V2500 gf issue 01

Notice to Holders
The information in this document is the property of
International Aero Engines AG and may not be copied, or
communicated to a third party, or used, for any purpose other
than that for which it is supplied without the express written
consent of International Aero Engines AG.
Whilst this information is given in good faith, based upon the
latest information available to International Aero Engines AG,
no warranty or representation is given concerning such
information, which must not be taken as establishing any
contractual or other commitment binding International Aero
Engines AG or any of its subsidiary or associated companies.
This training manual is not an official publication and must not
be used for operating or maintaining the equipment herein
described. The official publications and manuals must be used
for those purposes: they may also be used for up-dating the
contents of the course notes.

V2500 ABBREIVATIONS
ACAC Air Cooled Air Cooler
ACC Active Clearance Control
ACOC Air Cooled Oil Cooler
AIDRS Air Data Inertial Reference System
Alt Altitude
APU Auxiliary Power Unit
AMM Aircraft Maintenance Manual
BDC Bottom Dead Centre
BMC Bleed Monitoring Computer
BSBV Booster Stage Bleed Valve
CFDIU Centralised Fault Display Interface Unit
CFDS Centralised Fault Display System
CL Climb
CNA Common Nozzle Assembly
CRT Cathode Ray Tube
DCU Directional Control Unit
DCV Directional Control Valve
DEP Data Entry Plug
DMC Display Management Computer
ECAM Electronic Centralised Aircraft Monitoring
ECS Environmental Control System
EEC Electronic Engine Control
EGT Exhaust Gas Temperature
EHSV Electro-hydraulic Servo Valve
EIU Engine Interface Unit
EIS Entered Into Service
EVMS Engine Vibration Monitoring System
EVMU Engine Vibration Monitoring Unit
EPR Engine Pressure Ratio
ETOPS Extended Twin Engine Operations
FADEC Full Authority Digital Electronic Control
FAV Fan Air Valve
FCOC Fuel Cooled Oil Cooler
FCU Flight Control Unit
FDRV Fuel Diverter and Return to Tank Valve
FSN Fuel Spray Nozzle
FMGC Flight Management and Guidance Computer
FMV Fuel Metering Valve
FMU Fuel Metering Unit
FOB Fuel On Board
FWC Flight Warning Computer
HCU Hydraulic Control Unit
HIV Hydraulic Isolation Valve
HEIU High Energy Ignition Unit (igniter box)

HP High Pressure
HPC High Pressure Compressor
HPT High Pressure Turbine
HPRV High Pressure Regulating Valve
HT High Tension (ignition lead)
IDG Integrated Drive Generator
IAE International Aero Engines
IDG Integrated Drive Generator
IFSD In-flight Shut Down
IGV Inlet Guide Vane
lbs. Pounds
LE Leading Edge
LGCIU Landing Gear and Interface Unit
LGCU Landing Gear Control Unit
LH Left Hand
LP Low Pressure
LPC Low Pressure Compressor
LPCBV Low Pressure Compressor Bleed Valve
LPSOV Low Pressure Shut off Valve
LPT Low Pressure Turbine
LRU Line Replaceable Unit
LT Low Tension
LVDT Linear Voltage Differential Transformer
MCD Magnetic Chip Detector
MCDU Multipurpose Control and Display Unit
MCLB Max Climb
MCT Max Continuous
Mn Mach Number
MS Micro Switch
NAC Nacelle
NGV Nozzle Guide Vane
NRV Non-Return Valve
N1 Low Pressure system speed
N2 High Pressure system speed
OAT Outside Air Temperature
OGV Outlet Guide Vane
OP Open
OPV Over Pressure Valve
OS Overspeed
Pamb Pressure Ambient
Pb Burner Pressure
PRSOV Pressure Regulating Shut Off Valve
PRV Pressure Regulating Valve
PSI Pounds Per Square Inch
PSID Pounds Per Square Inch Differential
PMA Permanent Magnet Alternator

P2 Pressure of the fan inlet
P2.5 Pressure of the LP compressor outlet
P3 Pressure of the HP compressor outlet
P4.9 Pressure of the LP turbine outlet
QAD Quick Attach/Detach
SAT Static Air Temperature
SEC Spoiler Elevator Computer
STS Status
TAI Thermal Anti Ice
TAT Throttle Angle Transducer
TAP Transient Acoustic Propagation
TCT Temperature Controlling Thermostat
TDC Top Dead Centre
TE Trailing Edge
TEC Turbine Exhaust Case
TFU Transient Fuel Unit
TRA Throttle Resolver Angle
TLA Throttle Lever Angle
TLT Temperature Limiting Thermostat
TM Torque Motor
TO Take-off
TOBI Tangential out Board Injector
TX Transmitter
UDP Uni-directionally Profiled
VIGV Variable Inlet Guide Vane
VSV Variable Stator Vane

V2500 GENERAL FAMILARISATION COURSE NOTES CONTENTS
PREFACE
SECTION 1 ENGINE INTRODUCTION
SECTION 2 PROPULSION SYSTEM, FIRE PROTECTION AND VENTILATION
SECTION 3 ENGINE MECHANICAL ARRANGEMENT
SECTION 4 ELECTRONIC ENGINE CONTROL
SECTION 5 POWER MANAGEMENT
SECTION 6 FUEL SYSTEM
SECTION 7 OIL SYSTEM
SECTION 8 HEAT MANAGEMENT SYSTEM
SECTION 9 COMPRESSOR AIRFLOW CONTROL SYSTEM
SECTION 10 SECONDARY AIR SYSTEMS
SECTION 11 ENGINE ANTI-ICE SYSTEM
SECTION 12 ENGINE INDICATATIONS
SECTION 13 STARTING AND IGNITION SYSTEM
SECTION 14 THRUST REVERSE
SECTION 15 TROUBLESHOOTING

© IAE International Aero Engines AG 2000
IAE V2500 General Familiarisation Introduction
IAE V2500 Line and Base Maintenance for Engineers
This is not an Official Publication and must not be used for
operating and maintaining the equipment herein described.
The Official Publications and Manuals must be used for
these purposes.
These course notes are arranged in the sequence of
instruction adopted at the Rolls Royce Customer Training
Centre.
Considerable effort is made to ensure these notes are
clear, concise, correct and up to date. Thus reflecting
current production standard engines at the date of the last
revision.
The masters are updated continuously, but copies are
printed in economic batches. We welcome suggestions for
improvement, and although we hope there are no errors or
serious omissions please inform us if you discover any.
Telephone:
Outside the United Kingdom (+44) 1332 - 244350
Within the United Kingdom 01332 –244350
Your instructor for this course is:
----------------------------------------------------------------------------
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IAE International Aero Engines AG (IAE)
On March 11, 1983, five of the worlds leading aerospace
manufacturers signed a 30 year collaboration agreement
to produce an engine for the single isle aircraft market with
the best proven technology that each could provide. The
five organisations were:
• Rolls Royce plc - United Kingdom.
• Pratt and Whitney - USA.
• Japanese Aero Engines Corporation.
• MTU-Germany.
• Fiat Aviazione -Italy.
In December of the same year the collaboration was
incorporated in Zurich, Switzerland, as IAE International
Aero Engines AG, a management company established to
direct the entire program for the shareholders.
The headquarters for IAE were set up in East Hartford,
Connecticut, USA and the V2500 turbofan engine to power
the 120-180 seat aircraft was launched on January 1st
1984.
Each of the shareholder companies was given the
responsibility for developing and delivering one of the five
engine modules. They are:
• Rolls Royce plc - High Pressure Compressor.
• Pratt and Whitney – Combustion Chamber and High
Pressure Turbine.
• Japanese Aero engine Corporation (JAEC) - Fan and
Low Pressure Compressor.
• Motoren Turbinen Union (MTU) - Low Pressure
Turbine.
• Fiat Aviazione - External Gearbox.
Note: Rolls Royce have developed and introduced the
wide chord fan to the V2500 engine family.
The senior partners Rolls Royce and Pratt and Whitney
assemble the engines at their respective plants in Derby
England and Middletown Connecticut USA. IAE is
responsible for the co-ordination of the manufacture and
assembly of the engines. IAE is also responsible for the
sales, marketing and in service support of the V2500.
Note: Fiat Aviazione have since withdrawn as a risk-
sharing partner, but still remains as a Primary Supplier.
Rolls Royce now has responsibility for all external gearbox
related activity.
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IAE V2500 Engine/Airframe Applications
The V2500 engine has been designated the ‘V’ because
International Aero Engines (IAE) was originally a five-
nation consortium. The ‘V’ is the Roman numeral for five.
The 2500 numbering indicated the first engine type to be
released into production. This engine was rated at
25000lbs of thrust.
For ease of identification of the present and all future
variants of the V2500, IAE has introduced an engine
designation system.
• All engines possess the V2500 numbering as a generic
name.
• The first three characters of the full designation are
V25. This will identify all the engines in the family.
• The next two figures indicate the engines rated sea
level takeoff thrust.
• The following letter shows the aircrafts 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 all
applications for the Boeing McDonnell Douglas MD-90
aircraft.
• The V2500-A collectively describes all the applications
for the Airbus Industries aircraft.
This is irrespective of engine thrust rating.
The number given after the alpha indicates the mechanical
standard of the engine. For example;
• V2527-A5.
The only engine exempt from these idents is the current
service engine, which is already certified to the designated
V2500-A1. There is only one standard of this engine rating
and is utilised on the Airbus A320 aircraft.
Note:
The D5 variant is now no longer in production, however
the engine is still extensively overhauled and re-furbished.
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Introduction to the Propulsion System
The V2500 family of engines share a common design
feature for the propulsion system.
The complete propulsion system comprises the engine
and the nacelle. The major components of the nacelle are
as follows:
• The intake cowl.
• The fan cowl doors.
• Hinged ‘C’- ducts with integral thrust reverser units.
• Common nozzle assembly.
Intake Cowl
The ‘pitot’ style inlet cowl permits the efficient intake of air
to the engine whilst minimising nacelle drag.
The intake cowl contains the minimum of accessories. The
two main accessories that are within the intake cowl are:
• P2/T2 probe.
• Thermal anti icing ducting and manifold.
Fan Cowl Doors
Access to the units mounted on the fan case and external
gearbox can be gained easily by opening the hinged fan
cowling doors.
The fan cowl doors are hinged to the aircraft pylon in four
positions.
There are four quick release – adjustable latches that
secure the fan cowl doors in the closed position.
Each fan cowl doors has two integral support struts that
are secured to the fan case to hold the fan cowl doors in
the open position.
C - Duct Thrust Reverser units
The ‘C’-ducts is hinged to the aircraft pylon at four
positions per ‘C’-duct and is secured in the closed position
by six latches located in five positions.
The ‘C’-ducts is held in the open position by two integral
support struts.
Opening of the ‘C’-ducts allows access to the core engine.
Common Nozzle Assembly (CNA)
The CNA exhausts both the fan stream and core engine
gas flow through a common propulsive nozzle.
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Engine
The V2500 is a twin spool, axial flow, and high bypass
ratio turbofan type engine.
The engine incorporates several advanced technology
features, which include:
• Full Authority Digital Electronic Control (FADEC).
• Wide chord fan blades.
• Single crystal HP turbine blades.
• 'Powdered Metal' HP turbine discs.
• A two-piece, annular combustion system, which utilises
segmental liners.
Engine Mechanical Arrangement
The low-pressure (LP) system comprises a single stage
fan and multiple stage booster. The booster, which is
linked to the fan, has:
• A5 standard four stages.
• A1 standard three stages.
The boosters are axial flow type compressors.
A five-stage LP turbine drives the fan and booster.
The booster stage has an additional feature. This is an
annular bleed valve, which has been incorporated to
improve starting and handling.
Three bearing assemblies support the LP system. They
are:
• A single ball type bearing (thrust).
• Two roller type bearings (support).
The HP system comprises of a ten-stage axial flow
compressor, which is driven by a two-stage HP turbine.
The HP compressor has variable inlet guide vanes (VIGV)
and variable stator vanes (VSV).
• The A5 standard has one stage of VIGV and three
stages of VSV’s.
• The A1 standard has one stage of VIGV and four
stages of VSV's.
The HP system utilises four bleed air valves. These valves
are designed to bleed air from the compressors so as to
improve both starting and engine operation and handling
characteristics.
Two bearing assemblies support the HP system. They are:
• A single ball type bearing (thrust).
• A single roller type bearing (support).
The combustion system is of an annular design,
constructed with an ‘inner’ and ‘outer’ section.
There are twenty fuel spray nozzles supplying fuel to the
combustor. The fuel is metered according to the setting of
the thrust lever or the thrust management computer via the
FADEC system.
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The FADEC system uses pressures and temperatures of
the engine to control the various systems for satisfactory
engine operation. The sampling areas are identified as
stations and are common to all variants of the V2500
engine.
The following are the measurement stations for the V2500
engine:
• Station 1 - Intake/Engine inlet interface.
• Station 2 - Fan inlet.
• Station 2.5 – LPC Outlet Guide Vane (OGV) exit.
• Station 12.5 - Fan exit/ C-Duct by-pass air.
• Station 3 - HP Compressor exit.
• Station 4.9 - LP Turbine exit.
Engine stage numbering
The V2500 engine has compressor blade numbering as
follows:
Stage 1 - Fan.
Stage 1.5 - LPC booster
Stage 2 - LPC booster.
Stage 2.3 - LPC booster (A5 Only).
Stage 2.5 - LPC booster.
Stages (3-12) - HPC Stages.
Note the HPC is a ten-stage compressor.
The V2500 engine has turbine blade stage numbering as
follows:
Stages (1-2) - HP Turbine Stages.
Stages (3-7) - LP Turbine Stages.
V2500-A1 V2527-A5
EIS May 89 Dec 93
Take-off thrust (lb) 25,000 26,500
Flat rate temp (°C) 30 45
Fan diameter (ins) 63 63.5
Airflow (lb/s) 792 811
Bypass ratio 5.4 4.8
Climb-pressure ratio 35.8 32.8
Cruise sf (lbf/lb/hr) 0.543 0.543
Power plant wt. (lb) 7400 7500

SECTION 2
PROPULSION SYSTEM, FIRE PROTECTION
&
VENTILATION

IAE V2500 General Familiarisation Propulsion System, Fire Protection and Ventilation
Propulsion System Introduction
Purpose
The propulsion system encloses the Powerplant. They
provide the ducting for the fan bypass air and provide for
an aerodynamic exterior.
Description
The propulsion system comprises of the engine and the
following nacelle units:
• Intake cowl assembly.
• The L and R hand hinged fan cowl doors.
• The thrust reverser C-ducts.
• The common nozzle assembly (CNA).
• Engine mounts for the front and rear of the engine.
• Fire protection and ventilation system.
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Airframe Interfaces
Purpose
The airframe interfaces provide a link between the engine
and aircraft systems.
Description
The following units form the interface between the aircraft
and engine:
• The front and rear engine mounts.
• The bleed air off-takes.
• The starter motor air supply.
• Integrated Drive Generator (IDG) electrical power.
• Fuel supplies.
• Hydraulic fluid supplies.
• FADEC system interfaces.
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Propulsion System Access Panels
Purpose
The propulsion system access panels provide the engineer
with quick access to the components that require regular
or scheduled inspection.
The access panels allow the removal and installation of
Line Replaceable Units (LRU’s) during maintenance
activities.
Description
The access panels provided on the propulsion system are
as follows:
Engine Left Hand Side
Fan cowl door
Oil tank service door.
Master magnetic chip detector panel.
Zone 1 Ventilation Outlet Grille for the Fan Case.
Thrust reverser C-duct
Maintenance access panels for the thrust reverser
hydraulic actuators.
Translating cowl lockout pins.
Engine Right Hand Side
Intake cowl
Interphone jack.
Anti icing outlet grille.
P2/T2 probe access panel.
Fan cowl doors
Air-cooled oil cooler outlet.
Starter motor air valve access panel.
Zone 1 Ventilation Outlet Grille for the Fan Case.
Breathers overboard discharge.
Thrust reverser C duct
Maintenance access panels for the thrust reverser
hydraulic actuators.
Translating cowl lockout pins.
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Propulsion System Core Engine Access
Purpose
The propulsion system can be opened to allow access for
engineers both to the fan case and core engine.
Description
Fan cowl doors
The fan cowl doors are hinged from the aircraft strut at the
top and are secured by four latches at the bottom.
When in the open position they are supported by two
support struts per Fan Cowl.
Thrust reverser C ducts
The Thrust Reverser C-ducts are hinged from the aircraft
strut at the top by four hinged type brackets and are
secured by six latches at the bottom.
When in the open position they are supported by two
support struts per C-duct.
Propulsion System Materials and Weights
Intake cowl
The intake cowl is made up of the following materials:
• Intake D section is aluminium.
• Intake cowl is carbon fibre.
• Intake cowl weight is 238 lbs. (107.98 Kg).
Fan cowl doors
The fan cowl doors are made up of the following materials:
• Carbon fibre and aluminium.
• LH fan cowl door weight is 79 lbs. (35.84 Kg).
• RH fan cowl door weight is 86 lbs. (39.01 Kg).
Thrust Reverser C-ducts
The thrust reverser C ducts are made up of the following
materials:
• C-duct structure and translating cowls are carbon fibre
and aluminium.
• The thrust reverser C-duct weight is 578 lbs. (262.25
Kg).
Common nozzle assembly (CNA)
The CNA is made up of the following material:
• Titanium.
• CAN weight is 213 lbs.
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Intake Cowl
Purpose
To supply all the air required by the engine, with minimum
pressure losses and with an even pressure face to the fan.
Nacelle drag is also minimised due to the aerodynamically
streamlined design.
Location
The inlet cowl is bolted to the front of the LPC case (Fan).
Description
The intake cowl is constructed from hollow inner and outer
skins. These are supported by front (titanium) and rear
(Graphite/Epoxy composite) bulkheads.
Inner and outer skins are manufactured from composites.
The leading edge is a 'one piece' pressing in Aluminium.
The cowl weight is approximately 238 lbs.
The intake cowl has the following features:
• Integral thermal anti-icing system.
• P2T2 Probe.
• Ventilation Intake.
• Interphone socket.
• Engine attachment ring with alignment pins to ensure
correct location of the cowl on to the fan case.
• Door locators that automatically align the fan cowl doors
to ensure good sealing.
• Strut brackets to provide location for the left and right
hand fan cowl door support struts (front struts only).
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Fan Cowl Doors (FCD)
Purpose
The fan cowl doors provide for an aerosmooth exterior while
enclosing the fan case mounted accessories.
Location
They are located about the fan casing.
Hinged at the top to the aircraft strut and secured by four
latches at the bottom.
Description
The doors extend rearwards from the inlet cowl to overlap
leading edge of the 'C' ducts.
The A320 aircraft have a strake on the inboard cowl of each
engine, the right hand cowl on both engine 1 and left-hand
cowl on engine 2.
The A319 aircraft have strakes on both the left-hand and right
hand cowls on both engines 1 and 2.
Fan cowls are interchangeable between the A319 and A320
except for the strake configuration. Make sure the correct
configuration is installed.
The fan cowl doors are constructed from graphite skins
enclosing an aluminium honeycomb inner.
Aluminium is also used to reinforcement each corner to
minimises handling/impact damage and wear.
The fan cowl doors abut along the bottom centre line and are
secured to each other by 4 quick release and adjustable
latches.

