differentials across compartment walls and minimizes air leakage into the bearing chamber.

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 disconnectors on the oil tank. These ports are used to refill the oil tank promptly and precisely by allowing the airlines to quick-connect a pressurized oil line and a drain line.

The scavenge valve pressure transducer senses the pressure present in the scavenge line upstream of the scavenge valve and supplies a signal to the EIU. A pressure relief valve at the filter housing limits pump discharge pressure to approximately 450 psi to protect downstream components.

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 bearings. Except for the No 3 bearing damper and the No.4 bearing compartment, the pressure supplied to each location is controlled by a restrictor. There is a ”last chance” strainer at the entry of each compartment to prevent blockage by any debris / carbon flakes in the oil. The scavenge oil is then piped, either directly or through the de-oiler to the 5 stage scavenge pumps. There is a disposable cartridge type scavenge filter at the outlet of the scavenge pumps before returning to the oil tank. A valve allows oil to bypass the scavenge filter when the filter differential pressure exceeds 20 psi. A differential pressure warning switch. set at 12 psi gives cockpit indication of impending scavenge filter bypass. The oil pressure is measured as a differential between the main supply line pressure, upstream of any restrictors, and the pressure in the No.4 bearing compartment scavenge line, upstream of the two position scavenge valve. A low pressure warning switch, which is set for 60 psi, is provided in the main oil line before the bearing compartments and after the ACOC and FCOC at the same tapping 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



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Figure 1: Oil System Schematic 01



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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.



79-00-4



Figure 2: Oil System Schematic 02 3

4002EN XMTR-OIL QTY ENG 1 (ENG 2)

A

OIL QTY

121VU212 PRESSURISING

1 2

A

VALVE

TO SENSOR 4006KS

1 TO 9VDC FOR 0 TO

3 4

27,5 QUARTS 0,5 TO 0,6

A

A B

A RESERVE SPILL RESTRICTOR

QUARTS

ENG FUEL

PRESSURE TRIMMING

DEAERATOR

AIR COOLED OIL COOLER

VALVE

A A B

FUEL COOLED OIL COOLER

A

PRESSURE

RELIEF VALVE

OIL TANK

FAN AIR

STRAINER

POWER SUPPLY

1EN1 (1EN2) ENG 1 (ENG 2)

POWER SUPPLY

101PP (202PP) BUS 1 (BUS 2)

28VDC

PRESSURE

A

C

D E

1 TO 9VDC FOR 0 TO 400PSIA ^1 1PSIA

D A

P

FILTER SUCTION

4005EN TRANSDUCERNO.4 BEARING PRESSURE

PRESSURE PUMP

ENG 1 (ENG 2)

EIU

TO SENSOR 4007KS

101PP (202PP) BUS 1 (BUS 2)

28VDC

ANGLE GEAR BOX

A 3 2EN1 (2EN2) ENG 1 (ENG 2)

A

OIL / PRESS 121VU212

EXTERNAL GEAR BOX

1 2

POWER SUPPLY

E

1 TO 9VDC FOR 0 TO 400PSID 1PSID

4 3

A

F G

A

1 P STRAIN GAGE

RESTRICTOR

N° 1

2

3

BEARING COMPARTMENT

4

5

VENT AIR / OIL

BRG COMPT

BRG COMPT

OIL PRESSURE

P

OIL SCAVENGE

A

3 UNSD 2 H A

60PSID

4003EN XMTR OIL PRESS

4000EN SW LOW OIL PRESS

ENG 1 (ENG 2)

ENG 1 (ENG 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.

1KS1 EIU 1 85VU127 REED SWITCH OPENS WHEN P 210PSID

P 12

A 3

PSID

A

2 UNSD 3

CHIP DETECTOR STRAINERS

A

SCAVENGE PUMPS

4004EN THERMOCOUPLE

J

2

DE - OILER

A

SCAVENGE-OIL TEMP ENG 1 (2)

P

1

1KS2 EIU 2 86VU127

4006EN INDICATORPOS NO.4 SCAVENGE VALVE -60 TO 250°C

ENG 1 (ENG 2)

A

4001EN SW SCAVENGE FILTER DIFFERENTIAL PRESS

5 4 6 2 1 3

ENG 1 (ENG 2)

SCAVENGE

FILTER

TWO - POSITION SCAVENGE VALVE UNSD

DE - OILED AIR OVER - BOARD

K

A L



79-00-5



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 jets 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 sealing air to escape to the de oiler.



79-00-6



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 compartment forcing the air and oil out. The air and oil passes through the no.4 bearing scavenge valve and then into the de oiler.



79-00-8



Figure 4: Centre Bearing Compartment No.4



79-00-9



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 bearing 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 compartment, which is characterized by tail pipe smoke, tail pipe fire, high oil consumption and/or oil wetness in the LPT, all of which cause maintenance disruption. Oil feed tube blockage causes oil starvation of the number 5 bearing, which can result in bearing damage.



79-00-10



Figure 5: Rear Bearing Compartment No.5



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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 should be checked and oil added, if necessary, 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.

Figure 6: De-arator

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 probe 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.



79-00-12



Figure 7: Oil Tank



79-00-13



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 element. The pressure filter housing is installed at the oil pressure pump. The pressure filter housing incorporates a pressure priming connection and a antidrain 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



79-00-15



Air Cooled Oil Cooler (ACOC) Location The ACOC is mounted on the engine fan case.

Operation The ACOC is a additional oil cooler which removes heat from the engine lubricating 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 below.

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 controlled by the EEC.



79-00-16



Figure 9: ACOC Air Flow



79-00-17



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 casing, before it is sent to the bearing compartments 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 vacuum brazed tubes through which fuel passes.

Bypass Valves • •

One is an oil pressure relief bypass valve which diverts the excessive oil pressure during engine cold start. The other is a fuel filter bypass valve which ensures fuel flow in the event of fuel filter clogging.



79-00-18



Figure 10: Fuel Cooled Oil Cooler



79-00-19



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 geabox. 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.



79-00-20



Figure 11: Scavenge Pump Assembly



79-00-21



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.


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 switch 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 indication in the cockpit. The oil temperature is sensed by a dual resistor 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.



79-00-22



Figure 12: Scavenge Filter



79-00-23

Training Manual A319/A320/A321 De-oiler


Figure 13: 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.



79-00-24



Figure 14: De-oiler



79-00-25

Training Manual A319/A320/A321 No 4 Bearing Scavenge Valve


Figure 15:

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 overboard loss of vent air and to prevent deteriation 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.


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 operation 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)



79-00-26



Figure 16: No 4 Bearing Scavenge Valve



79-00-27



No 4 Bearing Scavenge Valve Description

No 4 Bearing Scavenge Valve Indicating

Purpose

The EIU incorporates three logics allowing the monitoring of the scavenge valve operation as well as a No.4 bearing carbon - seal failure

Maintains the centre bearing compartment (No 4 bearing) seal deferential pressure by controlling the venting of the compartment 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 conditions.

High flow

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 signal to one of the three EIU logics. Two EIU logics provide a warning message to the ECAM: ENG 1 (2) BEARING 4 OIL SYS. (class 2) and a CFDS message, when the valve is not in the correct position according to the sensed burner pressure. One EIU logic provides a message on the lower ECAM: Eng. 1 (2) Bearing (class 2) and a fault message is set on the CFDS (EIU menu) when the No. 4 bearing compartment pressure is to high according to the valve position and a high burner press.(possible Carbon seal failure)

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 lower 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.



79-00-28



Figure 17: No.4 Bearing Scavenge Valve



79-00-29



Engine Oil Pressure The Oil pressure is directly linked to the opening and closing of the No.4 Bearing Scavenge Valve. A closing of the valve to minimum flow (at approx. 85% N2) will restrict the return scavenge flow to the deoiler. This will result in a pressure drop, because the ratio of the pressures will change. (the oil pressure is the differential pressure of the oil pressure feed line and the scavenge line). The No. 4 compartment scavenge oil pressure range is 0 to 160 PSI. Normal operating pressure is 0-145 PSI after three minutes of stabilization at idle speed.



79-00-30



Figure 18: Oil Pressure Chart

-!)./),02%3352%.OSCAVENGEPRESSURE 03)'

&/2-).)-5-)$,%/),02%3352%#/22%#4)/.2%&%24/4(% %.').%/0%2!4)/.!,,)-)43

%.').%/),4%-0%2!452% -/4 -534"% # &

/2-/2%

-!8

-/0

-).

-/0

.



79-00-31



Oil System Pressure Sensing General The oil pressure indicating system gives a cockpit indication of the engine oil system working pressure. The indication of this pressure comes electrically from an oil pressure transmitter on each engine. • The oil pressure transmitter is bolted to a bracket on the top left side of the engine 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 engine 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.



79-00-32



Figure 19: LOP Switch and Oil Press. Transmitter



79-00-33



Magnetic Chip Detectors (M.C.D.) A total of 7 M.C.D. ‘s are used 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 deoiler 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 MCD if the seal rings are not installed. A saftey mechanism 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 seal ring could result in an in-flight shutdown of the engine because of oil leakage.



79-00-34



Figure 20: Chip Detectors



79-00-35

Training Manual A319/A320/A321 Master Chip Detector


Figure 21: Master Chip Detector Post and Pre-Mod

The master chip detector is located in the combined scavenge return linie, on the scavenge filter housing.