Thrust Reverser C Ducts
Purpose
The thrust reverser C ducts provide for:
• An aerosmooth exterior to minimise drag.
• The fan bypass ducting.
• Reverse thrust for aircraft deceleration.
Location
The thrust reverser C ducts are hinged from the aircraft
strut at the top and are secured at the bottom by six toggle
type clamps.
Description
The thrust reverser C ducts extend rearwards from the fan
cowls to the common nozzle assembly (CNA).
The thrust reverser C ducts;
Form the cowling around the core engine (inner barrel) to
assist in stiffening the core engine (load-share).
Form the fan air duct between the fan case exit and the
entrance to the CNA.
House the thrust reverser operating mechanism and
cascades.
Form the outer cowling between the fan cowl doors and
CNA.
The thrust reverser C ducts are mostly constructed from
composites but some sections are metallic mainly
aluminium for example the inner barrel, blocker doors and
links.
The thrust reverser C-ducts can be opened for access to
the core engine. This allows maintenance to be carried out
on the core engine while the engine is installed to the
aircraft.
The thrust reverser C-ducts are heavy therefor hydraulic
actuation is required to open them. Normal engine oil is
used in a hand-operated pump.
The thrust reverser C-ducts are held in the open position
by two support struts.
• The forward strut is a fixed length.
• The rear strut is a telescopic support.
•

Combined Nozzle Assembly (CNA)
Purpose
The CNA allows the mixing of the hot and cold stream gas
flows to produce the resultant thrust.
Location
The CNA is bolted to the rear flange of the turbine exhaust
casing. There is no fixing to the bottom of the pylon.
Description
The CNA:
Forms the exhaust unit.
• Mixes the hot and cold gas streams and ejects the
combined flow to atmosphere through a single
propelling nozzle.
• Completes the engine nacelle.

Engine Mounts
Purpose
The engine mounts suspend the engine from the aircraft
strut.
The engine mounts transmit loads generated by the
engine during aircraft operation.
Location
The front engine mount is located at the rear of the
intermediate case at the core engine.
The rear engine mount is located on the LPT casing at
TDC.
Description
Forward engine mount
The forward engine mount is designed to transmit the
following loads:
• Thrust loads.
• Side loads.
• Vertical loads.
The front mount is secured to the intermediate case in
three positions:
A monoball type universal joint. This gives the main
support at the front engine mount position.
Two thrust links that are attached to:
• The cross beam of the engine mount.
• Support brackets either side of the monoball location.
Rear engine mount
The rear engine mount is designed to transmit the
following loads:
• Torsional loads.
• Side loads.
• Vertical loads.
The rear engine mount has a diagonal main link that gives
resistance to torsional movement of the casing as a result
of the hot gas passing through the turbines.
There is further support from two side links. These limit the
engine side to side movement and give vertical support.

FIRE PROTECTION AND VENTILATION

Fire Protection and Ventilation
Purpose
The purpose of fire protection is to give an indication to the
flight deck of a possible fire condition about the engine.
The purpose of the ventilation system is to provide a flow
of cooling air about the engine to reduce the risk of a fire
condition annunciation to the flight deck.
Location
The locations of the fire detection fire wires are about the
fan casing and core engine.
The location of the ventilation air is about the entire of the
fan case and core engine.
Description
The engine is ventilated to provide a cooling airflow for
maintaining the engine components within an acceptable
operating temperature.
Also to provide a flow of air that assists in the removal of
potential combustible liquids that may be in the area.
Ventilation is provided for:
• The fan case area (Zone 1).
• The core engine area (Zone 2).
Zones 1 and 2 are ventilated to:
• Prevent accessory and component over heating.
• Prevent the accumulation of flammable vapours.
Zone 1 ventilation
Ram air enters the zone through an inlet located on the
upper LH side of the air intake cowl.
The air circulates through the fan compartment and exits
at the exhaust located on the bottom rear centre line of the
fan cowl doors.
Zone 2 ventilation
Metered holes within the inner barrel of the “C” duct allow
pressurized fan air to enter the zone 2 area.
Air exhausting from the active clearance control (ACC)
system around the turbine area also provides ventilation
air for Zone 2.
The air circulates through the core compartment and exits
through the lower bifurcation of the C ducts via the thrust
recovery duct.
Ventilation during ground running
During ground running local pockets of natural convection
exist providing some ventilation of the fancase zone 1.
Zone 2 ventilation is provided by the fan duct pressure as
above during ground running and in flight.

Fire Detection System
Purpose
The fire detection system monitors the air temperature in
Zone 1 and Zone 2.
When the air temperature increases to a pre determined
level the system provides flight deck warning.
Location
The fire detection system is located:
• Routed around the high-speed external gearbox.
• At BDC of the core engine nearest to the combustor
diffuser case.
Description
The V2500 utilises a Systron Donner fire detection system.
It has a gas filled core and relies upon heat exposure to
increase the internal gas pressure. Thus triggering
sensors.
When the air temperature about the fan case and/or core
engine increases to a pre-determined level the system is
designed to detect this and display a warning message
and indications to the flight deck.
The system provides flight deck warning by:
• Master warning light.
• Audible warning tone.
• Specific ECAM fire indications.
• Engine fire push button illuminates.
Zone 1 and Zone 2 fire detectors function independently of
each other.
Each zone has two detector units which are mounted as a
pair, each unit gives an output signal when a fire or
overheat condition occurs.
The two detector units are attached to support tubes by
clips.
Nacelle air temperature (NAC)
Zone 2 has the nacelle air temperature sensor.
Indication is to the flight deck when a temperature
exceedance has occurred.

Fire Detection System and Detector Units
The fire detection system employs detector units called
firewires.
The firewires are mounted in pairs. This is necessary due
to the class1 level 3 message that they generate when a
fire or overheat condition exists.
The fire detection system comprises of the following units:
• The firewires send a signal to the Fire Detection Unit
(FDU).
• The FDU sends a signal to the Flight Warning
Computer (FWC).
• The FWC generates the flight deck indications for a fire
condition.
There is one FDU per engine. The FDU has two channels;
each channel is looking at a separate fire detector loop of
zones 1 and 2.
Under normal conditions both firewires require to be
indicating to the FDU to give a real indication to the flight
deck.
If there is a single loop failure of more than 16 seconds
then the remaining firewire will continue to operate. The
FDU will recognise the faulty fire loop.
The faulty loop will be indicated to ECAM as the following
message:
ENG 1 (2) FIRE LOOP A (B) FAULT
If there is a double loop failure then the FDU will recognise
this as a possible burn through and the fire message will
be generated to the flight deck.
Firewire detectors
Each of the fire wire detector units comprises of the
following:
• A hollow sensor tube.
• A responder assembly.
Sensor tube
The sensor tube is closed and sealed at one end and the
other open end is connected to the responder.
The tube is filled with helium gas and carries a central core
of ceramic material impregnated with hydrogen.
An increase in the air temperature around the sensor tube
causes the helium to expand and increase until the
pressure causes the alarm switch to close. The FDU
recognises this as an abnormal situation, hence fire
indication will be illuminated.
If a ‘burn through’ occurs, the pressure within the sensing
tube is lost and as a result of this the integrity switch
opens to give an indication to the FDU of a loop failure.
Responder
The responder has two pressure switches, one normally
open and the other normally closed.
• The normally open switch is the alarm indication.
• The normally closed switch is the fault indication.

Fire Detection System Fire Bottles
Purpose
The fire bottles provide a means of extinguishing a
potentially hazardous fire about the engine when a fire
annunciation to the flight deck has occurred.
Location
The engine fire bottles are located in the aircraft strut.
Access for maintenance is via a panel that can be found
on the left hand side.
Description
The fire bottles have the following features:
• Agent type is bromotrifluoromethane.
• Charged to a nominal pressure of 600 psi at 21 deg. C.
• Pressure switch.
• Discharge head.
• Discharge squibs.
The pressure switch is set to indicate bottle empty when
the pressure falls below 225 psi. The indication in the flight
deck is:
AGENT 1 (2) SQUIB DISC
This is an illuminating annunciator light on the overhead
panel.
The discharge head has a leak proof diaphragm that is
designed to rupture when:
• The squib is activated from the flight deck.
• Excessive pressure in the fire bottle. 1600 to 1800 psi
at 95 deg. C
The squib is an Electro Pyrotechnic Cartridge containing
explosive powder. Two filaments ignite the powder when
they are supplied with 28v dc.
There is facility to carry out a fire system test that will give
all the expected indications if all is functioning correctly.
The fire test switch is located on the fire push button panel
on the overhead panel.

Fire Detection System Indications and Controls
Purpose
The purpose of the fire detection system indications is to
alert the flight crew to a possible fire condition.
The controls allow the flight crew to react and deal with the
impending fire indication in the flight deck.
Location
The fire control panel is located on the overhead panel for
fire bottle operation and fire system test.

Nacelle Air Temperature (NAC)
Purpose
The nacelle air temperature gives an advisory indication to
the lower ECAM CRT if a temperature exceedance has
been experienced.
Location
The NAC sensor is located by the bifurcation panel at
bottom dead centre between the two-thrust reverser C
duct halves.
The NAC is in zone 2.
Description
Under normal conditions the NAC indication is not
displayed on the lower ECAM CRT.
When a temperature exceedance of 320 deg.c has
occurred the indication will appear to the lower ECAM
CRT.
This indication is displayed if;
The system is not in engine starting mode and one of the
two temperatures reaches the advisory threshold.

SECTION 3
MECHANICAL ARRANGEMENT

IAE V2500 General Familiarisation Mechanical Arrangement
Mechanical Arrangement General
The engine is an axial flow, high by-pass ratio, twin spool
turbo fan.
The general arrangement is shown below.
L.P. System
Four stage L.P. compressor - comprising:
• 1 Fan stage
• L.P. Compressor consisting of 4 stages driven by:
• Five stage L.P. Turbine
H.P. System
• Ten-stage axial flow compressor driven by a 2 stage
H.P. Turbine.
• Variable angle inlet guide vanes.
• Variable stator vanes (3 stages A5).
• Handling bleed valves at stage 7 and 10.
Customer service bleeds at stage 7 and 10
Combustion System
• Annular, two piece, with 20 fuel spray nozzles.
Gearbox
• Radial drive via a tower shaft from H.P. Compressor
shaft to fan case mounted Angle and Main gearboxes.
Gearbox provides mountings and drive for all engine
driven accessories and the pneumatic starter motor.
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Engine Main Bearings
The main bearing arrangement and the bearing numbering
system is shown below.
The 5 bearings are located in 3 bearing compartments:
• The Front Bearing Compartment, located at the centre
of the Intermediate Case, houses No's 1,2 & 3 bearings.
• The Centre Bearing Compartment located in the
diffuser/combustor case houses No 4 Bearing.
• The Rear Bearing Compartment located in the Turbine
Exhaust Case houses No 5 Bearing.
No 1 Bearing
• Shaft axial location bearing.
• Takes the thrust loads of the L.P. shaft.
• Single track ball bearing.
No 2 Bearing
• Radial support for the front of the L.P.turbine shaft.
Single track roller bearing utilising "squeeze film" oil
damping.
• No 3 Bearing
• H.P. shaft axial location bearing.
• Radial support for the front of the H.P.shaft.
• Takes the thrust loads of the H.P. shaft.
• Single track ball bearing.
• Mounted in a hydraulic damper, which is centred by a
series of rod springs (squirrel cage).
No 4 Bearing
• Radial support for turbine end of H.P. shaft.
• Single track roller bearing.
No 5 Bearing
• Radial support for the turbine end of the L.P. shaft.
• Single track roller bearing.
Squeeze film oil damping.
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Engine Internal Cooling and Sealing Airflows
Purpose
To provide sealing air for the bearing chambers so as to
prevent oil loss.
To provide cooling air for the engines internal components
keeping them within designed operating temperatures.
Location
The air used for internal cooling and sealing is taken from the
compressor stages of:
• LPC stage 2.5
• HPC stage 8.
• HPC stage 10.
• HPC stage 12.
• The fan bypass provides external cooling air.
Description
Fan air is used to provide:
• Air for the Active Clearance Control (ACC) system.
This is used to control the tip clearances of the turbine
blades.
• Air through the Air Cooled Air Cooler (ACAC). This is
used for the precooling of the ‘buffer air’.
Buffer air is used to provide:
• Cooling, sealing and scavenge air for the No.4 Bearing
Chamber.
LPC stage 2.5 air is used for
• Sealing of the front and rear of the Front Bearing
Chamber
HPC stage 7 air is used for airflow control for compressor
stability and aircraft services bleed supply.
HPC stage 8 is used for:
• Sealing the hydraulic seal of the Front Bearing
Chamber and the sealing of the No. 5 Bearing
Chamber.
HPC stage 10 air is used for:
• Airflow control and aircraft services supply.
• ‘Make up’ air supply for the HPT stage 2 disc and
blades.
• Cooling air for the HPT stage 2 NGVs.
HPC stage 12 air is used for:
• Combustion chamber cooling.
• HPT stage 1 blades and NGVs cooling.
• The supply to the ACAC for buffer air cooling and
sealing of the no. 4 bearing chamber.
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Modular Construction
Modular construction has the following advantages:
• lower overall maintenance costs
• maximum life achieved from each module
• reduced turn-around time for engine repair
• reduced spare engine holdings
• ease of transportation and storage
• rapid module change with minimum ground running
• easy hot section inspection
• vertical/horizontal build strip
• split engine transportation
• compressors/turbines independently balanced
Module Designation
Module No Module
31 Fan
32 Intermediate
40 HP System
− 41 - HP Compressor
− 45 - HP Turbine
50 LP Turbine
60 External gearbox
Note:
The module numbers refer to the ATA chapter reference
for that module.
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Module 31
Description
Module 31 (Fan Module) is the complete Fan assembly
and comprises:
• 22 Hollow fan blades
• 22 Annulus Fillers
• Fan Disc
• Front and Rear Blade Retaining Rings
The blades are retained in the disc radially by the dovetail
root.
The front and rear blade retaining rings provides axial
retention. Blade removal/replacement is easily achieved by
removing the front blade retaining ring and sliding the
blade along the dovetail slot in the disc.
22 annulus fillers form the fan inner annulus.
The nose cone and fairing smooth the airflow into the fan.
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Module 32 - Intermediate Case
The Intermediate Module comprises of:
• Fan Case
• Fan Duct
• Fan Outlet Guide Vanes (OGV)
• LP Compressor ( A5 - 4 stage)
• LP Compressor Bleed Valve (LPCBV)
• Front engine mount structure
• Front bearing compartment which houses Nos. 1, 2
and 3 bearings
• Drive gear for the power off-take shaft (gearbox drive)
• LP stub shaft
• Inner support struts
• Outer support struts
• Vee groove locations for the inner and outer barrels of
the 'C' ducts

Module 32 - Intermediate Case
Instrumentation
The following pressures and temperatures are sensed and
transmitted to the E.E.C.
• P12.5
• P2.5
• T2.5
The rear view of the intermediate case is shown below.

Module 40 HP Compressor
Description
The HP compressor assembly (Module 40 is a 10 stage
axial flow compressor. It has a rotor assembly and stator
case. The compressor stages are numbered from the
front, with the first stage is stage being designated as
stage 3 of the whole engines compressor system. Airflow
through the compressor is controlled by variable inlet
guide vanes (VIGV), variable stator vanes (VSV) and
bleed valves.
The rotor assembly has five sub-assemblies
(1) Stages 3 to 8 HP compressor disks
(2) A vortex reducer ring.
(3) Stages 9 to 12 HP compressor disks
(4) The HP compressor shaft.
(5) The HP compressor rotating air seal.
The five sub-assemblies are bolted together to make the
rotor. The compressor blades in stages 3 to 5 are attached
to the compressor disks in axial dovetail slots and secured
by lockplates. The stages 6 to 12 compressor blades are
installed in slots around the circumference of the disks
through an axial loading slot. Lock blades, lock nuts and
jack screws hold the blades in position.
The HP compressor stator case has two primary sub-
assemblies, the HP compressor front and rear cases.

Module 40 HP Compressor
The HP compressor front case assembly has two split
cases bolted together along the engine horizontal centre
line.
The front case assembly contains the VIGV’s, the stages 3
to 5 VSV’s and the stage 6 stator vanes.
The front outer case provides a mounting for the VIGV and
VSV actuator. The front case assembly is bolted to the
intermediate case and to the rear outer case.
The HP compressor rear case assembly has five inner ring
cases and an outer case. Flanges on the inner cases form
annular manifolds, which provide stages 7 and 10 air
offtakes.
The five inner cases are bolted together, with the front
support cone bolted at the stage 7 case and the stage 11
case bolted to the rear outer case. The five inner cases
contain the stages 7 to 11 fixed stator vanes.
The rear outer case is bolted to the diffuser case and to
the rear flange of the HP compressor front case.
Access is provided in the compressor cases for borescope
inspection of the compressor blades and stator vanes

Combustion Section
The combustion section includes the diffuser section, the
combustion inner and outer liners, and the No 4 bearing
assembly.
Diffuser Casing
The diffuser section is the primary structural part of the
combustion section.
The diffuser section has 20 mounting pads for the
installation of the fuel spray nozzles. It also has two
mounting pads for the two ignitor plugs.
Combustion Liner
The inner and outer liners form the combustion liner.
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 (50 per inner and outer liner). The segments
can be replaced independently during engine overhaul.

HP Turbine
Description
The primary parts of the HP turbine rotor and stator
assembly are:
The HP Turbine Rotor Assemblies (Stage 1 and Stage 2)
The HP Turbine Case and Vane Assembly
The HP turbine rotor assemblies are two stages of turbine
hubs with single-crystal, nickel-alloy blades. The two-hub
configuration removes a bolt flange between hubs. This
decreases the weight and enables faster engine assembly.
The blades have airfoils with high strength and resistance
to creep. Satisfactory blade tip clearances are supplied by
active clearance control (ACC) to cool the case with
compressor air.

LP Turbine
Description
The primary parts of the Low Pressure Turbine (LPT)
module are:
• LPT Five Stage Rotor
• LPT Five Stage Stator Vanes
• Air Seals
• LPT Case
• Inner and Outer Duct
• LPT Shaft
• Turbine Exhaust Case (TEC)
The LP turbine has a five stage rotor which supplies power
to the LP compressor through the LPT shaft. The LPT
rotor is installed in the LPT case where it is in alignment
with the LPT stators. The LPT case is made from high-
heat resistant nickel alloy and is a one part welded
assembly. To identify the LP turbine module, an
identification plate is attached to the LP turbine case at the
136degrees position.
The LPT case has two borescope inspection ports at
125.27 and 237.10 degrees. The ports are used to
internally examine the adjacent engine sections:
• Trailing Edge (TE), Stage 2, HPT Blades
• Leading Edge (LE), Stage 3, LPT Blades
• Trailing Edge (TE), Stage 3, LPT Blades
The five LPT disks are made from high heat resistant
nickel alloy. The LPT blades are also made from nickel
alloy and are attached to the disks by fir-tree roots. The
blades are held in axial position on the disk by the rotating
air seals (knife-edge).

Module 60 - External Gearbox
Purpose
The gearbox assembly transmits power from the engine to
provide drives for the accessories mounted on the gearbox
front and rear faces.
During engine starting the gearbox also transmits power
from the pneumatic starter motor to the core engine.
The gearbox also provides a means of hand cranking the
HP rotor for maintenance operations.
Location
The gearbox is mounted by 4 flexible links to the bottom of
the fan case.
• Main gearbox 3 links.
• Angle gearbox 1 link.
Description
The external gearbox is a cast aluminium housing that has
the following features;
• Individually replaceable drive units.
• Magnetic chip detectors.
• Main gearbox 2 magnetic chip detectors.
• Angle gearbox 1 magnetic chip detector.
The following accessory units are located on the external
gearbox;
Front Face Mount Pads
• De-oiler.
• Pneumatic starter.
• Dedicated generator.
• Hydraulic Pump.
• Oil Pressure pump and filter.
Rear Face Mount Pads
• Fuel pumps (and fuel metering unit FMU).
• Oil scavenge pumps unit.
• Integrated drive generator (IDG).
The Oil sealing for the gearbox to accessory drive links is
provided by a combination of carbon and ‘O’-ring type
seals.
The carbon seals can be replaced while the engine is on
wing.

Engine View Right Hand Side
The following components are located on the right hand
side of the engine.
1. Stage 10 make-up air valve for supplementary turbine
cooling.
2. IDG harness interface.
3. Harness interface.
4. Start air and anti ice ducting interface.
5. Electrical harness interface.
6. Air starter duct.
7. Engine electronic control.
8. Anti ice duct.
9. Relay box.
10.Anti ice valve.
11.Starter valve.
12.10th
stage handling bleed valve solenoid.
13.No.4 bearing scavenge valve.
14.Air-cooled oil cooler (ACOC).
15.Intergrated drive generator (IDG).
16.Exciter ignition boxes.
17.Fuel distribution valve.
18.HPC stage 7B handling bleed valve.
19.LPT and HPT active clearance control valves (ACC).
20.HPC stage 10 handling bleed valve.
21.Engine rear mount.
22.Booster bleed valve slave actuator.
23.Front engine mount.
24.HPC 10th
stage cooling air for the HPT 2nd
stage NGVs.
25.Solenoids for the three off HPC 7th
stage handling
bleed valves.
26.Solenoid for the HP10 make-up cooling air control
valve.
27.Solenoid for the HP10 cabin bleed pressure
regulating/shut-off valve (PRSOV).

Engine View Left Hand Side
The following components are located on the left-hand
side of the engine.
1. Fan cowl door hinged brackets (4 off).
2. Thrust reverser hydraulic control valve (HCU).
3. Hydraulic tubes interface.
4. Fuel supply and return to wing tank.
5. C duct front hinge.
6. Thrust reverser hydraulic tubes interface.
7. Over pressuerization valve (OPV).
8. 2.5 bleed master actuator.
9. C Duct floating hinges.
10.Fan Air Valve (FAV).
11.C Duct rear hinge.
12.Opening actuator mounting brackets.
13.C Duct compression struts (3off).
14.Cabin bleed air pre cooler duct interface.
15.Cabin bleed air system interface.
16.Pressure regulating valve (PRV).
17.Air-cooled air cooler (ACAC).
18.HPC 10th
stage cabin bleed offtake pipe.
19.HPC 10th
stage pressure regulating/shut-off valve
(PRSOV).
20.HPC 7th
stage bleed valve (HPC7 C).
21.HPC 7th
stage cabin bleed non-return valve (NRV).
22.VIGV/VSV actuator.
23.Fuel pumps and fuel metering unit.
24.High speed external gearbox.
25.Hydraulic pump.
26.Engine oil tank.
27.IDG oil cooler.
28.LP fuel filter.
29.Fuel cooled oil cooler (FCOC).
30.Savenge oil filter pressure differential switch.
31.Fuel return to tank valve (part of item 32).
32.Fuel diverter valve (part of item 31).
33.Oil pressure differential transmitter.
34.Low oil pressure switch.

Engine Combined Drains System
Purpose
To provide an early indication of a system or component
failure by evidence of a fluid leak.
Location
The drains systems of tubes are located about the engine.
The drains mast is located at BDC of the fan case. It
protrudes from the bottom of the fan cowl doors.
Description
This provides a combined overboard drain through a
drains mast. The drains are for fuel and oil from the core
module components, the LP compressor/intermediate
case components and the external gearbox.

SECTION 4
ELECTRONIC ENGINE CONTROL

IAE V2500 General Familiarisation Electronic Engine Control
Electronic Engine Control Introduction
The V2500 uses a Full Authority Digital Electronic Engine
Control (FADEC).
The FADEC comprises the sensors and data input, the
electronic engine control unit (EEC) and the output
devices, which include solenoids, fuel servo operated
actuators and pneumatic servo operated devices. The
FADEC also includes electrical harnesses.
Engine Electronic Control
The heart of the FADEC is the Engine Electronic Control
(EEC) unit - shown below. The EEC is a fan case mounted
unit, which is shielded and grounded as protection against
EMI - mainly lightning strikes.
Features
• Vibration isolation mountings.
• Shielded and grounded (lightning strike protection).
• Size - 15.9 X 20.1 X 4.4 inches.
• Weight - 41 lbs.
• Two independent electronic channels.
• Two independent power supplies, the EEC utilises
67.53 Watts of power from either the three phase AC
from a dedicated engine mounted alternator, or 28
Volts DC from an aircraft source.
• A two way Pressure Relief Valve maintains the units
differential pressure (< 5 PSID).
• Six ‘screened’ pressure ports provide the required
pressure inputs to both channels.
• Built in handle facilitates removal and handling.
• Has three control modes in each channel. Engine
Pressure Ratio (EPR) – which is the Primary thrust
control Mode. N1 Rated and Un-rated and also
provides Auto Starting and Thrust Reverser control. (To
be covered in detail later).
• Schedules engine operation to provide maximum
engine performance and fuel savings.
• Provides improved engine starting (Auto Start) and
transient characteristics (acceleration/deceleration).
• Provides maximum engine protection and is more
flexible to readily adapt to changes in engine
requirements.
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The Engine Electronic Control (EEC) Description
The EEC is a dual channel control unit that utilises a split
housing design.
The assembled unit is sealed with a housing seal and a
protective shield provides channel separation.
The control assembly is separated into two modules, each
containing one control channel.
Each module contains two multi-layer printed circuit
boards assemblies, which enable it to function
independently of the other channel.
A mating connector provides ‘Crosstalk’, for partial or
complete channel switching and fault isolation logic when
the two modules are joined.
This connector also provides for the exchange of ‘cross-
link data’, cross wiring and hardwired discretes between
the two channels.
The EEC has two identical electronic circuits that are
identified as Channel A and Channel B. Each channel is
supplied with identical data from the aircraft and the
engine.
This data includes throttle position, aircraft digital data, air
pressures, air temperatures, exhaust gas temperatures
and rotor speeds.
The EEC, to set the correct engine rating for the flight
conditions uses this data. The EEC also transmits engine
performance data to the aircraft.
This data is used in cockpit display, thrust management
and condition monitoring systems.
Each of the EEC channels can exercise full control of all
engine functions. Control alternates between Channel A
and Channel B for consecutive flights, the selection of the
controlling channel being made automatically by the EEC
itself.
The channel not in control is nominated as the back up
channel
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.
Electronic Engine Control
Harness (electrical) and Pressure Connections
Two identical, but separate electrical harnesses provide
the input/output circuits between the EEC and the relevant
sensor/control actuator, and the aircraft interface.
The harness connectors are 'keyed' to prevent
misconnection.
Note:Single pressure signals are directed to pressure
transducers - located within the EEC - the pressure
transducers then supply digital electronic signals to
channels A and B.
The following pressures are sensed: -
• Pamb ambient air pressure - fan case sensor
• Pb burner pressure (air pressure) P3/T3 probe
• P2 fan inlet pressure - P2/T2 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
Electrical Connections
Front Face
J1 E.B.U. 4000 KSA
J2 Engine D202P
J3 Engine D203P
J4 Engine D204P
J11 Engine D211P
Rear Face
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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Engine Electronic Control (EEC.)
Overview
The EEC provides the following engine control functions:-
• Power Setting (E.P.R.).
• Acceleration and deceleration times.
• Idle speed governing.
• Overspeed limits (N1 and N2).
• Fuel flow.
• Variable stator vane system (V.S.V.)
• Compressor handling bleed valves.
• Booster stage bleed valve (B.S.B.V.).
• Turbine cooling (10 stage make-up air system).
• Active clearance control (A.C.C.).
• Thrust reverser.
• Automatic engine starting.
• Oil and fuel temperature management.
Note:
The fuel cut off (engine shut down) command comes from
the flight crew and is not controlled by the EEC.
Fault Monitoring
The EEC has extensive self test and fault isolation logic
built in. This logic operates continuously to detect and
isolate defects in the EEC.
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Electronic Engine Control (EEC) Data Entry Plug
Purpose
The Data Entry Plug (DEP) provides discrete data inputs
to the EEC. Located on to 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 (Selected from multiple rating options).
• Engine Serial No.
Location
The data entry plug is located on the channel B side
electrical connectors of the EEC.
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DATA ENTRY PLUG (DEP)

Electronic Engine Control
Failures and Redundancy
Improved reliability is achieved by utilising dual sensors,
dual control channels, dual selectors and dual feedback.
• Dual sensors are used to supply all EEC inputs except
pressures, (single pressure transducers within the EEC
provide signals to each channel - A and B).
• The EEC uses identical software in each of the two
channels. Each channel has its own power supply,
processor, programme memory and input/output
functions. The mode of operation and the selection of
the channel in control is decided by the availability of
input signal and output controls.
• Each channel normally uses its own input signals but
each channel can also use input signals from the other
channel required i.e. if it recognises faulty, or suspect,
inputs.
• An output fault in one channel will cause switchover to
control from the other channel.
• In the event of faults in both channels a pre-determined
hierarchy decides which channel is more capable of
control and utilises that channel.
• In the event of loss of either channels, or loss of
electrical power, the systems are designed to go to the
fail safe positions.

Operation and Control
EEC Power Supplies
The electrical supplies for the EEC are normally provided
by a dedicated alternator, which is mounted to and driven
by the external gearbox.
Dedicated Alternator
The unit is a permanent magnet alternator which has two
independent sets of stator windings and supplies two
independent, 3 phase, frequency wild AC outputs to the
EEC These unregulated AC supplies are rectified to 28
volts DC within the EEC
The Dedicated Alternator also supplies the N2 (HP
Compressor speed) signal for the EEC. This is provided by
the frequency of a single phase winding in the stator
housing. This source is the ‘primary’ speed signal and is
used by both Channels of the EEC and for the Flight Deck
instrument display of engine actual speed. Should this
signal fail, there is a ‘Back-up’ signal which is derived from
one of the three phase windings of Channel ‘B’ power
generation.
There is no speed signal generation provided by the output
of the coil windings of the Dedicated Alternators Channel
‘A’ power supply.
The EEC also utilises aircraft power to operate some
engine systems:-
• 115 volts AC 400 Hz power is required for the ignition
system and inlet probe anti-icing heater
• 28V DC is required for some specific functions, which
include the thrust reverser, fuel on/off and ground and
test power for EEC maintenance.
In the event of a dedicated alternator total failure the EEC
is supplied from the aircraft 28V DC bus bars, 28V DC
from the same source is also used by the EEC during
engine starts until the dedicated alternator comes 'on line'
at approximately 10% N2.
The dedicated alternator comes on line and supplies the
EEC power requirement when the N2 reaches
approximately 10%.
Switching between the aircraft 28V supply and dedicated
alternator power supplies is done automatically by the
EEC.
The dedicated alternator is cooled by 12.5 cooling air.
piped from the fan exit pressure probe, which is mounted
in the upper fan case splitter fairing.
.