LOCKED INDICATION MARK

OMNI SEAL

The Master Chip Detector is accessible through its own access panel in the L/H fan cowl. 5 SEAL RING

If the master M.C.D. indicates a problem then each of the other M.C.D.‘s is inspected to indicate the source of the problem.

LOCKED INDICATION MARK

Do not try to install the MCD if the seal rings are not installed. A saftey mechanism 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 seal ring could result in an in-flight shutdown of the engine because of oil leakage. 6 MAGNETIC CHIP DETECTOR HOUSING

4 MAGNETIC CHIP DETECTOR

SBE79-0042 LOCKED INDICATION HOLE

BAYONET PIN

SLOT

LOCKED INDICATION HOLES

2 SEAL RING 3 MAGNETIC CHIP DETECTOR HOUSING 1 MAGNETIC CHIP



79-00-36



Figure 22: Magntic Chip Detectors



79-00-37



IDG Oil Servicing IDG Oil Pressure Fill A quick fill coupling situated on the transmission 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 overflow drain ensures a correct quantity of oil.

Oil Filter A clogged filter indication is provided by a local visual pop out indicator. The indicator 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 engine, 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 filling 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 operating temperatures and damage/ disconnect to IDG.



79-00-38



Figure 23: IDG Oil Servicing



79-00-39



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 flashes 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.



79-00-40



Figure 24: ECAM Oil Indication



79-00-41

Training Manual A319/A320/A321 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.


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.

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.



79-00-42



Figure 25: Basic Schematic



79-00-43





79-00-44



24 Electrical Power - V2500A



24-22-1


Electrical Power 24-22 AC Main Generation

24-22 AC Main Generation

The generator is driven at a constant speed of 12000 RPM and cooled by oil spraying.

General

Generator Control Unit Supply

Each engine drives its associated Integrated Drive Generator (IDG) through the accessory gearbox. The drive speed varies according to the engine rating.

The Permanent Magnet Generator supplies the exciter field through the Generator Control Relay and the Generator Control Unit through a Rectifier Unit.

The IDG is split in two parts: the drive and the generator.

The Generator Control Unit (GCU) supply from the aircraft network is duplicated (Back up supply).

The IDG is cooled and lubricated by the IDG oil system.

The excitation control and regulation module keeps the voltage at the nominal value at the Point Of Regulation (POR).

Generator Drive Using the variable speed input, the generator drive produces a constant speed on the output shaft via a variable ratio differential.

Generator Operation Control

The output constant speed is regulated at 12000 RPM.

The generator is controlled by the corresponding generator push button. When pressed in, if the generator speed is high enough, the generator is energized.

Speed Control

If the delivered parameters are correct (Power Ready relay closed) the Generator Line Contactor (GLC) closes to supply its network.

A mechanical governor, acting on a hydraulic trim unit, controls the differential gear in order to maintain the constant output speed.

Generator Monitoring

The differential gear also controls the oil system pumps in order to lubricate and cool the IDG components.

The FAULT light comes on when any generator parameter is not correct or when the Generator Line Contactor is open.

Control and Monitoring

During the AVIONICS SMOKE procedure, the FAULT light does not come on when the GEN1 LINE push button is set to off.

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 failure signal is sent to SDAC 1 and 2 through the Electrical Generation 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.


24-22-2



Figure 1: IDG Location



24-22-3

Training Manual A319/A320/A321 Integrated Drive Generator


DO NOT PUSH THE IDG DISCONNECT PUSHBUTTON SWITCH FOR MORE THAN 3 SECONDS.

The IDG disconnection signal is inhibited when the corresponding engine is not running.

THERE MUST BE AT LEAST 60 SECONDS BETWEEN TWO OPERATIONS OF THE SWITCH.