IAE International Aero Engines AG 2000
IAE V2500 General Familiarisation Power Management
Power Management
Purpose
The power management system is designed to allow the
control of engine power by either manual or auto throttle
control.
Location
The aircraft throttle is located in the flight deck. This is in
reference to the TLA resolvers.
The EEC is engine intermediate case mounted. This is in
reference to the TRA signal that is derived from TLA.
Description
The throttle control lever (Thrust Lever) is based on the
"fixed throttle" concept, there is no motorised movement of
the throttle levers.
Each throttle control lever drives dual throttle resolvers,
each resolver output is dedicated to one EEC channel.
The throttle lever angle (TLA) is the input to the resolver.
The resolver output, which is fed to the EEC, is known as
the Throttle Resolver Angle (TRA).
The relationship between the throttle lever angle and the
throttle resolver angle is linear therefor;
1 deg TLA = 1.9 deg TRA
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Throttle Control Lever Mechanism
The throttle control mechanism for one engine is shown
below.
The control system consists of:
• The throttle control lever.
• The mechanical box.
• The throttle control unit.
The throttle control lever movement is transmitted through
a rod to the mechanical box. The mechanical box
incorporates 'soft' detents which provides selected engine
ratings, it also provides "artificial feel" for the throttle
control system.
The output from the mechanical box is transmitted by a
second rod to the throttle control unit. The throttle control
unit incorporates two resolvers and six potentiometers.
Each resolver is dedicated to one EEC. channel, the
output from the potentiometers provides T.L.A. signals to
the aircraft flight management computers.
A rig pin position is provided on the throttle control unit for
rigging the resolvers and potentiometers.
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Bump Rating Push Button (A1 Engined Aircraft only)
In some cases (optional) 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.
Note:
Bump Ratings can be selected, regardless of TLA only in
EPR mode when aircraft is on ground.
Bump Ratings can be de-selected at any time by actuating
the bump rating push button, as long as the aircraft is on
the ground and the Thrust Lever is not in the Max Take-Off
detent.
In flight, the bump ratings are fully removed when the
Thrust Lever is moved from the Take-Off detent to or
below the Max Continuous detent.
The Bump Rating is available in flight (EPR or N1 mode)
under the following conditions;
• Bump Rating is initially selected on ground.
• Take-Off, Go Around TOGA Thrust position set.
• Aircraft is within the Take-Off envelope.
When Bump Rating is selected a ‘B’ appears next to the
associated EPR display. Use of Bump must be recorded.
When one Bump button is selected, both engines are
Bump Rated.
Pressing Bump again deselects Bump Rating.
Flexible Takeoff (A1 & A5Engined Aircraft)
Definition of Flexible Takeoff:
In many instances, the aircraft takes off with a weight
lower than the maximum permissible takeoff weight. When
this happens, it can meet the required performance with a
decreased thrust that is adapted to the weight: This is
called ‘Flexible Takeoff’ and the thrust is called ‘Flexible
Takeoff Thrust’. The use of Flexible Takeoff Thrust saves
engine life.
The maximum permissible takeoff weight decreases as
temperature increases, so it is possible to assume a
temperature at which the actual takeoff weight would be
the limiting one. This temperature is called ‘Flexible
Temperature’ or ‘Assumed Temperature’ and is entered
into the FADEC via the MCDU PERF TO page in order to
get the adapted thrust.
Note! If the thrust ‘Bump’ is armed for takeoff and flexible
thrust is used, the pilot must use the Takeoff Performance
determined for the non-increased takeoff thrust (without
Bump).
• Thrust must not be reduced by more than 25% of the
full rated thrust.
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Throttle Control Lever Mechanism
The throttle control lever moves over a range of 65
degrees, from minus 20 degrees to plus 45 degrees. An
intermediate retractable mechanical stop is provided at 0
degrees.
Forward Thrust Range
The forward thrust range is from 0 degrees to plus 45
degrees.
• 0 degrees = forward idle power.
• 45 degrees = rated take off power.
Two detents are provided in this range;
• Max climb (MCLB) at 25 degrees.
• Max continuous (MCT)/Flexible (de-rated) take off
power (FLTO) at 35 degrees.
Reverse Thrust Range
Lifting the reverse latching lever allows the throttle to
operate in the range 0 degrees to minus 20 degrees. A
detent at minus 6 degrees corresponds to thrust reverse
deploy commanded and reverse idle power, minus 20
degrees is max reverse power.
Auto Thrust System (ATS)
The Auto Thrust System can only be engaged between 0
degrees and plus 35 degrees.
Thrust Rating Limit
Thrust rating limit is computed according to the thrust lever
position. If the thrust lever is set in a detent the FADEC will
select the rating limit corresponding to this detent.
If the thrust lever is set between two detents the FADEC
will select the rating limit corresponding to the higher
mode.
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EEC/Fuel System Interface
Purpose
To allow the throttle signal from the flight deck to be
received by the EEC. The EEC will convert this signal into
a fuel flow error in order to change the fuel flow for a
power level change.
Description
Movement of the pilots throttle control lever is sensed by
the dual resolvers that signal the TRA to the EEC.
The EEC computes the fuel flow that will produce the
required thrust.
The computed fuel flow request is converted to an
electrical current (I) which drives the torque motor in the
Fuel Metering Unit (FMU) which modulates fuel servo
pressure to move the Fuel Metering Valve (FMV) and sets
the fuel flow.
Movement of the FMV is sensed by a dual resolver which
is located in the fuel-metering unit next to the FMV.
The dual resolver translates the fuel metering valve
movement into an electrical feedback signal that is fed
back to the EEC.
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IAE V2500 General Familiarisation Fuel System
Fuel System Introduction
Purpose
The primary purpose of the fuel system is to provide a
completely controlled continuous fuel supply in a form
suitable for combustion, to the combustion system.
Description
Control of the fuel supply is by the EEC via the Fuel
Metering Unit (FMU). High pressure fuel is also used to
provide servo pressure (actuator muscle) for the following
actuators;
• BSBV actuators.
• VSV actuator.
• ACC actuator.
• ACOC actuator.
The major components of the fuel system include;
• High and low pressure fuel pumps (dual unit).
• Fuel/oil heat exchanger.
• Fuel filter.
• Fuel metering unit (FMU).
• Fuel distribution valve.
• Fuel injectors (20).
• Fuel diverter and back to tank valve (FDRV).
The fuel system controls are on the centre control pedestal
and the indications are in the form of an annunciator light
and ECAM messages.
Operation
The aircraft pumps deliver the fuel to the engine LP pump.
The LP pump boosts the initial fuel delivery to a pressure
so as to prevent low pressure entry into the HP pump.
Nominal pressure 150psi.
The fuel flows into the fuel oil heat exchangers for the
engine and IDG.
Depending on the mode of operation the heat
management system is in depends on which direction the
fuel will flow.
From the engine FCOC the fuel passes through the LP
fuel filter. The filter has a 40 micron filtration capability.
The fuel is received by the HP pump and is boosted to a
nominal 1000 psi. The HP pump has pressure relief set at
1360 psi.
The FMU meters the fuel and the excessive HP fuel is
diverted back into the LP supply. The FMU is controlled by
signals from the EEC.
The fuel flow meter gives indication to the upper ECAM
screen of real time fuel flow in KG/H.
The distribution valve filters the fuel and splits the supply
into ten separate outlets.
The ten outlets supply fuel to two fuel spray nozzles per
outlet. The fuel spray nozzles have small filters within
them. This gives last chance filtration prior to fuel
atomisation.
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Fuel System Controls and Indications
Controls
The fuel metering valve is controlled by the selection of the
master lever located on the centre control pedestal.
The EEC has biased control of the FMU PRSOV for fuel
selection to on and fuel selection to off, if N2 is below 50%
and the start sequence is in auto.
The command for fuel selection to off when the indicated
N2 speed is above 50% is from the master lever.
Indications
The fuel temperature sensor is used by the EEC for the
function of the heat management system.
The fuel filter differential pressure switch annunciates to
the lower ECAM screen a message of FILTER CLOG. This
message is located in the right hand upper memo box.
The message of FILTER CLOG will occur when the fuel
filter differential pressure exceeds 5 psi.
If there is a disagreement between the selection of the
master lever and the PRSOV position then a fault exists.
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Fuel Pumps
Purpose
The fuel pumps are designed to ensure that the fuel
system recieves fuel at a determined pressure in order to
allow the atomisation of fuel in the combustion chamber.
Description
The combined fuel pump unit consists of low pressure and
high pressure stages that are driven from a common
gearbox, output shaft.
LP fuel pump
Purpose
To provide the necessary pressure increase to;
• Account for pressure losses through the Fuel Cooled
Oil Cooler and the LP fuel filter.
• Suppress cavitation.
• Maintain adequate pressure at the inlet to the HP
stage.
Description
Shrouded, radial flow, centrifugal impeller, with an axial
inducer.
HP Stage
Purpose
To increase the fuel pressure to that which will ensure
adequate fuel flow and good atomisation at all engine
operating conditions.
Description
Two gear (spur gear) pump.
• Provides mounting for fuel metering unit (FMU).
• Integral relief valve.
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Fuel Cooled Oil Cooler
Purpose
To transfer heat from the oil system to the fuel system to;
• Reduce the temperature of the engine lubricating oil
under normal conditions.
Prevent fuel icing.
Location
The fuel and oil heat exchanger is located on the left hand
side of the intermediate case. In the nine o’clock position.
Description
The fuel and oil heat exchanger is a single pass for the
flow of fuel and multi pass for the flow of oil.
The fuel and oil heat exchanger has the following features;
• A single casing houses the Fuel Cooled Oil Cooler and
the LP fuel filter.
• Provides location for the fuel diverter and back to tank
valve (unit not shown).
• Fuel temperature thermocouple.
• Fuel differential pressure switch.
• Oil system bypass valve.
• Fuel/oil tell tale leak indicator.
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Low pressure Fuel Filter
Purpose
To remove solid contaminants from the LP part of the fuel
system.
Location
The LP fuel filter is located in the LP fuel filter housing that
is integral with the fuel and oil heat exchanger.
Description
The LP fuel filter is a woven, glass fibre, disposable, 40
micron (nominal) type.
The LP fuel filter and housing have the following features;
• A differential pressure switch, which generates a flight
deck message, FUEL FILTER CLOG, if the differential
pressure across the filter, reaches 5 psid.
• A by-pass valve which opens and allows fuel to by-
pass the filter if the differential pressure reaches 15
psid.
• A fuel drain plug, used to drain filter case or to obtain
fuel samples.
• Fuel temperature sensor.
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Fuel Metering Unit (FMU)
Purpose
The FMU has three functions for fuel control. They are;
• Fuel metering to the combustion chamber.
• Control of the opening and closing off of the fuel supply
to the combustion chamber.
• Overspeed protection.
Location
The FMU is mounted on the combined fuel pumps
assembly.
The combined fuel pumps assembly is located on the rear
face of the high-speed gearbox, left hand side.
Description
The FMU is the interface between the EEC and the fuel
system.
All the fuel delivered by the HP fuel pumps, which is more
than the engine requires is passed to the FMU.
The FMU, under the control of the EEC, meters the fuel
supply to the fuel spray nozzles.
The HP fuel pressure also provides a servo operation
(muscle) for the following actuators;
• Booster stage bleed valve (BSBV) actuators.
• Variable stator vane (VSV) actuator.
• Active clearance control (ACC) actuator.
• Air cooled oil cooler (ACOC) actuator.
Excessive HP fuel supplies that are not required, other
than that for actuator control and metered fuel to the
combustor, is returned to the LP system via the spill valve.
In addition to the fuel metering function the FMU also
houses the overspeed valve and the pressure raising and
shut off valve.
The overspeed valve under the control of the EEC
provides overspeed protection for the LP (N1) and HP (N2)
rotors.
The pressure raising and shut off valve provides a means
of isolating the fuel supplies to start and stop the engine.
Note:
There are no mechanical inputs to, or outputs from, the
FMU.

IAE V2500 General Familiarisation Fuel SystemFuel System
Fuel Metering Unit (FMU)
Service Bulletin V2500-ENG-73-0172
This Service Bulletin introduces a Woodward Governor
Company FMU similar to the existing unit except for a
‘Common Flow/High Flow’ maximum fuel flow stop
assembly. This allows the unit to be switched to suit all
V2500-A5 model applications. This is considered
logistically advantageous for mixed fleet operators.
The changes introduced are:
a) The external single set fuel flow stop mechanism
has been deleted.
b) An external switchable two-position maximum
fuel flow stop has been introduced which can be
set for either A319/A320 or A321 aircraft
applications
c) A single reversible nameplate is introduced
which, in conjunction with stop setting letter and
FMU dataplate directive, will facilitate clear
unambiguous identification of each flow setting.
d) A security seal system is introduced onto the
above switchable fuel flow stop and reversible
nameplate.
e) To facilitate installation of the security seal lock
wire, the two existing retaining cap screws have
been replaced by lockwire compatible
equivalents.
FMU Part Number Position Setting Letter
FMU 8061-636 0
FMU 8061-637 X
(i)To switch 8061-636 to 8061-637, carryout
switch procedure in accordance with Woodward
Governor Company Service Bulletin 83724-73 Fuel
Metering Unit (FMU)
Service Bulletin V2500-ENG-73-0172 (Continued)
(ii) To switch 8061-637 to 8061-636, carryout
switch procedure in accordance with
Woodward Governor Company Service
Bulletin 83724-73-0004.
a) Re-connect engine harness and LP fuel tube
(Refer to AMM 73-22-52)
b) Close access to the engine (Refer to AMM
71-13-00)
c) Do an ‘idle’ check (Refer to AMM 71—00-00)
or a wet motor leak test (Refer to AMM 71-00-
00)
d) Do the operational tests of the starter and
FMU (Refer to AMM 80-13-51)
Do the operational FADEC test as per (AMM 73-22-00)

Fuel Distributor Valve
Purpose
The fuel distributor valve receives fuel from the FMU and
carries out three functions;
• Last chance filtration of the metered fuel.
• Distribution of the metered fuel through ten fuel supply
tubes to the fuel spray nozzles.
• Upon shut down allows fuel drain back (pressure
reduction) for prevention of fuel leaks into the combustor
upon engine shut down.
Location
The fuel distribution manifold is located on the right hand side
of the combustion diffuser casing. It is in the 4 o’clock
position.
Description
The fuel distributor manifold has the following features;
• Integral fuel filter - with by-pass valve.
• Single fuel metering (check) valve.
• Spring loaded closed upon engine shut down.
• Fuel pressure opened.
• Ten fuel outlet ports.

Fuel Spray Nozzles (FSN)
Purpose
To inject the fuel into the combustion chamber in a form
suitable for combustion by;
• Atomising the fuel.
• Mixing it with HPC delivery air.
• Controlling the spray pattern.
Location
The fuel spray nozzles are equi spaced around the
circumference of the combustor diffuser casing.
Description
Parker Hannifin manufactures the Airspray fuel nozzles.
The fuel spray nozzles have the following features;
• 20 fuel spray nozzles.
• Inlet fitting houses fuel filter.
• Internal and external heat shields to reduce coking.
• Transfer tubes for improved fuel leak prevention.

Fuel System Operation
Fuel Metering Unit Description
A simplified schematic representation of the Fuel Metering
Unit is shown below.
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 (2)
rotors.
• Isolation of fuel supplied for starting/stopping the
engine.
Three valves arranged as follows carry out these three
functions;
• The Fuel Metering Valve.
• The Overspeed Valve.
• The Pressure Raising and Shut Off Valve (PRSOV).
Fuel metering valve
The fuel metering valve varies the fuel flow according to
the EEC command.
The positional feedback to the EEC is by a rotary variable
displacement transducer (RVDT).
The overspeed valve
The overspeed valve protects the engine against an
exceedance of;
• N1 shaft speed. (109%)
• N2 shaft speed. (105.7%)
The feedback to the EEC of the valve operation is by a
micro switch.
The pressure raising and shut of valve (PRSOV)
The PRSOV is an open and close type valve. The PRSOV
controls the fuel to the combustor.
When the valve is in the pressure raising state it is said to
be open.
When the valve is in the shut off state it is said to be
closed.
Note:
The EEC has command to open the PRSOV upon an
engine start.
The EEC has command to close the PRSOV in auto start
mode and when the N2 is below 50%.
Above 50% N2 the close command is from the master
lever in the flight deck only.
Pressure drop governor and spill valve
The pressure drop governor controls the pressure
difference across the FMV.
The spill valve is controlled by the pressure drop governor.
The spill valve is designed to vary the excessive HP fuel
pressure return to the LP system.

IAE V2500 General Familiarisation Engine Oil System
Engine Oil System Introduction
Purpose
The oil system is a self contained, full flow recirculating
type design to ensure reliable lubrication and cooling
under all circumstances.
Description
Oil cooling is controlled by a dedicated Heat Management
System which ensure that engine oil, IDG oil and fuel
temperatures are maintained at acceptable levels while
ensuring the optimum cooling configuration for the best
engine performance.
The engine oil system can be divided into three sections.
These sections are;
• Pressure feed.
• Scavenge.
• Venting.
Pressure feed
The pressure feed system uses the full flow generated by
the pressure pump. The pressure pump moves the oil
through;
• The pressure filter.
• Fuel oil heat exchanger.
The oil is then distributed to the engine bearings and gear
drives.
Scavenge
The scavenge system is designed to retrieve the oil that is
present in the bearing chambers and gearbox for cooling
and recirculation.
There are six scavenge pumps that are designed to suck
the oil and pass it through;
• Magnetic chip detectors.
• A scavenge filter and master chip detector.
Prior to returning the oil back to the oil tank.
Venting
The venting system is designed to allow the air and oil mix
that develops in the bearing chambers and gearbox to
escape to the de oiler.
No.4 bearing does not have a scavenge pump. It relies
upon the build up of air pressure in the bearing chamber to
force the air and oil through the no.4 bearing scavenge
valve and into the de oiler.
Indications
There are flight deck indications that allow the oil system
to be monitored.
There are also messages generated ECAM for further
flight crew awareness.
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Engine Oil System Indications
The operation of the engine oil system may be monitored
by the following flight deck indications;
• Engine oil pressure.
• Engine oil temperature.
• Oil tank contents.
In addition ECAM alerts may be given for the following
non-normal conditions: -
• Low oil pressure.
• Scavenge filter clogged or partly clogged (high
differential pressure).
• No 4 compartment scavenge valve inoperative.
The oil system parameters are displayed on the Engine
page on the Lower ECAM screen.
Oil temperature (deg.c)
Normal indication to ECAM is GREEN.
156°C or above flashing green indication.
156°C or above more than 15 minutes or 165°C without
delay steady amber indication.
Upper ECAM message ENG 1(2) OIL HI TEMP-Level 2.
Oil low temperature alert, throttle above idle and engine
running.
Upper ECAM message ENG 1(2) OIL LO TEMP-level 2.
Single chime.
Master caution light.
Oil quantity
Less than 5 quarts flashes green.
Oil pressure
390 psid or above indication flashes.
60-80 psid amber indication.
Upper ECAM amber message ENG OIL LO PR level 1.
60 psid or below red indication.
Master warning light.
Continuous repetitive chime.
Upper ECAM red message level 3;
ENG 1(2) OIL LO PR
THROTTLE 1(2) IDLE
Scavenge filter clog
If the filter differential pressure is greater than 12 psi oil
filter clog message appears on Engine page, lower ECAM.
Oil Consumption
Acceptable oil use is not more than 0.6 US pts/hr (0.5 Imp
pts/hr).
Oil increase of 100 cc’s or more analyse sample for fuel
contamination.
master caution light
single chime
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Oil System Bearings and Gears Lubrication
Front Bearing Compartment (Bearings no. 1, 2, 3)
Purpose
Bearings and gears require oil for;
• Lubrication.
• Cooling.
• Vibration suppression.
Location
The following bearings and gears are located in the front
bearing compartment;
• Ball bearing no.1.
• Roller bearing no.2.
• Ball bearing no.3.
Description
The bearing chamber utilises hydraulic seals and carbon
seals to contain the oil within the bearing chamber.
The front seal has LPC booster stage 2.5 air passing
across the seal in order to prevent oil loss.
The rear seal has LPC 2.5 air passing across the seal in
order to prevent oil loss.
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.
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Centre Bearing Compartment (Bearing no.4)
Purpose
Bearings require oil for;
• Lubrication.
• Cooling.
Location
The following bearing is located in the centre bearing
compartment;
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;
• 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.
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Rear Bearing Compartment (Bearing no.5)
Purpose
Bearings require oil for;
• Lubrication.
• Cooling.
Location
The following bearing is located in the rear bearing
compartment;
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;
• 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.
The presence of vent air in the scavenge oil is used to
pressurise the oil tank.
Excess air pressures that develop in the oil tank vent to
the de oiler.
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High speed external gearbox
Purpose
Gears require oil for;
• Lubrication.
• Cooling.
Location
The following module is located at the six o’clock position
on the intermediate module.
Description
The high speed external gearbox is a one piece casting
consisting of the following;
• Gear trains.
• Oil jets.
• Two scavenge outlets with strainers.
• Vent out to the de oiler.
• Integrally mounted oil tank.
• Angle gearbox.
• Accessory units.
The gear ratios differ to suit the rotational operating
speeds of the accessory units.
The high speed external gearbox gears are lubricated by;
• Oil jets directing the oil onto the gears.
• Splash lubrication caused by the motion of the gears.
The high speed external gearbox has;
• Scavenge oil recovery by two scavenge pumps.
• Vent air outlet to allow the vent air to escape to the de
oiler.

Oil Tank
Purpose
To store the dedicated engine oil supply.
Location
Located to the top LH side of the external gearbox.
Description
The engine oil tank has the following features;
Pressurised hot tank.
Oil quantity transmitter.
• Gravity fill port with safety flap.
• Sight glass oil level indicator.
• Internal 'cyclone' type de aerator.
• Tank pressurisation valve (6 psi) ensures adequate
pressure at inlet to oil pressure pump.
• Strainer in tank outlet to pressure pump.
• Provides mounting for scavenge filter and master
magnetic chip detector (MCD).
The oil tank has the following for oil capacity;
• Tank capacity is 29 US quarts.
• Usable oil 24 US quarts.
There is an anti siphon tube that supplies a small flow of
oil back to the tank.
This flow of oil splashes across the sight glass providing a
cleaning action that prevents the build up of impurities.
On early A1 engines the oil tanks were fitted with a
Prismalite oil level indicator, no sight glass was fitted.

Pressure Pump and Pressure Filter Assembly
Purpose
The pressure pump is designed to produce oil pressure for
distribution in the bearing chambers and high speed
external gearbox.
The pressure filter gives initial filtration of the oil as it
leaves the oil tank.
Location
The pressure pump and filter are one assembly. They are
located on the front face of the high speed external
gearbox. Mounted to the left hand side.
Description
The pressure pump and filter assembly has the following
features;
• Cold start pressure limiting valve.
• Flow trimming valve (Line adjustment not permissible).
• 125 micron filtration strength pressure filter.
• Oil priming ports.
• Anti drain valve.

Air/Oil Heat Exchanger (Air Cooled Oil Cooler)
Purpose
The air cooled oil cooler acts as a second cooler for the oil
system.
The heat management system of the EEC controls the
operation of this unit.
Location
Attached to the fan casing on the lower RH side.
Description
The air oil heat exchanger is normally closed when the
engine fuel and oil temperatures are operating within their
required temperature ranges.
During certain conditions of engine operation the fuel and
oil temperatures may experience high temperatures.
The air cooled oil cooler acts to cool the oil in order for the
oil to cool the fuel.
The air cooled oil cooler has the following features;
• Corrugated fin and tube with a double pass design.
• Oil by pass valve.
• Modulated air flow as commanded by EEC (heat
management system) Air flow regulated by air control
valve.
• Electro-hydraulic servo valve operated system.
Note:
The oil has a continuous flow through the air cooled oil
cooler. This is regardless of wether the valve is open or
closed.
Main oil cooling is carried out by the fuel oil heat
exchanger.

Fuel/Oil Heat Exchanger
Purpose
The purpose of the fuel oil heat exchanger is to;
• Cool the engine oil.
• Heat the fuel.
Location
The location of the fuel oil heat exchanger is on the left
hand side of the intermediate case.
Description
The fuel oil heat exchanger is a;
• Single pass fuel flow.
• Multi pass oil flow.
Forms an integral unit with the LP fuel filter.
A differential pressure relief valve permits oil to by-pass
part of the cooler if the oil pressure is high during initial
engine running, following a cold start.

Scavenge Pumps Unit
Purpose
Returns scavenge oil to the tank.
Location
All 6 scavenge pumps are housed together as a single unit
on the rear of the high speed external gearbox, left hand
side.
Description
The scavenge pumps assembly consists of six gear type
pumps that are designed to retrieve the oil from the
gearbox and bearing chambers. Thus returning the oil
back to the oil tank.
The scavenge pump capacity is determined by the gear
width.

De oiler
Purpose
To separate the air and oil mixture that develops in the
bearing compartments and gearbox.
To return the oil back to the oil tank and eject the air
overboard.
Location
The de oiler is located on the front face of the high speed
external gearbox, fight hand side.
Description
The de oiler has the following features;
• Provides mounting for the No.4 bearing chamber
scavenge valve.
• Overboard vent.
• Provides location for No.4 bearing Magnetic Chip
Detector housing.
The de oiler is also called a centrifugal separator. This due
to the fact that it relies upon the centrifugal force of a
rotating body to separate the oil from the air.
The oil is centrifuged outwards and into the gearbox.
The air is induced inwards an out through the discharge
slots.
.

Scavenge Filter
Purpose
To trap solid contaminants.
Location
Mounted to the rear of the oil tank.
Description
The scavenge filter has the following features;
• The filter is disposable.
• The filter has a by pass valve.
• There is a differential pressure switch monitoring filter
for contamination.
• The master magnetic chip detector housing is located
on the filter housing.
Mesh type filter that is disposable. The scavenge filter has
a nominal 30 micron filtration capability.

No 4 Bearing Scavenge Valve
Purpose
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, of which the de oiler is located on the front
face of the high speed 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
When the valve is at the high flow position the engine is at
low power.
Therefor the valve is fully open and the pressure
differential is maintained across the carbon seal.
This has the effect of allowing the buffer air to pressurise
the bearing compartment without an excessive flow that
could dry the seal.
Low flow
When the valve is at the low flow position the engine is at
high power.
Therefor the valve is fully closed and the pressure
differential is maintained across the carbon seal.
This has the effect of increasing the back pressure of
scavenge oil and vent air that is leaving the bearing
compartment.
Increasing the back pressure maintains the seal differential
within acceptable limits.
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.

Magnetic Chip Detectors (MCD)
Purpose
The magnetic chip detectors (MCDs) give on condition
monitoring of the gears and bearing assemblies. They can
indicate when a wear mechanism is present.
Location
The MCDs are located about the high speed external
gearbox.
The Master MCD is located at the inlet to the scavenge
filter.
Description
A total of 7 MCD’s are used in the oil scavenge system.
The MCDs have the following common features;
• The MCDs are of a bayonet style.
• Dual seal rings.
• Baulking pin preventing complete insertion in the case
of a missing seal.
• Separate removable housing.
Each bearing compartment and gearbox has its own
dedicated MCD (two in the case of the main gearbox).
The No4 bearing is located in the de-oiler scavenge outlet.
An additional MCD is located in the combined scavenge
return line, at the scavenge filter inlet. This is known as the
Master MCD.
The Master MCD is accessible through its own access
panel in the LH fan cowl door.
If the master MCD indicates a problem then each of the
other MCD’s are inspected to indicate the source of the
problem.
Access to the dedicated MCD’s is by opening the L and R
hand fan cowls.

Master Magnetic Chip Detector
Purpose
The master MCD gives indication for all gears and
bearings of the engine.
It allows periodic inspection without the requirement to
inspect all MCDs.
Location
The master MCD can be accessed from a dedicated panel
on the left hand side fan cowl door.
Description
The master magnetic chip detector is located in the
scavenge filter case. The MCD is at the inlet to the filter.
If debris is found on the MCD head then further
confirmation as to the source can be had by inspecting all
the MCDs.
Further confirmation still can be had by inspecting the
scavenge filter.

Differential Oil Pressure Transmitter and Low Oil
Pressure Warning Switch
Purpose
The pressure transmitter is designed to give real time
indication to the ECAM of differential oil pressure.
The low pressure switch is designed to give warning of
minimum operating differential oil pressure to ECAM.
Location
The differential oil pressure transmitter and low oil
pressure switch are located on the left hand side
intermediate case. Located in the 10 o’clock position.
Description
The pressure transmitter and low oil pressure switch
differential pressures are sampled from;
• Pressure feed to the no.4 bearing.
• Scavenge oil from the no.4 bearing.

DETV252106
PRESSURE TRANSMITTER AND LOW PRESSURE SWITCH

SECTION 8
HEAT MANAGEMENT SYSTEM

IAE V2500 General Familiarisation Heat Management System
Heat Management System
Purpose
The system is designed to provide adequate cooling, to
maintain the critical oil and fuel temperatures within
specified limits, whilst minimising the requirement for the
fan air offtake.
Location
The following units are located about the engine fan case;
• The engine FCOC.
• The engine ACOC.
• The IDG FCOC.
• The fuel diverter and back to tank valve.
• The aircraft outer wing fuel tanks.
Description
Three sources of cooling are available: -
• The LP fuel passing to the engine fuel system.
• The LP fuel that is returned to the aircraft fuel tanks.
• Fan air.
There are four basic configurations between which the flow
paths of fuel in the engine LP fuel system are varied. The
configurations are;
• Mode 1.
• Mode 3.
• Mode 4.
• Mode 5.
Within each configuration the cooling capacity may be
varied by control valves that 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 that permits/inhibits fuel spill to aircraft
tanks.
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Air/Oil Heat Exchanger Air Modulating Valve
Purpose
To govern the flow of cooling (fan) air through the air/oil
heat exchanger, as commanded by the EEC heat
management control system.
Location
The air/oil heat exchanger (ACOC) is attached to the right
hand side of the engine fan case.
It is in the four o’clock position as viewed from the rear of
the engine looking forwards.
Description
The ACOC is a plate type heat exchanger. It is operated
by an electro hydraulic servo valve mechanism.
The following are features of the ACOC;
• Fail safe position is valve open for maximum cooling.
• Fire seal forms an air tight seal between the unit outlet
and the cowling orifices.
• Control by either channel A or B of EEC.
• Valve position feed back signal via LVDT to each
channel of EEC.
• Valve positioned by fuel servo pressure acting on a
control piston.
• Fuel servo pressure directed by the electro hydraulic
servo valve assembly which incorporates a torque
motor.
The valve is operated via signals from the EEC heat
management system.
The electro-hydraulic servo valve directs a controlling fuel
pressure to the operating piston.
Depending on which side of the piston the fluid is present
depends whether the valve opens or closes.
An LVDT gives positional feedback to the EEC of the
valves position.
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Air Control Valve Electro Hydraulic Servo Valve
(EHSV)
Purpose
To provide the 'muscle' to move the air control valve to the
EEC commanded position.
Location
Bolted to the air control valve casing.
Description
Two stage directional flow valve. The stages are;
• Stage 1 is an electrically activated torque motor and
'Jet pipe'.
• Stage 2 is a spool valve.
The following are features of the EHSV;
• Two independent torque motor windings - one
connected to each channel of EEC.
• Operation, from either channel of EEC.
• Jet pipe protected by 90 micron filter.
• Biased to ensure air control valve fully open at engine
start condition.
• Single fuel servo supply from fuel metering unit (FMU).
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Fuel Diverter and Back to Tank Valve
Purpose
The fuel diverter valve and the back to tank valve together
form a single unit. Command signals of the EEC control
the two valves.
The two valves in turn manage the flow of high pressure
fuel. This is done to optimise the heat exchange process
that takes place between the fuel and oil.
Location
The unit is bolted to the rear of the fuel/oil heat exchanger.
Description
Fuel Diverter Valve
This valve is a two position valve and is operated by a dual
coil solenoid. The control signals to energise/de-energise
the solenoid come from the EEC
• Solenoid energised - mode 1 or 3.
• Solenoid de-energised (fail safe position) mode 4 or 5.
Back to Tank Valve
This valve is a modulating valve and will divert a proportion
of the LP fuel back to the aircraft tanks as controlled by the
EEC.
The interface between the EEC and the valve is a
modulating torque motor; the torque motor will direct HP
servo fuel to position the valve.
• Fail safe position is with the valve fully closed - no fuel
return to tank.
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Heat Management System Operation
The following are the four modes of control for the heat
management system. The system is fully automatic as
controlled by the EEC.
The four modes will be in effect when certain
aircraft/engine operating conditions exist.
Mode 1 (normal mode)
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.
Mode 3
Mode 3 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.
Mode 4
Mode 4 is the mode adopted when the burned fuel flow is
low. For example;
Low engine speeds.
High HP fuel pump inlet temperature.
In this mode the fuel/oil heat exchanger is operating as a
fuel cooler. The excessive heat is passed to the engine oil.
The ACOC extracts the heat from the oil that has been
heated up by the hot fuel.
Mode 5
Mode 5 is the mode that is used when the system
conditions demand operation as in mode 3, but is not
permitted due to;
• IDG oil system temperature is excessive.
• The fuel spill to the aircraft tank is not permitted
because of high spill fuel temperatures.
Mode 5 is the adopted position for the fail safe engine
conditions.
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SECTION 9
COMPRESSOR AIRFLOW CONTROL

IAE V2500 General Familiarisation Compressor Airflow Control
Compressor Airflow Control System
Introduction
Description/operation
The engine incorporates two air bleed systems and a
variable stator vane (V.S.V.) system, which together are
used to: -
• Ensure stable airflow through the compressor at "off
design" conditions.
• Ensure smooth, surge free, accelerations and
decelerations (transient conditions).
• Improve engine-starting characteristics.
• In re-stabilising the engine if surge occurs (surge
recovery).
The complete system comprises three sub-systems, which
are: -
• An L.P. compressor air bleed located at engine station
2.5 and known as the Booster Stage Bleed Valve
(B.S.B.V.)
• H.P. compressor air bleeds on stages 7 and 10.
• The V.S.V. system which comprises variable inlet guide
vanes, at the inlet to the H.P. compressor, and 4
stages of variable stator vanes on the A1 and 3 stages
on A5 engines.
The EEC controls all three systems
A schematic overview of the complete airflow control
system is shown below.
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L.P. Compressor Bleed Valve (LPCBV) - A5
Purpose
The LPCBV bleeds air from the rear of the L.P.
compressor at engine station 2.5, the bleed air is vented
into the fan air duct.
The bleed valve provides improved surge margins during
starting, low power and transient operations.
The bleed valve is controlled by the EEC and is fully
modulating, between the fully open and fully closed
positions this is a function of: -
• N1 corrected speed
• Altitude
• Aircraft forward speed (Mach Number)
For starting the bleed valve is fully open and will
progressively close during engine acceleration, during
cruise and take off the valve is fully closed. For
decelerations and engine operation in reverse thrust the
valve is opened. In the event of an engine surge the valve
is opened to enhance recovery.
LPCBV Mechanical Arrangement
The L.P. Compressor Bleed Valve is a continuous ring
Type valve, which rotates and slides forward to open and
rearward to close. Ten support arms support the ring. Two
of the support arms are driven via a lever and actuating
rod by both the LPCBV master actuator and the slave
actuator.
The two actuators utilise H.P. fuel pressure (from the
FMU) as the hydraulic medium and are hydraulically linked
to ensure simultaneous movement. The master actuator
interfaces with the EEC via a torque motor control and
LVDT feedback.
The mechanical linkage is shown below.
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Booster Stage Bleed Valve
Component Description Actuators
The two B.S.B.V. actuators utilise H.P. fuel as a hydraulic
operating medium.
The actuators are located on the rear of the intermediate
casing on either side of the H.P. compressor, as shown
below.
Only one of the actuators, the one on the left hand side,
interfaces with the EEC This actuator is called the Master
actuator, the right hand side actuator is called the Slave
actuator.
The two actuators are hydraulically linked by two tubes,
which pass across the top of the H.P. compressor case.
The master actuator incorporates a Linear Variable
Differential Transducer (L.V.D.T.) which transmits actuator
positional information back to the EEC
The slave actuator incorporates two overload relief valves,
which prevent overpressurisation of the actuators in the
case of faults, such as a mechanically seized actuator.
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Variable Stator Vane System (VSV) A5
Introduction
Variable Incidence Stator Vanes control the entry of air
into the H.P. compressor. The variable vanes control the
angle at which the air enters the first five stages of the
H.P. compressor.
The angle varies with the H.P. compressor speed (N2);
this reduces the risk of blade stall and compressor surge.
The four stages of variable incidence stators comprise inlet
guide vanes to stage 3 and stage 3, 4, and 5 stator vanes.
Mechanical Arrangement
Each vane has pivots at its inner and outer ends, which
allow the vane to rotate about its longitudinal axis.
The outer end of each vane is formed into a shaft, which
passes through the compressor case and is attached by a
short lever to a Unison ring, (one unison ring for each
stage).
Short rods to a crankshaft connect the four unison rings. A
short rod to an actuator that utilises H.P. fuel as a
hydraulic operating medium connects the crankshaft.
Signals from the EEC direct H.P. fuel to extend/retract the
actuator. Actuator movement causes the crankshaft to
rotate, and, through the unison rings, reposition the
variable stator vanes.
The actuator incorporates an L.V.D.T. which signals
actuator positional information back to the EEC
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Handling Bleed Valves Introduction
Handling bleed valves are fitted to the H.P. compressor to
improve engine starting, and prevent engine surge when
the compressor is operating at off-design conditions.
A total of four bleed valves are used, three on stage 7 and
one on stage 10.
The handling bleed valves are ‘two position’ only - fully
open or fully closed, and are operated pneumatically by
their respective solenoid control valve.
The solenoid control valves are scheduled by the 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.
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Handling Bleed Valves - Location
The diagram below shows the location of the four bleed
valves and the solenoid control valve.

Handling Bleed Valves - Operation
The bleed valves and the solenoid control valves all
operate in the same manner. The operation of one bleed
valve only is described.
Bleed Valves
The bleed valve is a two-position valve and is either fully
open or fully closed. The bleed valve is spring loaded to
the open position and so all the bleed valves will be in the
correct position - open - for engine start.
When the engine is started the bleed air will try to close
the valve. The valve is kept in the open position by servo
air (P3) supplied from the solenoid control valve, (solenoid
de-energised) as shown below.
The EEC will close the bleed valves at the correct time
during acceleration. The bleed valve is closed by the EEC,
which energises the solenoid control valve, as shown
below.
Energising the solenoid control valve vents the P3 servo
air from the opening chamber of the bleed valve, and the
valve will move to the closed position.
During an engine deceleration the reverse operation
occurs and the bleed valve opens.

Compressor Airflow Control
Handling Bleed Valves – Troubleshooting Continued:
If a bleed valve fails to close when required to do so, under
certain conditions the engine may exceed the
recommended EGT operating limits thus preventing the
aircraft from taking off.
This will be caused as a result of the EEC trying to achieve
Take-off EPR but with a reduced volume of air being
supplied to the combustion chamber for mixing with fuel,
ignition and subsequent expansion.
Therefore the EEC makes up for the shortfall in the
available volume of air and simply demands the FMU to
provide more fuel to compensate.
The resultant ‘over-fuelling’ provides the required EPR, but
with the penalty of increased EGT.
TSM Supporting data 75-00-00-301 (Bleed Valve
Troubleshooting Hints) give comprehensive
recommendations in the diagnosis of bleed valve related
problems.
Rigorous troubleshooting would reduce large number of
NFF cases.
The lubrication of bleed valves should not be carried out.
Engine Parameter shift/mismatch during climb/cruise
Engine Parameter shifts due to an open bleed valve that
are not noticed at engine start are more likely to become
evident at higher EPR power settings.
This increases the likelihood of an EGT
Exceedance/Overlimit.
Additionally, an open bleed valve as a result of bleed valve
system problems will also result in unexplained engine
parameter shifts.
Possible causes:
• Bleed valve(s) stuck open (Non-detected FADEC fault).
• Solenoid valve sticking in the de-energised position
(Non-detected FADEC fault).
• An electrical failure of the solenoid valve which results
in the solenoid moving to the de-energised position
(FADEC fault).

Compressor Airflow Control
Handling Bleed Valves – Troubleshooting Continued:
Engine Operation Impact (Transient Manoeuvres and
Surge Recovery).
Engine Operation Impact
During transient manoeuvres (acceleration/deceleration)
and surge recovery, HP compressor stability is maintained
by opening particular bleed valves as defined by the EEC
logic.
For transient manoeuvres on the ground, the 7C are
opened and during flight both the 7A and 7C are opened.
In both cases the 7C are opened based upon a transient
detect and is closed after a set period of time has elapsed.
There are no valves opened at take-off power or steady
state cruise. For surge recovery, the 7A, 7C and 10 stage
bleed valves are opened to maintain compressor stability.
Possible causes
Engine problems (stall/surge) on transient operation can
be the result of:
• Bleed valve(s) not being opened during the transient
(acceleration/deceleration)
• Bleed valve closing early.
• Bleed valves not being open can be due to the bleed
valve sticking in the closed position (Non-detected
FADEC fault), or the solenoid sticking in the energised
position (Non-detected FADEC fault).
• Bleed valves closing early can be due to the solenoid
sticking such that the full de-energised position is not
obtained (Non-detected FADEC fault), bleed valve
internal seal wear and leakage of P3 servo air.