Figure 2: IDG Description 115V 400Hz IDG

Variable Input Speed

Disconnect Mechanism

45009120 RPM

Input Stepup Gear

Engine Accessory Gear Box

Differential Gear

Variable Unit Reset Handle

Constant Output Speed 12000 RPM

Fixed Unit

Hydraulic Trim Unit Mechanical Governor

Drive

Oil System Charge Pump Deaerator Scavenge

Generator Permanent Magnet Generator

P M G S U P P L Y T O G C U

3 Phase 400 Hz Generator

F I E L D E X C I T A T I O N

Pumps IDG 1

Oil

GEN 1 FAULT

FAULT

ELEC Panel

Generator Control Unit

OFF



24-22-4



Figure 3: IDG Oil Cooling and Warning IDG 1 GCU 1 Drive

Oil System

Oil in Temp Sensor

Oil out Temp Sensor

Charge Pressure Switch

Input Speed Sensor

Fuel/Oil Heat Exchanger

Oil Inlet Temp

G P C U

C F D I U

E G I U

S D A C 1

T° Rise Indication

Oil Outlet Temp Overheat Temp >185°C Low Pressure Low Speed IDG 1

Disconnect Solenoid

E/W Display Master Caut

S D A C 2

SC System Display

Charge Pump

Generator U S E R S

Differential Gear Hydraulic Trim Unit

Oil in Temp Sensor

Charge Pressure Switch

FAULT

Cooler Bypass Valve

Governor

Generator

IDG De-Activator

Relief Valve

Scavenge Pump Oil Filter

Oil out Temp Sensor

Fuel/Oil Heat Exchanger

Oil Sump

Pressure Fill Port

Clogging Indicator Fuel



System

24-22-5

Training Manual A319/A320/A321 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 MINUTES TO COMPLETE. FAILURE TO OBSERVE THE OVERFLOW TIME REQUIREMENTS CAN CAUSE HIGH IDG OIL LEVEL CONDITION RESULTING IN ELEVATED OPERATING TEMPERATURES AND DAMAGE TO THE IDG. Figure 5: IDG Front View

A A

B

Normal (Reset)

B A320

Red

1

OVER FULL

Yellow

ΔP Indicator Button (Silver End, Red Cylindrical Side)

2 Green Red

ADD OIL

ADD OIL

CFM-66


3

Extended


24-22-6

Training Manual A319/A320/A321 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 green band, oil servicing is required. The yellow band corresponds to the oil thermal expansion margin.

Figure 6: Servicing of IDG

Red Band

1

OVER FULL

Yellow Band Vent Valve (Vacuum)

Filter Clogging Indicator

Green Band

Oil Filter

Electrical Connectors

2

ADD

ADD

OIL

OIL

3

Red Band

Disconnect Reset Handle

ΔP INDICATOR BUTTON

Oil Level Indicator NORMAL (RESET)

Oil Out Port

DPI RESETS REFER TO APPROPRIATE DOCUMENTATION FOR DETAILS OF THE ALTERNATE DPI PROCEDURE

Oil IN Port

1 2 3 Overflow Drain Port


EXTENDED

Pressure Fill Port

Case Drain Plug


4

REMOVE IDG

DPI RESET LABEL

24-22-7



Figure 7: Servicing of IDG Step One 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.

Overflow Drain Hose Pressure Fill Hose

Step Two Remove pressure fill hose only. Install dust cap.

Overflow Drain Hose

Allow to drain the overflowdrain about 20 minutes!

Dust CAP

Step Three Remove overflow drain hose when drainage slows to drops. Install dust cap.

Overflow Drain Hose



Dust CAP

24-22-8



Figure 8: IDG Oil Filter / IDG Installation 4 3

2

Phase Lead Installation

1

Alternate Configuration

Terminal Block Stud

Square Washer Terminal Block

Generator Terminal Lead Assembly

QAD Ring Bracket Lockwire O-Rings

Bracket Tension Bolt

Tension Bolt



24-22-9

Training Manual A319/A320/A321 AC Main System The two engine generators provide the AC main generation. The AC main generation 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 supply 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 panel, control the generators 1 and 2 respectively via the GCU.

Figure 9: Main AC Distribution System



24-22-10



26 Fire Protection - V2500A



26-12-1


Fire Protection 26-12 Engine Fire and Overheat Detection

26-12 Engine Fire and Overheat Detection Fire Detectors Each engine fire detection system consists of two independent loops A and B connected in parallel to the Fire 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.



26-12-2

Training Manual A319/A320/A321 Figure 1: Engine Fire Warning/Extinguishing



Figure 2: Fire Panel/Engine Panel


26-12-3



Figure 3: Fire Detector - Pylon

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26-12-4



Figure 4: Fire Detector - Engine

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26-12-5



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 control 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 supplies 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.



26-12-6



Figure 5: General and Fire Warning



26-12-7



Figure 6: Loop Fault Warning



26-12-8



Figure 7: Detection Fault Warning



26-12-9



Figure 8: Fire Detector - Schematic



26-12-10



Figure 9: Fire Detector - Alarm and Fault States



26-12-11



Electrical Circuits Fire Detection Circuit The engine fire and overheat detection system 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 the 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 detectors, the FDU, and the indication warnings. It initiates the loops and squib tests.



26-12-12



Figure 10: Fire Detection Circuit



26-12-13



Figure 11: Fire Extinguishing Circuit



26-12-14



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 related 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.