SECTION 10
ENGINE SECONDARY AIR SYSTEMS

IAE V2500 General Familiarisation Engine Secondary Air Systems
Engine Secondary Air systems Introduction
Purpose
The secondary air systems serve the function of the
engine and aircraft.
Description
The secondary air system is made up of the following;
• Active clearance control system (ACC).
• 10th
stage make up air system.
• Aircraft services bleed system.
• Air cooled air cooler (ACAC) for the no.4 bearing
cooling and sealing.
Active Clearance Control (ACC)
The system improves engine performance by ensuring that
the HPT and LPT operate with optimum turbine blade tip
clearances.
This is achieved by directing a controlled flow of cooling air
to reduce the thermal growth of the turbine casings.
This minimises the increase in turbine blade tip clearances
which otherwise occurs during the climb and cruise
phases.
10th
Stage Make Up Air System
The purpose of this system is to provide additional cooling
airflow to the HPT stage 2 disc and blades.
The cooling air used is taken from the 10th stage manifold
and is controlled by a two position pneumatically operated
valve.
The valve position is controlled by the EEC as a function of
corrected N2 and altitude.
Aircraft Services Air Offtake System
The engine supplies the aircraft with bleed air taken from
HPC stages 7 and 10.
The air that is taken from the engine is used for the
following;
• Cabin pressurisation and conditioning.
• Wing anti icing.
• Engine cross feed starting.
• Hydraulic system pressurisation.
• Water system pressurisation.
The required air is bled from the HPC of each engine.
Air cooled air cooler (ACAC)
HPC12 air is used for cooling and sealing the no.4 bearing
in the centre bearing compartment.
The ACAC pre cools the HPC12 air prior to the air being
passed to the centre bearing compartment.
The cooled HPC12 air is commonly known as buffer air.
The ACAC uses fan bypass air as the cooling medium.

SECONDARY AIR SYSTEMS INTRODUCTION
DETV250280
HPC STAGE 7 AIR OFFTAKE HPC STAGE 10 AIR OFFTAKE
MAKE UP AIR VALVE
AIR COOLED AIR
COOLER (ACAC)
ACC ACTUATOR
MODULATING AIR
CONTROL VALVE
ACC TUBES FOR THE
LPT AND HPT

Active Clearance Control (ACC)
Purpose
The system improves engine performance by ensuring that
the HPT and LPT operate with optimum turbine blade tip
clearances.
This is achieved by directing a controlled flow of cooling air
to reduce the thermal growth of the turbine casings.
This minimises the increase in turbine blade tip clearances
which otherwise occurs during the climb and cruise
phases.
Location
The ACC system is mainly located about the core engine.
Description
The ACC system consists of the following items;
• LPT and HPT cooling manifolds.
• Operating actuator with LVDT feedback.
• Modulating air control valve unit.
• EEC control.
• Fan bypass air cooling medium.
The EEC controls the ACC system by monitoring the
following parameters;
• Corrected N2.
• Aircraft altitude.
From these two parameters the EEC will signal the
operating actuator.
The operating actuator moves in a linear motion by the
influence of fuel pressure. The EEC receives feedback of
the actuator position by an LVDT.
The operating actuator moves a linkage that controls the
valves in the modulating air control unit for the LPT and
HPT case cooling.
The HPT and LPT casings are cooled by fan bypass air
that is ducted from the fan bypass.
Failsafe Position
Upon the event of fuel pressure loss and/or EEC power
failure the ACC modulating air control valves will adopt the
following positions;
• LPT is –44% open.
• HPT is closed.

Active Clearance Control Components
ACC Actuator
Purpose
The ACC actuator provides the movement to the
modulating air valves so it can vary the LPT and HPT
cooling airflows.
Location
The ACC actuator is located on the right hand side of the
core engine in the 5 oclock position. It is mounted on the
compressor casing.
Description
The ACC actuator consists of the following;
• Linear motion two directional piston.
• Dual track LVDT.
• Electro hydraulic torque motor.
• Filter.
The ACC actuator receives signals from the EEC. The
torque motor will direct high pressure fuel to one of the two
sides of the piston. This is dependant on the EEC
command signal.
Piston movement will result in a movement in the push pull
rod that links the ACC actuator and the modulating air
valve.
The LVDT will feedback the piston position to the EEC.
At engine shut down or nil servo pressure the ACC
actuator will assume the failsafe position. A spring in the
ACC actuator will force the piston to the failsafe position.
Modulating Air Control Valve Unit
Purpose
The modulating air control valve receives ducted air from
the fan bypass stream and regulates, as per ACC actuator
input, the flow rate to the LPT and HPT ACC manifolds.
Location
The modulating air control valve is located on the right
hand side of the core engine in the 5 oclock position. It is
mounted on the turbine casing.
Description
The modulating air control valve has two separate valves.
They are;
• HPT valve.
• LPT valve.
The two valves are designed to operate to allow the
optimum airflow to the respective casings.
The failsafe position is;
• HPT is closed.
• LPT is –44% open.

Page 10-6
ACC ACTUATOR
Revision 1
DETV250282
ACC ACTUATOR AND MODULATING AIR VALVE
FAN BYPASS
AIR INLET
MODULATING AIR
CONTROL VALVE TORQUE MOTOR
LVDT FEEDBACK
SERVO FUEL SUPPLIES

Active Clearance Control Components
HPT ACC Manifold
Description
The HPT ACC manifold is designed to impinge cooling air
onto the turbine casing about the rotor blade path. This is
done to reduce the rotor blade tip to rotor path gap.
Location
The HPT ACC manifold is located on the HPT casing.
Description
The assembly consists of left and right hand tube
assemblies, which are a simple push fit into the manifold.
The tube assemblies are sealed off at their upper ends.
Air from the air control valve enters the manifold and is
directed to the left and right tubes.
Air outlet holes on the inner face of the tubes direct the air
onto the HPT casings.
LPT ACC Manifold
Description
The LPT ACC manifold is designed to impinge cooling air
onto the turbine casing about the rotor blade path. This is
done to reduce the rotor blade tip to rotor path gap.
Location
The LPT ACC manifold is located on the LPT casing.
Description
The assembly consists of upper and lower tube
assemblies with integral manifolds; both ends of the
cooling tubes are sealed.
Air from the air control valve enters a supply tube, which
then splits to feed air into two tubes that supply the upper
and lower manifolds. The manifolds direct the air into the
cooling air tubes.
Air outlet holes on the inner surfaces direct the air onto the
LPT cases.

ACC LPT MANIFOLD
HALVES
ACC HPT MANIFOLD
HALVES
DETV250283
HPT AND LPT DISTRIBUTION MANIFOLDS

10th
Stage Make Up Air System
Purpose
The 10th
stage make up air valve system allows additional
cooling air to the HPT stage 2 disc and blades.
Description
The 10th
stage make up air valve system consists of the
following components;
• EEC control.
• Make up valve control solenoid.
• Two position type on/off valve.
• Microswitch positional feedback.
The EEC uses input signals of;
• Corrected N2.
• Altitude.
These inputs are used by the EEC to signal the control
solenoid to operate.
The control solenoid manages the flow of P3 (HPC stage
12) air for the pneumatic operating medium.
The two position make up air valve either opens to flow
stage 10 air or closes for no flow.
The solenoid is de energised when the valve is in the open
position. This is the fail safe position.
The microswitch gives a positional feedback signal to the
EEC indicating of either;
• Valve open.
• Valve closed.
The valve is open for all conditions of flight/engine
operation except for cruise.
In cruise the valve is closed.

10TH
Stage Make Up Air System Components
10th
Stage Make Up Valve
Purpose
The 10th
stage make up valve purpose is to supply air to
supplement the normal airflows around the no.4 bearing
housing and the HPT disc and blades.
Location
The 10th
stage make up air valve is located at the top of
the HPC casing.
Description
The 10th
stage make up air valve consists of the following
components;
• Operating piston.
• Microswitch feedback.
• Valve body.
The valve is a two positional type. It can either flow HPC
stage 10 air or cut it off from the engine. There is no
modulation.
The operating piston is spring loaded to the open position
when servo air is not present in the piston chamber.
Servo air is used to close the valve.
The micro switch gives positional feedback of the piston
position hence the valves condition.
The fail safe position is valve open.
Control Solenoid valve
Purpose
The control solenoid valve purpose is to manage the flow
of servo air pressure to the make up air valve.
Location
The control solenoid is located on the right hand side of
the fan case approximately in the 4 o’clock position.
Description
The control solenoid consists of the following components;
Solenoid pack.
Pilot valve.
Valve body.
The solenoid control valve will direct the flow of servo air
pressure to port when it is de-energised.
The solenoid control valve will direct the flow of servo air
pressure to the make up valve when it is energised.

Aircraft Services Air Offtake System
Purpose
To provide the following aircraft systems with engine
ducted air supply;
• Cabin pressurisation and conditioning.
• Wing leading edge anti icing.
• Engine cross bleed starting.
• Hydraulic system pressurisation.
Location
The bleed air offtakes are taken from;
• HPC stage 7 for high power conditions.
• HPC stage 10 for low power conditions.
Description
HPC air is taken from the engine and ducted towards the
aircraft services.
The HPC stage 7 offtake has a non return valve installed
before the two offtakes join. The NRV protects against
HPC stage 10 air from reverse flowing back into the HPC
stage 7 of the engine.
The HPC stage 10 offtake has a control valve called the
high pressure valve (HPV).
After the two offtakes come together as one there is a
pressure regulating valve (PRV). The PRV is controlled by
a switch located in the flight deck.
The over pressurisation valve (OPV) protects the system
against excessive pressures.
The precooler prepares the bleed air to an acceptable
temperature before it enters the environmental control
system (ECS).
The pre cooler utilises fan bypass air to cool the HPC
bleed air.
The temperature limiting thermostat (TLT) controls the
PRV when an over temperature has been experienced.
The temperature controlling thermostat controls the pre
cooler valve but if bleed temperature cannot be maintained
it will signal for the PRV to close.
The bleed monitoring computer controls the functions of
bleed air system making the system fully automatic.

Aircraft Services Air Offtake System Operation
The following is the operation of the aircraft services air
offtake system.
The bleed monitoring computer (BMC) controls the
opening and closing of the;
• Over pressure valve (OPV).
• Pressure regulating valve (PRV) (this is also controlled
by a selector switch in the flight deck.
• Fan air valve (FAV).
• High pressure regulating valve (HPRV).
The bleed monitoring computer monitors the following;
• Temperature limiting thermostat (TLT).
• Temperature controlling thermostat (TCT).
• Pressure sensor downstream of the HPRV.
• Pressure sensor downstream of the PRV.
The selection of HPC stage 7 (also called IP bleed air) or
HPC stage 10 is automatically done by the BMC.
High pressure regulating valve (HPRV)
The HPRV will regulate HPC stage 10 air to 36 psi.
The HPRV will close if downstream pressure is greater
than 100 psi and/or HPC stage 7 greater than 36 psi.
Pressure regulating valve (PRV)
The PRV is spring loaded open when there is no
pneumatic air available.
The PRV will regulate either airflow to 44 psi.
The PRV starts to open at approximately 8 psi.
The PRV will close when HPC stage 7 air is greater than
36 psi.
Over pressure valve (OPV)
The OPV will start to close at 75 psi.
The OPV will be fully closed at 85 psi.
The OPV will reopen at 35 psi.
Temperature limiting thermostat (TLT)
TLT normal indication to BMC is 200 deg.c.
The TLT over temperature is 245 deg.c. Above this value
will reduce PRV pressure to 17.5 psi.
The TLT maximum temperature is 257 deg.c. Above this
value and the PRV is closed.
Temperature controlling thermostat (TCT)
The TCT maximum temperature is 257 deg.c. Above this
value and the PRV is closed.
The pressure sensors feedback pressure signals to the
BMC.

Air Cooled Air Cooler (ACAC)
Purpose
The ACAC purpose is to pre cool HPC12 air. The ACAC
uses fan bypass air as the cooling medium.
Location
The ACAC is located on the turbine casing. Bottom left
hand side in the 5 o’clock position.
Description
The ACAC is a fin and tube type design.
The fan bypass airflow that is utilised by the ACAC
extracts heat from the HPC12 air.
The HPC12 air is taken off the engine through a singular
tube.
The HPC12 air enters the ACAC and the heat exchange
process takes place between the fan bypass air and the
hot HPC12 air.
The fan bypass air is ejected to atmosphere.
The cooled HPC12 air leaves the ACAC and is distributed
to the centre bearing compartment through three tubes.
The tubes enter the diffuser casing in three positions. They
are;
• 12 o’clock.
• 3 o’clock.
• 9 o’clock.
The cooled HPC12 (buffer) air enters the cooling jacket of
the centre bearing chamber.
The buffer air protects the no.4 bearing from excessive
heat exposure.
The buffer air enters the bearing compartment to prevent
oil loss.
This also pressurises the bearing chamber to allow the oil
and air mix to leave the bearing chamber and enter the de
oiler.
The centre bearing compartment does not have an oil
scavenge pump.

FAN BYPASS AIR INLET
AIR COOLED AIR
COOLER (ACAC)
AIR COOLED AIR COOLER (ACAC)
HPC12 (BUFFER) AIR INTO
THE CENTRE BEARING
COMPARTMENT
FAN BYPASS AIR
OVERBOARD DUMP
CENTRE BEARING
COMPARTMENT COOLING
JACKET
HPC12 AIR INTO
ACAC
COOLED HPC12 AIR
OUT OF THE ACAC
DETV250376

SECTION 11
ENGINE ICE PROTECTION SYSTEM

IAE V2500 General Familiarisation Engine Ice Protection System
Engine Ice Protection System Introduction
Purpose
Ice may form in the inlet cowl when the engine is operating
in conditions of low temperature and high humidity.
Ice build about the inlet cowl leading edge could affect
engine performance and could cause engine damage from
ice ingestion.
To prevent ice formation ice protection systems have been
incorporated into the engine.
• The inlet cowl leading edge is thermally ice protected.
• The P2/T2 probe mounted in the inlet cowl is thermally
ice protected.
• The spinner of the fan module is ice protected by a
flexible rubber tip.
Description
The engine ice protection system description is as follows;
Inlet cowl ice protection
The inlet cowl is thermally heated to prevent ice formation
at the leading edge of the intake lip.
The ice protection system for the inlet cowl is controlled
from the flight deck by a selector switch. The switch will
control the opening and closing of the TAI valve.
The valve will allow the airflow taken from the HPC stage 7
to flow to the distribution manifold in the inlet cowl leading
edge lip.
The distribution manifold will allow hot air to enter the inlet
cowl leading edge.
The excessive air is ejected overboard via an outlet
located on the right hand side of the inlet cowl.
Fault indications for the ice protection system are as
follows;
• The flight deck anti icing selector switch illuminates.
• An ECAM warning message is generated.
P2/T2 probe heater
The P2/T2 probe is continuously heated during engine
operation by an integral 115V heating coil.
Spinner
The spinner is protected against ice build up by a solid
rubber nose tip that vibrates naturally to break up and
dislodge the ice immediately it starts to form.
Ground running
Icing conditions may occur when the outside air
temperature (OAT) is less than;
• 5.5 deg.c (42 deg.f).
• The humidity is high for example rain, sleet, snow, fog
(visibility is less than one mile).
If the above conditions exist the ice protection system
must be operated as soon as the engine stabilises at low
idle conditions after an engine start.

Component Description
Anti Icing Control Valve
Purpose
The anti icing control valve allows the flow of HPC stage 7
air to enter the TAI manifold in the intake cowl.
Location
The anti icing valve is located on the right hand side of the
fan case in the 4 o’clock position.
Description
The anti icing control valve has the following function;
• On/off selection from the flight deck to flow warm air to
the TAI manifold.
The anti icing valve is made up of the following items;
• Valve body.
• Linear moving piston.
• Control solenoid.
• Air filter.
• Butterfly valve.
• Micro switch.
• Manual override (as per MMEL requirements).

Engine Ice Protection SystemIAE V2500 General Familiarisation
Revision Page 11-4
ANTI ICE VALVE
FILTER
LOCKOUT PIN
ELECTRICAL
CONNECTOR
VALVE BODY
ENGINE ANTI ICE VALVEENGINE ANTI ICE VALVE
1
DETV250273DETV250273

Operation
Valve closed
The control solenoid is energised.
The ball valve of the control solenoid is hard against the
ambient vent outlet. This prevents upstream air from
escaping to vent.
The pressure acting on the piston at position area A is
greater than the pressure acting on position area B. This
along with spring pressure holds the butterfly valve in the
closed position.
Valve open
The control solenoid is de energised.
The ball valve is no longer held by the control solenoid
against the ambient vent. The ball valve moves by spring
pressure against the orifice which allows upstream air to
enter area A.
This now prevents air passage to area A of the piston.
The air pressure now remains at piston area B only. This
pressure is greater than the spring pressure alone therefor
the piston moves against the spring pressure.
The resultant movement opens the butterfly valve and
allows HPC stage 7 air to flow towards the TAI manifold.
Fail safe position
The fail safe position is as follows;
• Solenoid de energised.
• Servo air pressure at piston area B only.
• Butterfly valve in the open position.
Manual override
The valve has provision for being secured in either the;
• Locked position.
• Open position.
This requirement is necessary when a valve has failed.
The MMEL will advise of the actions required to allow
despatch of the aircraft.

ECAM Indications
The engine anti icing valve has a micro switch which will
feedback the valve position in relation to the selector
switch position.
When the anti icing system is functioning normally a
caption will appear;
• ENG A. ICE
The caption appears on the upper ECAM lower right hand
side.
The selector switch will have the on indicator illuminating
in the colour of blue.
If a disagreement exists between the selector switch and
the microswitch output signal to the EEC a fault has been
detected. The fault detection occurs when one of the
following situations exist;
• A valve failure to open.
• A valve failure to close.
The fault portion of the selector switch will illuminate in the
colour of amber when a disagreement exists.
Th upper ECAM screen will display a WARNING and
STATUS message of;
• ENG 1(2) VALVE CLSD.
• ENG 1(2) VALVE OPEN.
These messages relate to the switch position and the
intended valve position.
The messages are engine specific.
The message is status related therefor it becomes a
despatch critical message. Advice from the MMEL is
required.
Note
It is advisable not to lock the TAI valve in the open position
for the higher thrust engines.
If the valve fails in the closed position it is advisable to
avoid icing conditions.
If the valve fails in the open position there will be a thrust
limit penalty.
One or both may be inoperative provided the valve has
failed in the open position and the performance penalties
are applied and OAT does not exceed ISA +35 deg.c.
Engine anti ice valve fault;
For ER operations only one valve allowed to be failed in
the closed position and providing the aircraft is not
operating in icing conditions.
Engine anti ice fault light;
One or both valves may be inoperative provided the faulty
valve is deactivated and considered inoperative in the
open position.

Anti Ice Valve Manual Override
The anti ice valve can be manually locked;
• Open.
• Close.
This to allow for despatch of the aircraft for continued service.
Despatch of the aircraft is according to MMEL instructions.