26-12-15



Figure 12: General ... Supply Interfaces



26-12-16



Figure 13: Engine Fire Extinguisher Bottle



26-12-17



Figure 14: Squib and Low Pressure Switch



26-12-18



Figure 15: Distribution Lines Installations



26-12-19





26-12-20


Fire Protection 26-99 CFDS System Report / Test

26-99 CFDS System Report / Test



26-99-1

Training Manual A319/A320/A321 Figure 1: Fire Protection-System Report/Test



Figure 2: Engine or APU FDU-System Report/Test


26-99-2

Training Manual A319/A320/A321 FDU - Bite


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 tests performed by each FDU are: • Power up test • MCDU test • In Operation test

The power up test will be initiated if the computer power supply has been de-energized 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 aircraft on ground.

The MCDU test is identical to the power up test.

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 provided 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 failures in both loop detection circuits.

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.



26-99-3

Training Manual A319/A320/A321 Figure 4: FDU Fault Messages-Examples



Figure 5: MCDU Messages from FDU


26-99-4



30 Ice and Rain Protection - V2500A



30-00-1


30-00 Eng. Air Intake Ice Protection

Power Plant V2500A 30-00 Eng. Air Intake Ice Protection

System Control ON - (PB-Switch In, Blue)

System Description

The ON light comes on in blue. (valve solenoid deenergized). ENG ANTI ICE ON is indicated on the ECAM MEMO page.

Engine Air Intake Anti-Ice Air Source The air bled from the 7th stage of the 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 exhausts via a flush duct in the aft cavity of the intake cowl.

When the anti ice valve is open (valve position sw. NOT CLOSED), the zone controller sends a signal to the 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)

Valve

Anti ice system is OFF (valve solenoid energized).

For each Engine, hot bleed air is ducted via an ”ON/OFF” valve.

FAULT - (PB Switch In, Amber)

The valve is pneumatically operated, electrically controlled and spring loaded closed.

Fault light illuminates amber when valve not fully open.

Upon energization of the solenoid, the valve will close.

FAULT - (PB-Switch Out, Amber)

In case of loss of electrical power supply 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.

Fault light illuminates amber.

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 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.

The ”FAULT” light comes on during transit or in case of abnormal operation. When the anti-ice valve is open, the zone controller determines the bleed air demand for the Full Authority Digital Engine Control (FADEC) system. ECAM Page If at least one of the two engine air intake anti-ice systems is selected ”ON”, a message appears in GREEN on the ”ECAM MEMO” display.



30-00-2



Figure 1: Engine Nacelle A/I Architecture

7

1

FADEC

CABIN ZONE CONTROLLER


2

OPEN POSITION SIGNAL


30-00-3



System Control Schematic Figure 2: Control Schematic



30-00-4



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).



30-00-5



Figure 3: Engine Anti-Ice Duct and Valve



30-00-6





30-00-7





30-00-8



36 Pneumatics - V2500A



36-10-1


36-10 General

Pneumatic 36-10 General

HP Bleed Valve The HP bleed valve operates pneumatically and is connected by a sense line to a pressure regulator valve (PRV 4001HA).

Distribution - Description and Operation

It is springloaded closed and starts to open at 8 psi HP stage air pressure. It regulates the downstream pressure to 36 psi when open.

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.

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 energized. 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 exchanger (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 exceeds 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



36-10-2

Training Manual A319/A320/A321 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 valve (FAV). The fan air valve is springloaded 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 signal to both BMCs (Bleed Monitoring Computer) for pressure monitoring and indication. 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 system 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 overtemperature in ducts and give temperature indication.

This system is installed in the MID 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 the 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 protection 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.


36-10-3

FAN AIR VALVE CONTROL THERMOSTAT (7170HM1)

TO BMC 1 CONTROL TEMPERATURE SENSOR (6HA1) TO BMC 2

TEMP CONTROL T0 200 C CTS

BLEED PRESSURE REGULATED VALVE CONTROL SOLENOID (10HA1)

OVER BOARD

TRANSFER PRESSURE TRANSDUCER (7HA1) TO BMC 2

Pr

VENTED IF PRV IS CLOSED

(REG 44 PSI)

REGULATED PRESSURE TRANSDUCER (8HA1)

PRV

TP

Pt

VENT

TP

TP

VENT HP BLEED OVERRIDE SOLENOID ENG 1 (4029KS )

S

- 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

HP BLEED OVERRIDE SOLENOID

TO BMC 1

TO BMC 1

- 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

TEMPERATURE LIMITING THERMOSTAT

-

REGULATES 36 PSI AND CLOSES IF: 10 t h STAGE PRESS >100PSI OR 7t hSTAGE PRESS > 36 PSI OR COUPLING SENSE LINE IS VENTED (PRV CLOSE OR HP OVERRIDE SOLENOID ENERGIZED)

HIGH PRESSURE BLEED VALVE (HPV)

SERVO PRESSURE TO ENGINE NACELLE ANTI-ICE VALVE

VALVE POS. TO BMC 1+2

COUPLING SENSE LINE

HPV

TP

HIGH PRESSURE BLEED VALVE (4000HM1) REG

TP

HP th 10 STAGE

TO HYDRAULIC RESERVOIR (ENG 1 ONLY)

STAGE

IP

T

th

S

TO LH BLEED AIR SYSTEM

T

CT

PRECOOLER (7150HM)

FAN AIR VALVE (9HA1)

OPV

7

IP CHECK VALVE (7110HM)

VALVE POS. TO BMC 1+2

BLEED PRESSURE REGULATOR VALVE (4001HA1)

VALVE POS. TO BMC 1

OVER PRESSURE VALVE (5HA1)

FAV

VALVE POS. TO BMC 1

TP

ENG START SYS.

F A N

REG

36-10-4 for training purposes only Sep08/Technical Training Copyright by SRTechnics

T

36-10 General

Pneumatic Training Manual A319/A320/A321 Figure 1: Bleed Air System



Figure 2: Bleed Air System Layout



36-10-5



Figure 3: Component Location, Engine and Pylon

Z400

A

(CTS ) Bleed Pressure Regulator Valve Control Solenoid

(CT)

10HA

Fan Air Valve Control Thermostat

9HA

A

7150HM Bleed Air Precooler Exchanger

Z400

Fan Air Valve (FAV)

Wing Anti Ice

Bleed Pressure Regulator Valve (PRV)

to Starter Valve

(T)

6HA

5HA Overpressure Valve (OPV)

7170HM

Heat Exchanger Outlet Sensor

(Pt)

7HA

Bleed Transfer Pressure Transducer

4001HA 8HA (Pr) Bleed Regulated Pressure Transducer

Fan Air Inlet (Air from Fan)

4000HA HP Bleed Valve (HP)

7110HM IP Bleed Check Valve (IPC)



36-10-6

Training Manual A319/A320/A321 BMC Bleed Monitoring Computer The two BMCs monitor the operation of the HP bleed valve (close/open microswitch signals, and transfer pressure level). They receive and process the signals and transmit the information per data bus by the System Data Acquisition Concentrator (SDAC) to the ECAM system which generates the system display. The indications are: pressure, temperature 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 Interface Unit (CFDIU). The CFDIU generates maintenance information which is displayed on the Multi Function Control Display Unit (MCDU) if the MCDU MENU is selected.


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 abnormal 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.

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 operation and failure to be detected 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 temperature. 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 pressures. Additionally, they monitor ambient overheat in pylons, wings and the fuselage. 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.)



36-10-7



Figure 4: BMC Interfaces



36-10-8

Training Manual A319/A320/A321 High Pressure Bleed Valve The HP Bleed Valve is a 4 in. dia. butterfly-type valve which operates as a shutoff 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 necessary 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.


allow the HP bleed valve actuator opening chamber supply with reduced pressure air. When chamber (4) is vented to ambient the clapper (5) leaves its lower seat position 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

A319 only: A solenoid is installed on a bracket 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) increases. 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,

3

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 pneumatically.

Safety Valve Closing Chamber

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.

Opening Chamber

2

The test intake is used for checking correct valve operation on the ground by directly supplying the regulator. The HP bleed valve downstream pressure supplies the HP bleed valve actuator closing chamber through distribution clapper (3). Indeed when downstream pressure reaches the value determined by spring preloading.

Opening/Closing

1 Test Intake

reduced pressure air supplies chamber (4) of the opening/closing sub assembly though a jet to control the position of clapper (5) against its lower seat position and


6


Butterfly

Microswitch

Bleed Air Flow

36-10-9

Training Manual A319/A320/A321 Bleed Pressure Regulator Valve (PRV) (1)The PRV is a 4 in. dia. butterfly-type valve, normally spring-loaded closed in absence 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.


The test intake is used for checking correct valve operation on the ground by directly supplying the regulator. The downstream pressure supplies the actuator closing chamber through distribution clapper (4). Indeed when downstream pressure reaches the value determined by spring preloading. Figure 6: Bleed Pressure Regulator Valve 5

6

7 to Solenoid Thermostat

from HP Bleed Valve

It is controlled in closed position by crew action on: • ENG FIRE pushbutton switch • ENG BLEED pushbutton switch.

Closing Chamber

The PRV closes pneumatically in case of impending reverse flow to the engine. The Overpressure Valve (OPV) installed downstream of the PRV protects the system against damage if overpressure occurs.

4

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 valve to close at 450 plus or minus 25 deg.C.

Regulation The 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 actuator opening chamber.

2

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 contact with its seat (4), downstream pressure still supplies the actuator closing chamber despite a reduced pressure air value lower than the nominal regulation threshold.