ANTI ICE
VALVELOCKOUT PIN DEACTIVATION
HOLE
ANTI ICE VALVE MANUAL OVERRIDE
DETV250277
FADEC POWER
SELECT SWITCH
ENGINE START
CONTROL PANEL HAND TUR
POINT
NING
HAND TURNING
POINT

IAE V2500 General Familiarisation Engine Indications
FADEC/Aircraft Interface Introduction
Purpose
The FADEC system supplies the aircraft systems with the
relevant engine data in order to assist the aircraft to carry
out its functions.
Description
The aircraft systems that it uses to interpret the engine
data and display it for flight crew use is;
• Electronic centralised aircraft monitor (ECAM) system.
• Flight warning computer (FWC).
• Data management computer (DMC).
• System data acquisition concentrator (SDAC).
• Engine interface unit (EIU).
The ECAM system receives engine and aircraft data and
displays this on two cathode ray tubes (CRTs). The ECAM
system is designed to give the flight crew primary and
secondary engine/aircraft data.
The flight warning system monitors all data that relates to
a class 1 type indication. This is regarded as the highest
priority type annunciation.
The system display can be transferred to the navigation
display (ND) CRT by a selector switch.

ECAM Indications
The ECAM system displays both engine and aircraft data.
The upper and lower ECAM CRTs display engine/aircraft
data in digital and analogue form.
Upper ECAM CRT
The upper ECAM CRT will display the following engine
and aircraft data;
• EPR command.
• EPR actual.
• EGT.
• N1 rotor shaft speed.
• N2 rotor shaft speed.
• Fuel flow.
• Fuel on board (FOB).
• Slat and flap position.
The upper ECAM CRT display is also used to give warning
information of class 1 type alerts. This is given in the form
of a message.
Note:
A1 series of engines have bump switches to enhance the
take off performance. Whenever the bump is selected the
alpha B will appear next to the EPR gauge.
The B will disappear when;
• Mn of 0.45 is reached.
• The aircraft altitude has reached 15000 ft.
Lower ECAM CRT
The lower ECAM CRT will display the following engine and
aircraft data;
• Fuel used.
• Oil quantity.
• Oil pressure.
• Oil temperature.
• Engine vibration for N1 and N2.
• Nacelle air temperature (NAC).
• Total air temperature (TAT).
• Static air temperature (SAT).
• Aircraft gross weight.
Note:
The NAC will only appear on the lower ECAM CRT when
an exceedance has occurred.
During engine start up the start air valve position, bleed air
pressure and igniter selection are displayed in the NAC
position.
No.4 bearing scavenge valve indication is a pre
modification standard.

ECAM Indications Upper CRT
Engine pressure ratio (EPR)
Actual EPR indication is green.
EPR max is the thick amber index.
EPR TLA angle is the white circle.
Transient EPR is the blue arc.
Idle indication flashing green for 10 seconds then steadies
for both engines at idle in flight.
REV indication for thrust reverser status.
Exhaust gas temperature (EGT)
Actual EGT indication is normally green.
When EGT exceeds 610 deg.c the indication remains
green the pointer pulses amber.
The values pulse red when EGT at red line.
EGT over-temperature is the red mark. If an over-
temperature occurs a red mark appears at the max value
achieved. It will disappear after a maintenance action
through the MCDU.
Max permissible EGT red line at beginning of red arc.
During engine start the max permissible will be at starting
value.
Max EGT is the thick amber index. This is not displayed
during engine start.
N1
Actual N1 indication is normally green.
Pulses red when N1 limit is exceeded.
Pulses amber when N1 exceeds N1 rating limit in N1
mode.
Max permissible N1 is the red line indication at beginning
of the red arc.
N1 overspeed occurs a red mark appears at the max value
achieved. It will disappear after a maintenance action
through the MCDU.
N2
Actual N2 indication is normally green.
N2 goes red when limit is exceeded also a red cross
appears next to the digital value.. It will disappear when a
maintenance action through the MCDU.
N2 indication is highlighted and boxed grey during engine
start sequence.
Thrust limit mode
TOGA, FLX, MCT, CL and MREV are displayed in blue.
EPR rating limit is displayed in green.
De rate temp indication is displayed in blue.
Actual fuel flow
Actual fuel flow is displayed in green and gives real time
indication of fuel flow for left and right engines.

ECAM Indications Lower CRT
Fuel Used
Fuel used indication is normally green.
Freezes at last value when engine is shut down and resets
at next engine start.
Last two digits are dashed if fuel used indication is
inaccurate due to loss of fuel flow for 1 minute.
Oil quantity
Oil quantity indication is normally green.
At 5 quarts the advisory level is reached and the indication
pulses.
At 7 quarts and above the pulsing stops.
Oil pressure
The oil pressure indication is normally green.
The indication pulses if the oil pressure exceeds 390 psi
increasing or 385 psi decreasing.
Between 80 and 60 psi the indication is amber.
Below 60 psi the indication is red.
Oil temperature
Oil temperature indication is normally green.
The indication pulses above 156 deg.c increasing and 150
deg.c decreasing.
The indication becomes amber with an ECAM warning if
temperature exceeds 165 or above 156 for more than 15
minutes or the temperature is below minus10 deg.c.
Ignition and start valve position
The ignition and start valve positions are displayed during
start up only.
The selected igniters are displayed in green.
The bleed pressure indication is normally green.
If the pressure goes below 21 psi or suffers an over
pressure the indication is amber as long as the start valve
is not closed.
Nacelle temperature (NAC)
The NAC is not normally displayed. It will appear pulsing
green (advisory) when an exceedance has occurred.
Vibration
Vibration indication is normally green.
The indication pulses green if vibration is above 5.0 units
(advisory).
Oil filter and fuel filter
No indication if both filters are normal.
The message CLOG will appear in amber when the
differential pressure across the filter has been exceeded.
An ECAM message will also be generated.

Revisi 1 Page 12-8on
DETV250378
LOWER ECAM CRT

ECAM System Fault Monitoring
Purpose
The ECAM system is designed to constantly monitor for
engine /aircraft parameter deviations from the normal. The
deviation can then be annunciated to the flight crew.
Description
Normal parameter indication is;
• Green.
Approaching parameter deviation the indication is;
• Flashing green.
Warning condition
The parameter deviation indication is;
• Steady red indication.
• Master warning light on glare shield.
• Repetitive audible chime.
• ECAM message.
Caution condition
The parameter deviation indication is;
• Steady amber indication.
• Master caution light on glare shield.
• Audible chime.
• ECAM message.
ECAM Messages
The ECAM displayed messages are enunciated to the
flight crew in the order of priority.
The alert level classification for faults is as follows;
• Level 3 Red warning with repetitive chime.
• Level 2 Amber caution with chime.
• Level 1 Amber caution with no chime.
The upper ECAM CRT will display all warning type
messages that are generated. This will display in the left
memo box.
The lower ECAM CRT will display messages of caution
and status.
The Lower ECAM CRT also has the facility to display other
systems of the engine and aircraft.

ECAM System Pages
Purpose
The ECAM pages give the flight crew a detailed parameter
and system status of the aircraft and engine systems.
These pages can also assist in troubleshooting.
Description
There are twelve pages of information covering the
systems of the engine and aircraft. The pages are as
follows;
• Engine.
• Bleed air.
• Cabin pressure.
• Electric’s.
• Hydraulics.
• Fuel.
• Auxiliary power unit (APU).
• Conditioning.
• Doors.
• Wheels.
• Flight controls.
• Engine/air.
The pages can be called up by either the flight crew
manually or automatically according to the flight phase the
aircraft is in.
ECAM Status Page
The status page has information of faults that affect the
redundancy of a system.
Status messages do not directly affect the aircraft
operation but reference to the MMEL is required before
aircraft despatch.
By depressing the status select button the status screen
will appear.
Status information that has occurred during flight will be
alerted to the flight crew by a pulsing STS on the upper
ECAM CRT warning memo box.
This occurs when the engines are shut down.

Flight Deck Centre Pedestal
The following are some of the Engine related controls and
interfaces:
1. Captains Multipurpose Centralised Display Unit
(MCDU)
2. Systems Display Control Panel
3. First Officers Multipurpose Centralised Display Unit
(MCDU)
4. Engine No. 2 Thrust Lever
5. Engine No. 2 Master Switch
6. Ignition Mode Selector Switch
7. Printer
8. Engine No. 1 Thrust Lever
9. Engine No. 1 Master Switch

Flight Deck Overhead Panel
The following are some of the Engine related interfaces found
on the Overhead Panel:
1. N1 Mode Selector Switches for No. 1 & No. 2 engine
2. Engine Manual Start Switches for No. 1 & No. 2 engine
3. Engine and Auxiliary Power Unit (APU) Fire Panel
4. FADEC ground power switches for No. 1 & No. 2 engine

Shaft Speed Indicating System
Purpose
The speed indicating system provides signals of;
• N1 shaft speed.
• N2 shaft speed.
The indications are used for ;
• The ECAM CRT display.
• EEC control.
A dedicated signal is used for trim balancing purposes.
Location
The N1 speed sensors are located in the front bearing
chamber mounted on the no.2 bearing support.
The N2 speed indication is the output signals from the
dedicated generator.
Description
The speed indicating description is as follows;
N1 System
The N1 indication is supplied by three pulse probes. The
pulse probes operate by monitoring the passage of a
phonic wheel.
The phonic wheel passage across the pulse probe
generates an output signal relative to a percentage of a
revolution.
Hence if the phonic wheel has 60 teeth then 60 pulses
represents a complete revolution of the N1 shaft.
N2 System
The N2 indication is supplied by a dual output signal from
channel B of the dedicated generator.
• An output goes to the channel B side of the EEC.
• An output goes to the EVMU.
Fan Trim Balance
The fan trim balance probe is located in the same place as
the speed pulse probes. This probe supplies a dedicated
signal for monitoring of LP system unbalance.
The probe is also different from the speed probes. It
cannot be utilised to give N1 speed indication.
The pulse probe monitors a datum tooth of the phonic
wheel. This tooth is in line with the no.1 fan blade.

N1Speed Indicating System Operation
The probes comprise of two pole pieces, a permanent
magnet, and a coil wound on to one of the pole pieces.
The pole pieces span two teeth of the phonic wheel. The
phonic wheel is an integral part of the fan stub-shaft and
has 60 teeth.
As the shaft rotates and the teeth of the phonic wheel pass
the pole pieces and a voltage pulse is produced in the
winding. The number of pulses produced is directly
proportional to the speed of the shaft.
This signal is passed to the EEC and is used to display N1
speed on the flight deck and also for the engine control
circuits as required.
Trim Balance Probe
The signal from this probe is only used during trim balance
operations and provides the phase relationship between
any out of balance forces present and a datum position.
The trim balance probe senses the passage of one
specially modified tooth on the phonic wheel and produces
one pulse per revolution.

Exhaust Gas Temperature (EGT) Indicating System
Purpose
EGT is displayed to the flight deck via the ECAM system to
give the flight crew an indication of the engine
temperature.
This allows the engines to be operated within the
temperature limitations as advised by IAE.
Location
The EGT thermocouples are located at the exhaust outlet.
The EGT T/C leads come together at a junction box
located at BDC of the turbine casing.
Description
The EGT is measured by 4 thermocouples, which are
located in the support struts of the turbine exhaust case
(engine station 4.9).
The 4 thermocouples are connected to the junction box by
a thermocouple harness. The materials used for the
thermocouples and harness are;
• Chromel (CR).
• Alumel (AL).
An extension harness connects the EGT junction box to
channels A and B of the EEC.
Indication
The EGT indication appears on the upper ECAM display
unit. The ECAM provides the EGT indication;
• In analogue dial gauge format.
• In digital format.
EGT is below 610 deg.c.
The actual EGT indication is normally green.
EGT is > 610 deg c.
The indication pulses and changes colour to amber.
EGT is > 635 deg c.
• The indication becomes red.
• The MASTER WARN light comes on, accompanied by
the repetitive audible chime.
The following message appears on the ECAM upper CRT;
EGT OVERLIMIT
• The maximum value reached is memorised.
• A small red line remains positioned on the analogue
scale at that value (max pointer).
Note:
The small and large nuts that secure the EGT leads to the
junction box must torque check and tightened during the A
check until further notice.
Single and dual channel failures have occurred due to
loose EGT securing nuts.

P3/T3 Sensor
Purpose
To give the EEC an input signal of;
• P3 pressure for fuel scheduling and surge detection.
• T3 temperature for trend monitoring.
Location
The P3/T3 sensor is located on the combustor casing at
the one o’clock position.
Description
The P3/T3 sensor is a dual-purpose aerodynamically
shaped probe. It measures the pressure and temperature
of the air stream at the inlet of the diffuser case.
The resultant data is transmitted to the EEC for control
purposes. At the EEC the pressure enters a transducer.
The temperature signal is received as a resistance value.

Engine Pressure Ratio (EPR) Indicating System
Purpose
To indicate to the flight deck a parameter that is the
representation of engine thrust.
Location
The main components of the EPR system is the P2/T2
probe and the P 4.9 pressure rakes.
They are located;
P2/T2 probe in the intake cowl at approximately TDC.
P4.9 pressure rakes are in the exhaust duct of the LPT.
Description
The engine pressure ratio (EPR) is used to set and control
the engine thrust EPR. EPR is;
P4.9
P2
P2 is measured by the P2/T2 Probe.
P4.9 is measured by a pressure rake.
The pressures from these sensors are routed to the EEC.
The EEC processes the pressure signals to form actual
EPR and transmits the EPR value to the ECAM for display
on the upper screen.
Each of the two EEC channels carry out this operation
independently.
EPR Indications
The actual EPR is displayed in green. The display
changes to Amber if EPR exceeds limit.
The display changes to red, with master caution and single
chime if EPR exceeds red line limit (TBD).
The associated indications are:
• EPR maximum has a thick amber line.
• EPR limit maximum EPR value corresponding to thrust
limit mode, which can be any one of the five conditions
that follow;
Take off/go around mode (TO/GA).
Flexible take-off mode (FLX).
Maximum continuous thrust mode (MCT).
Climb mode (CLB).
Flex TO Temperature is an assumed temperature entered
by the flight crew through the MCDU to the FMS facility.
EPR reference is the predicted EPR value according to
TRA.

EPR System P2/T2 Sensor
Purpose
The P2/T2 sensor is a dual purpose probe which
measures the total air temperature and pressure in the
inlet air stream. The temperature and pressure signals are
fed to the EEC.
Location
The sensor is installed at the 12 o'clock position in the air
inlet cowl.
Description
The temperature is measured by two platinum resistance
elements. Each channel of the EEC monitors one of the
elements.
The pressure signal is fed to a pressure transducer in the
EEC.
The sensor is electrically heated to provide anti ice
protection.
The EEC software corrects any temperature signal errors
caused by heating.
Note:
The probe anti icing heater utilises 115V AC from the
aircraft electrical system.
Higher thrust engines (V2533) have a longer probe.
The relay box on the right hand side of the fan case
controls the selection of voltage to the probe heater unit.
EPR System P4.9 Rake
Purpose
The P4.9 rakes send a pressure signal to the EEC for the
EPR system.
Location
The P4.9 rakes are located in the exhaust OGVs. They are
in the 3, 6 and 9 o’clock position.
Description
The P4.9 pressure rakes send a pressure signal down a
common tube to a transducer within the EEC.

Engine Vibration Indicating System
Purpose
The system monitors engine vibration for engine 1 and
engine 2.
Location
The vibration transducer is located on the engine fan case
in the 11 o’clock position.
Description
Monitoring is done by a vibration transducer on each
engine fan case. This produces an electrical signal in
proportion to the vibration detected and sends it to the
engine vibration monitoring unit (EVMU).
Two channels come from each engine. The EVMU
provides signals of;
• Vibration.
• N1 (LP shaft speed).
• N2 (HP shaft speed).
These are displayed on the engine page of the ECAM.
The vibration transducer is installed on the fan case at the
top left side of the engine. It is attached with bolts and is
installed on a mounting plate.
Indications
The engine vibration indications are displayed in green on
the lower ECAM display unit on the engine and cruise
pages.
The ECAM display unit receives the information through
the ARINC 429 data bus via the SDAC 1 and SDAC 2.
If the advisory level is reached, the indication flashes (0.6-
sec bright, 0.3-sec normal).
If the indication is not available, 2 amber crosses replace
the corresponding indication.
Note:
A5 engines have a dual cable.
D5 engines have a single cable.

SECTION 13
ENGINE STARTING AND IGNITION

IAE V2500 General Familiarisation Engine Starting and Ignition System
Engine Starting and Ignition System
Purpose
The engine starting system is to allow the engine to
achieve minimum power conditions (low idle). The engine
requires assistance to achieve this condition.
The engine ignition system provides the electrical spark
that is required to ignite the fuel air mix in the combustor.
The ignition system is used for;
• Engine starting on ground and in flight.
• Prevention of a flame out by providing a continuous
spark during engine running when the aircraft is flying
through a heavy rainstorm for example.
Description
The system comprises of the following;
• Pneumatic starter motor.
• Starter air control valve.
• Dual ignition system.
• Pneumatic ducting.
• Start control panels on the flight deck for auto starts,
manual starts and starter motor operation.
• ECAM indications.
Starting of the engine for the Airbus A319, 320 and 321
can be done either;
• Manually.
• Automatically.
Basic start sequence
Whichever method of starting is selected the EEC has
control.
The start sequence is initiated in the flight deck.
Upon selection for engine start an electrical signal is sent
to open the starting air valve.
The starting air valve opens and admits the air supply into
the starter motor.
The starter motor rotates the high speed external gearbox
that in turn rotates the radial drive shaft (tower shaft) that
in turn rotates the HP system (N2).
As the HP system gathers momentum of rotation the LP
system starts to rotate.
• At 10% N2 the dedicated generator comes on line.
• At 16% N2 the ignition system becomes live.
• At 18% N2 the FMV opens to admit fuel to the
combustor.
At 43% N2 the starter air valve is deselected.
At 50% N2 and above the EEC auto start protection is
cancelled.
At approximately 60% N2 the engine is at minimum power
conditions (low idle).
Note;
Above 50% N2 the command for engine shut down is done
from the master lever only.

Starter Air Duct
Purpose
To provide a means of supplying air to the starter motor.
Location
The starter air duct is located on the right hand side of the
engine fan casing (intermediate module).
Description
Air supplies for the pneumatic starter motor may be
supplied from;
• The aircraft APU.
• Cross bleed from the other engine if already running.
• Ground starter trolley.
Minimum duct pressure for starting should be between 30
and 40 psi.
All ducting in the system is designed for high pressure and
high temperature operation.
Gimbal joints are incorporated to permit working
movement.
E-type seals located between all mating flanges prevent
air leakage, mating flanges are secured by Vee-band
coupling clamps.