Bleed Air Flow Butterfly Test Intake 1



Microswitch

3

36-10-10

Training Manual A319/A320/A321 Overpressure Valve (OPV)


Figure 7: Overpressure Valve

The OPV is a 4 in. dia. butterfly-type valve, whose operation is fully pneumatic. In normal conditions the valve is spring-loaded open.

Regulator Assembly

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.

Microswitch Closing Chamber

Pneumatic Actuator

Test Port

Safety Devices and Indications A

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 cannot 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)

Bleed Air Pressure



Butterfly

36-10-11

Training Manual A319/A320/A321 Fan Air Valve (FAV)


Figure 8: Fan Air Valve

The FAV is a 5.5 in. dia. butterfly-type valve, 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 temperature to 200 plus or minus 15 deg.C (27 deg.F).

A

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. A

Safety Devices and Indications B

The FAV is equipped with a test port which serves to perform an "in situ" test.

Position Indicator

A manual override serves to close the valve mechanically on the ground. Microswitch Electrical Connector

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)

B

Thermal Fuse Vent Screw Test Intake Pressure Tapping (Motive Pressure)



36-10-12

Training Manual A319/A320/A321 Fan Air Valve Control Thermostat CT


Figure 9: 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).

CT (7170HM)

(2) Detailed Description The fan air valve control thermostat contains two mains parts: • a temperature sensing element • a pressure regulator.

Chamber A Clapper

Regulation When the temperature downstream of the precooler exchanger is below the required 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 thermostat to the opening chamber of the FAV.

Pressure Reducing Valve

Chamber B to the Opening Chamber of Fan Air Valve

Air Venting

Between both values the FAV butterfly has an intermediate position.

Filter

Precooled Air Outlet

Regulating Probe



36-10-13

Training Manual A319/A320/A321 Temperature Limitation CTS


Figure 11: 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 dilatation. This cause a modification of the butterfly 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 full open and the bleed pressure limited to 17.5 psig.

Closure of bleed pressure regulator valve When the solenoid is energized, its valve 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.

Non Return Assembly

Solenoid Assembly Solenoid

Electrical Connector

Plunger

Upstream Precooler Pressure

Solenoid Valve

The Bleed Pressure Regulator Valve Control Solenoid has no direct effect on the HP Bleed Valve (HPV) operation. Figure 10: CTS Location CTS (10HA)

A

Regulator Assembly

to PRV


Air Vent

A Filter

Solenoid Assembly

Non Return Assembly

to Pressure Regulator Valve (PRV)

Downstream Precooler Pressure

Thermostat Assembly

Sensing Tube

from Precooler Upstream

Invar Rod

Attachment Plate Thermostat Assembly Note:

LH Side Shown RH Side is Symmetrical



36-10-14



Bleed Transfer Regulated Pressure Transducers Pt

Operation

The pressure transducer is a piezo-resistive type cell. It senses the bleed transfer/ regulated pressure and transforms it into a proportional current voltage.

The pressure to be measured is ducted to the transducer via a sense line. It acts on the integrated strain gage of the piezo-resistive cell to generate an electrical signal proportional to the pressure variation. The signal is transmitted to the bleed monitoring computer.

Each pressure transducer consists of: • a measuring electronic cell • an electrical connector • a pressure port. A

Tube

Z420/480

Wing Anti Ice Duct

Pylon Loop

Control Temp. Sensor 6HA1 (6HA2) CTS


TCT Z410/470

TLT Transfered Pressure Transducer Label 7HA1 (7HA2)

A Precooler 7159HM

Housing

TPT


RPT Regulated Pressure Transducer 8HA1 (8HA2)



36-10-15


Temperature Control Description and Operation To improve the operation of the engine bleed air, the temperature limitation function 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 temperature 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 Computer (BMC) generates a class 2 maintenance message ''AIR BLEED'' on the ECAM STATUS page. An associated maintenance message ''THRM (Thermostat), FAV (Fan Air Valve) or sense line'' can be seen on the MCDU.

Figure 12: Temperature Control



36-10-16


CFDS MCDU Pages


On the SYSTEM REPORT/TEST AIR BLEED page: AIR BLEED title is replaced by PEU,

On the SYSTEM REPORT/TEST menu page:

PRINT key is deleted.

AIR BLEED key is replaced by PNEU key. Figure 13:

Classic

Enhanced


SYSTEM REPORT / TEST ENG>

A319/320/321

TOILET>

A318

SYSTEM REPORT / TEST AIR BLEED
SYSTEM REPORT / TEST PNEU


ENG>

PRINT*


36-10-17





36-10-18


Study Questions IAE V2500

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 pylon 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

For training purposes only

Study Questions IAE V2500-1



b) Thrust levers between Idle and Climb c) Thrust levers in any position

b) The EEC only c) The MASTER lever to OFF position 13.Is same EEC channel always in command?

20.What power level is set if A/THR is lost

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

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

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 installed at the LP compressor case

22.In A/THR mode, will changes in the engine power cause movement of the thrust lever?

15.Which sensors are located at the diffuser case?

Answer:

Answer:

23.What type is the LP pump

16.What is the primary mode setting? a) N1 b) EGT c) EPR

Answer:

24.HP fuel system is protected against excessive pressure by? 17.How is the FF signal transmitted from the EEC to the Fuel Metering Valve (FMV)

Answer:

Answer: 25.How is icing / waxing of the fuel prevented at low temperatures 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) 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:

a) Thrust levers between Idle and FLEX / MCT






c) Wet motoring.

27.What positions the spill valve Answer:

34.Of which type is the ignition system? a) Low voltage, low energy. b) Low voltage, high energy. c) High voltage, high energy.

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

35.From where can igniters A and B be selected? a) The engine panel. b) The ECAM control panel. c) The EEC.

29.The overspeed and PRSOV are opened by? a) electrically by actuators b) pneumatically c) metered fuel pressure

36. What causes an automatic continuous relight selection? a) IGN START selected. b) Flame-out detected. c) EEC failure.

30.Fuel samples can be taken from? Answer:

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.

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 overcomes 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

38.Where is electrical power for the Ignition coming from? Answer:

32. How does the starter operate? a) Electrically. b) Pneumatically. c) Hydraulically.

75-00 Airflow 1. What is the copressor airflow control system used for?

33.In which case can the EEC abort the start? a) Automatic start and manual start. b) Automatic start only.




Training Manual A319/A320/A321 2. Which compressor valve is fully modulating?


9. What is the VSV actuator fail safe position? a) VSV actuator fully extended

Answer:

b) VSV guid vanes fully closed c) VSV actuator fully retracted

3. Do the VIGV and the VSV have a seperate actuator?

10.Is the fuel diverter valve a modulating valve?

Answer:

Answer: 4. VIGV and VSV are fitted at which stages of the HP compressor on A1 and A5? Answer: 11.Is the fuel return to tank flow valve modulating? 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

12.What flight condition inhibit the fuel return to tank flow? Answer:

6. Do all stage 7 bleed valve operate in transient condition? Answer:¨ 13.Air is supplied to the active clearance control (ACC) system by?

7. Which handling bleed valve open, should a compressor surge occur

a) stage 10 air

Answer:7A, 7C and 10th

b) stage 12 air c) fan air

8. What is the fail safe position of a handling bleed valve, should the electrical control solenoid fail? Answer:


14.Make-up air valve is closed at which conditions? Answer:



Training Manual A319/A320/A321 15.Stage 10 HPC make-up air provides?


2. Which engine parameters are diplayed on the E/WD?

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

Answer:

16.What air is used for cooling the HPC 12 air in the ACAC 3. Which engine parameters are diplayed on engine system display?

Answer:

Answer: 17.What is the cooled HP 12 air used for? Answer: Answer: 4. Which speed signal is supplied to the EVMU? a) N1 only b) N1 and N2

18.What is the fail safe position of the engine anti-ice valve?

c) N2 only

a) closed b) last commanded position c) open

5. How many probes interact with the phonic wheel? Answer:

19.When and how is the P2/T2 probe heated? Answer: 6. How many thermocouples measure the EGT a) one b) three

77-00 Indication

c) four

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


7. EGT is measured at which engine station? Answer:



Training Manual A319/A320/A321 8. Where is the P2T2 probe located?


14.What senses the movement of the translating cowls Answer:

Answer:

15.How is the movement of the translating cowls synchronised?

9. Which shaft does the air starter rotate

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

Answer:

10. How is the selection of the accelerometer sensors performed?

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.

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

17.What does the CFDIU simulate during the MCDU reverse test? a) The TLA. b) The ground logic. c) The N2 condition.

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

18.What is the purpose of the auto-restow system? a) To restow the thrust reverser immediately after detection of any uncommanded movement. b) To restow the thrust reverser after detection of any uncommanded movement of at least 10% of travel. c) To automatically restow the thrust reverser on cancellation of a deploy command.

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)?




Training Manual A319/A320/A321 19.Which important safety procedure should be performed prior any maintenance work on or under the c-duct?


6. What is the pupose of the No.4 bearing scavenge valve? Answer:

Answer:

7. Name the two temp sensors in the oil system

79-00 Engine Oil

Answer:

1. How is the oil system protected from high pump delivery pressure? Answer:

2. Has the pressure filter a bypass valve fitted? 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

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

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.









IAE V2500 Beamer

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