Starter Air Control Valve
Purpose
The starter air control valve is designed to control the
admittance of air to the starter motor.
The valve is commanded from the flight deck via the EEC.
Location
The starter air control valve is located on the right hand
side of the engine fan casing (intermediate module).
Description
The starter air control valve consists of the following;
• Butterfly valve for airflow control.
• Pneumatically operated.
• Microswitch position indication for valve positional
status.
• Air filter to prevent valve operating mechanisms from
contamination.
• Failsafe position of the valve is closed.
• Provision of a manual override for abnormal start
attempts.
The starter air control valve is a pneumatically operated,
electrically controlled shut-off valve. The valve is
positioned on the lower right hand side of the engine fan
casing.
The starter air valve controls the airflow from the air
ducting 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.
Manual operation
The starter air valve can be opened/closed manually using
a 0.375 in square drive.
Access is through a panel in the right hand side fan cowl
door.
A valve position indicator is provided on the valve body.
A micro switch provides valve position feed back
information to the EEC.

Starter Air Valve
Operation
The starter air valve operation for opening and closing is
as follows;
Engine shutdown
With no pneumatic air supply available the valve is spring
loaded to the closed position.
Valve opening
Air upstream of the butterfly valve is filtered and routed
through an orifice in the solenoid valve.
Air upstream of the orifice is also admitted to the smaller
piston of the double acting actuator.
When the solenoid is energised the ball valve opens to
admit air to the larger piston whilst simultaneously closing
the vent port.
The air acting on the larger piston overcomes the
combined force of upstream air pressure acting on the
smaller piston and the actuator spring.
Movement of the actuator is translated through the linkage
to rotate the butterfly valve towards the open position.
Valve closing
When the solenoid is de-energised, at approximately 6000
rpm (43%) N2, the ball valve closes and air acting on the
larger piston is vented to atmosphere through the vent.
Air pressure and actuator spring pressure acting on the
smaller piston then closes the butterfly valve.
Any loss of air pressure will cause the butterfly valve to
close under the action of the actuator spring.

Pneumatic Starter Motor
Purpose
The purpose of the pneumatic starter motor is to provide
an initial rotational input to the high speed external
gearbox in order to assist the engine to achieve a stable
minimum power condition (low idle).
Location
The starter motor is located on the front face of the high
speed external gearbox.
Description
The starter motor consists of the following;
• Oil filler/level plug.
• Drain plug with a built in magnetic chip detector.
• QAD devices to allow for ease of maintenance.
The starter motor gears and bearings are lubricated by an
integral lubrication system.
A quick attach/disconnect adapter (QAD) attaches the
starter motor to the external gearbox. A quick detach Vee
clamp connects the starter motor to the adapter.
Note;
There are two standards of starter motor available for the
V2500 powerplant. The current being the synchronous
clutch engagement unit.
The synchronous clutch allows for smoother crash
engagements thus reducing the wear and damage caused
by such operations.

Starter Motor
Operation
The starter is a pneumatically driven turbine unit that
accelerates the HP rotor to the required speed for engine
starting.
The starter comprises of the following;
• A single stage turbine.
• A reduction gear train.
• A clutch and an output drive shaft.
These are 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 HP 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.
The engine HP turbine compressor assembly, the external
gearbox and starter output drive shaft continue to rotate.
When the starter output drive shaft rotational speed
increases above a predetermined rpm, Centrifugal force
overcomes the tension of the clutch leaf springs, allowing
the pawls to be pulled clear of the gear hub ratchet teeth
to disengage the output drive shaft from the turbine.

Engine Ignition System
Purpose
The ignition system is designed to provide the means of
igniting the air/fuel mix in the combustor.
Location
The ignition system units are located in the following
positions;
• The relay box is located on the right hand side of the
engine fan case.
• The high energy ignition units (HEIUs) are located on
the right hand side of the core engine. Mounted on the
HPC casing.
• The igniter plugs are located on the combustion diffuser
casing at fuel spray nozzle positions no. 7 and 8.
Description
Two independent ignition systems are provided.
The system is made up of the following units;
• Ignition relay box.
• Two ignition exciter units.
• Two igniter plugs.
• Two air cooled HT ignition connector leads.
The ignition system can operate in various modes. These
modes are as follows;
Dual igniter select;
• All in flight starts.
• Manual start attempts.
• Continuous ignition.
Single alternate igniter select;
• Auto starts.
Continuous ignition select;
• Engine anti ice.
• Take off.
• Approach.
• Landing.
• EEC failure.
Continuous ignition may also be selected manually.
The ignition exciters provide approximately 22.26 Kv and
the igniter discharge rate is 1.5/2.5 sparks per second.
Test
Operation of the ignition system can be checked on the
ground, with the engine shut down, through the
maintenance menu mode of the CFDS.

Ignition Relay Box
Purpose
Used for connection and the isolation of the high energy
ignition units.
Location
The relay box is located on the right hand side of the
engine fan casing.
Description
The ignition system 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.
Note;
The same relay box also houses the relay that controls the
115V AC supplies for P2/T2 probe heating.

Engine Starting Electrical Control
Engine Interface Unit (EIU)
Purpose
The EIU is an interface concentrator between the aircraft
and FADEC system.
Location
The EIU is located in the avionics bay.
Description
There are two engine interface units (EIU’s), one for each
engine. The EIU is an interface concentrator between the
aircraft and FADEC system.
The EIU main functions are;
• To concentrate data from the flight deck panels.
• To ensure the segregation of the two engines.
• To provide the EEC with electrical power supply.
• To give the necessary logic and information between
the engine and the aircraft systems.
• Receives discrete electrical signals from the cockpit.
Digitises these signals and transmits them to the EEC.
Also sends discrete signals to close air conditioning pack
flow valves and increase the airflow from the APU if
required.
Electronic Engine Control (EEC)
Purpose
To provide electronic signals for FADEC system unit
control.
Location
The EEC is located on the engine fan casing right hand
side.
Description
The EEC is the heart of the FADEC system and has
control of the FADEC system components and constantly
monitors their performance.
The EEC will make adjustments where necessary to
optimise the operation of the engine.
During the starting of the engine the EEC generates the
pneumatic starter valve opening/closing signal in respect
of control switch selection (rotary selector, master lever,
‘MAN START’ push button switch) and N2 speed signal.
The EEC will send any warning or caution message to the
flight warning computer (FWC).
The FWC will send this to the display management
computer (DMC) for indication on the ECAM upper or
lower CRT.

Engine Operation
Purpose
The purpose of engine operation on the ground is to
validate non FADEC detected faults (mechanical failures)
and to prove the integrity of an LRU or system after
maintenance has been carried out.
Flight deck
The flight deck is where the engines are operated from.
There are two panels utilised for starting the engine. These
panels are;
• The auto mode select panel (auto starts).
• The alternate mode select panel (manual starts).
Auto mode select panel
The auto mode select panel has the following;
• Two main engine master switches.
• Rotary switch.
The engine master switches have two positions;
• Cut off.
• On.
These switches activate the start air valve and the FMU,
via the EIU and EEC, when in the auto mode.
These switches activate the FMU, via the EIU and EEC,
when in the alternate (manual) mode.
The rotary switch has three positions;
• Crank.
• Mode norm.
• Ign/start.
The crank position allows operation of the starter motor
only.
The mode norm sets the ignition to auto function, for
example when anti icing is selected the ignition comes on
to a continuous operation.
The ign/start allows the engine to be started. This switch
must be in the ign/start position before selecting the
master levers to the on position.
Alternate mode (manual start pamel).
The alternate mode select panel allows the engine to be
started in the non auto function or manual mode.
Note:
The EEC software has a fuel flow reduction capability
upon the detection of a stall. This is known as fuel
depulse.
The depulse logic is designed to assist the engine in
recovery from a stall during starting.

Safety Zones
During run up operations, extreme care should be exercised
when operating the engines
Refer to the diagram below, which illustrates the inlet suction
hazard areas for the conditions at idle and take-off thrust.
.

Safety Zones - Jet Wake Hazard Areas
During run up operations, extreme care should be exercised
when operating the engines.
Refer to the diagram below, which illustrates the jet wake
hazard areas for the conditions at idle and take-off thrust.
Noise Danger Areas
All persons working near the engine while it operates must
wear ear protection.
Loud noise from the engine can cause temporary or
permanent damage to the ears.

SECTION 14
THRUST REVERSER SYSTEM

IAE V2500 General Familiarisation Thrust Reverser System
Thrust Reverser System Introduction
Purpose
The thrust reverser is designed to assist the aircraft in
decelerating quickly and safely upon landing.
It also assists deceleration during an aborted takeoff.
Location
The thrust reverser is an integral part of the C duct
assembly.
The C duct assembly is mounted to the aircraft strut by
four hinged brackets located at the top of the C ducts.
They are held in the closed position by six latch locks
located at the bottom of the C ducts.
Description
The reverser is a translating sleeve type system. It directs
the fan air rearwards for normal forward thrust or forwards
for thrust reverse.
When the thrust reverser system is in the stow position the
fan air exhausts at the common nozzle. This produces
forward thrust.
When the thrust reverser is deployed four linear motion
actuators cause the translating sleeves to move
rearwards.
This moves the blocker doors from an axial to a radial
position in the C duct fan exhaust area.
The blocker doors forces the fan air through the cascades
in a forward direction. The cascades are exposed
whenever the thrust reverser is deployed.
Reverse thrust is selected from the flight deck by the gated
reverse thrust levers. The EEC has control over the
operation of the thrust reverse system.
All signals to and from the thrust reverser are through the
EIU and EEC.
Thrust reverser system features
Electronic control.
Hydraulic actuation system.
Positional information feedback.
Actuator lock position sensors and feedback.
Electronic safety locks.
Automatic re-stow system.
Manual deployment and stow capability for maintenance.
Manual lockout to allow aircraft to be despatched with an
inoperative thrust reverser.

DETV250262
V2500 THRUST REVERSER SYSTEMV2500 THRUST REVERSER SYSTEM

Page 14-3Revision 1
• EIU inhibit relays for un-commanded in-flight
deployment.
Thrust Reverser Assembly
When the translating sleeves are in the forward thrust
position the path of the fan bypass air is in the normal
forward thrust.
Rearward movement of the translating sleeves unveils the
cascade deflectors and moves the blocker flaps from an
axial to a radial position.
This blocks the fan stream airflow and forces the fan efflux
through the cascade deflectors.
Methods of deployment
The thrust reverser can be deployed in one of two
methods;
• Using the engine/aircraft hydraulic system. Moving the
thrust reverse select levers that are mounted on the
main forward levers does selection.
• Manual input by two hand turning points for
maintenance purposes.
Safety features
The thrust reverse system operation is controlled by the
engine electronic control (EEC). The following are EEC
controlled functions of the thrust reverse system.
The thrust reverse system incorporates a double lock
safety system to protect against inadvertent deployment.
They are;
• Landing gear control unit (LGCU).
• Lock sensors on the locking actuators.
• Shut off valve that is signalled to operate from the
SEC.
• The thrust reverser also has a system that will return
the engine thrust to idle should the thrust reverse
system inadvertently deploy.
• Auto re-stow is a system that is designed to stow the
thrust reverser when an un-commanded deployment is
detected.
Indications
ECAM indications for fault annunciation of the thrust
reverser system status are done by use of proximity
sensors, relay select status, hydraulic system pressures
and LVDT feedback signals.
The locking actuator sensors detect unlocked conditions
and the LVDT detects transient and deployed conditions.
The signals are relayed from the EEC to the EIU and then
to the ECAM screens.

Page 14-5Revision 1
In the flight deck this unlocked condition is identified as an
amber coloured REV caption on the EPR indicating gauge.Operation
Thrust reverse is selected from the flight deck by pulling up
on the thrust reverse select levers. The select levers are
mounted on the front side of the main thrust levers. The
thrust levers have a gated feature that allows thrust
operation by the throttles in one direction only.
The EEC has control of the thrust reverse system
operation for deploy and stow.
The EIU inhibit relay controls the DCV power signal for the
control solenoid from the EEC to the DCV.
Deploy
Pulling up on the thrust reverse lever in the flight deck will
send a signal for thrust reverse select to the EEC. This will
also put the main throttles in the reverse thrust quadrant.
The EEC will look for the following conditions before thrust
reverse will be allowed;
• The EEC will check that the aircraft is on the ground by
checking the LGCU signal of the aircraft computers.
• The EEC will check that the engine is running by means
of a N2 signal.
• The EEC cannot deploy the thrust reverser until the EIU
inhibit relay is active.
• SEC control signal for the shut off valve.
The hydraulic isolation valve solenoid and the directional
control valve solenoid will both be energised for a deploy
condition.
This will admit high pressure hydraulic fluid to the stow and
deploy sides of the thrust reverse system.
The lower locking actuators will unlock and the EEC will
see a signal from the proximity sensor of reverser system
unlocked.
When the translating sleeves have moved to 78% of the
full deploy position the amber REV indication will change
to a green REV indication.
When green REV is indicated the full reverse thrust power
is available to the flight crew.
Stow
To stow the thrust reverse system the flight crew will return
the throttles to the idle detent position and select levers to
the down position. This will put the throttles back to the
forward thrust quadrant.
The hydraulic isolation valve solenoid is energised and the
directional control valve solenoid is de-energised for a
stow condition. This leaves high pressure hydraulic fluid
present on the stow side of the system.
As the translating sleeves move from deploy back to stow
the flight deck indication will change from green to amber
on the REV indication.
When the thrust reverser has reached the fully stowed
position the amber REV will go and the EPR gauge will
return to normal indication.
This will indicate that the thrust reverser is fully stowed and
locked.

Revision 1 Page 14.6

Controls and Indications
Thrust reverse is selected from the flight deck by use of
latching selector levers that are mounted on the main
throttle control levers.
Controls
Pulling the levers upwards will initiate the sequence of
events that will deploy the reverser system. The EEC, in
conjunction with the EIU, controls the deployment and
stowing of the reverser system.
The levers move into the thrust reverse quadrant therefor
while in this position throttle movement is only possible in
the thrust reverse mode.
Movement towards the maximum throttle stop for thrust
reverse is possible but the engine will only accelerate
when the EEC has feedback of the translating sleeve
status. The translating sleeve must be beyond 78% of the
fully deployed position.
Stowing the thrust reverser requires the latching select
levers to be pushed down and the main throttles will revert
back to normal forward thrust. This will also stow the
reverser system.
Indications
The thrust reverser system indications appear on the
ECAM CRTs. The EPR indication is used to display the
status of the thrust reverser.
Normal indication Thrust reverser stowed and
locked.
REV in colour amber Thrust reverser unlocked
and in transit.
REV in colour green Thrust reverser deployed.
Thrust reverser indications of non-normal conditions will be
indicated to the ECAM screens in the form of a message.
The following are the associated ECAM messages that
appear to the ECAM screens;
• Reverse Unlocked.
• Reverser Fault.
• Rev pressurised.
• Rev Switch Fault.
By entering the CFDS screens the faults can be
interpreted to pinpoint the location.
Revision 1 Page 14.7

Hydraulic System
The hydraulic system provides the force required to move
the translating sleeves for both deploy and stow
conditions.
The hydraulic system comprises of the following;
• Linear motion actuators.
• Flex shaft.
• Hydraulic control unit comprising of a HIV and DCV.
Linear Motion Actuators
There are four linear motion actuators per C duct set.
The two upper actuators are non-locking and incorporate
LVDT’s for feedback to the EEC.
The two lower actuators are locking they incorporate
proximity sensors to give indication to the EEC of lock and
unlock conditions.
Flex shaft
The four linear motion actuators are kept in
synchronisation movement by flexible shafts that have a
high torsion resistance.
Hydraulic Control Unit
The hydraulic control unit comprises of the following items;
• Hydraulic isolation valve (HIV).
• Directional control valve (DCV).
The HIV controls the presence of high pressure hydraulic
fluid in the thrust reverser system. The energising of a
control solenoid valve controls this valve. For stow and
deploy conditions this valve must be energised.
The DCV controls the flow direction of the high pressure
hydraulic fluid once it is in the thrust reverser system.
• The DCV will direct the high pressure hydraulic fluid to
the stow and deploy sides of the system for deploy
conditions.
• The DCV will direct the high pressure hydraulic fluid to
the stow side of the system for stow conditions.

Hydraulic Control Unit
Purpose
The hydraulic control unit (HCU) is designed to control the
safe passage of the high pressure hydraulic fluid to the
thrust reverser system.
Location
The HCU is located between the top of the engine fan
case and the aircraft strut. Access is gained by opening
the left hand side fan cowl door.
Description
The HCU is a self contained LRU designed to control the
flow of high pressure hydraulic fluid. The EEC and EIU has
control over the HCU control solenoids.
When the EEC detects a demand for thrust reverse
operation both EEC and EIU will signal the HCU control
solenoids.
The HCU has the following features;
• A hydraulic isolation valve (HIV) which controls the flow
of high pressure hydraulic fluid into the thrust reverser
system. The HIV function is controlled by a control
solenoid valve.
• A directional control valve (DCV) which controls the
direction of flow of the high pressure hydraulic fluid to
either the deploy or stow sides of the system. The DCV
function is controlled by a control solenoid valve.
• A pressure switch to feedback system pressure status
to the EEC.
• A filter with a clog indicator that gives a visual
indication of the filter condition when it becomes
contaminated.
• A bleed valve.
• Provision for locking out the valve operation for
maintenance and flight.

Hydraulic Control Unit Operation
Deploy
When the thrust reverser select levers are moved to the up
position the EEC detects that the thrust reverser system is
required.
• The EEC has selection control over the HIV.
• The EIU has selection control over the DCV.
The HIV control solenoid will be energised. This moves a
lock-pin away from the pilot valve chamber orifice. High
pressure hydraulic fluid enters the left hand side of the
chamber and forces the pilot valve to move to the right.
The pilot valve moves against a spring that exerts a
pressure on the right hand side of the pilot valve.
The pilot valve recess moves in line with the hydraulic fluid
supply tube. This admits high pressure fluid into the thrust
reverse system and initially to the stow side of the system.
The pressure switch moves to the high pressure indicating
position.
The DCV control solenoid will be energised. This moves a
lockpin away from the pilot valve chamber orifice. High
pressure hydraulic fluid enters the left hand side of the
chamber and forces the pilot valve to move to the right.
The pilot valve moves against a spring that exerts a
pressure on the right hand side of the pilot valve.
The pilot valve recess moves in line with the hydraulic fluid
supply tube. This admits high pressure fluid into the thrust
reverse system deploy side of the system.
There is now pressure present in both the stow and deploy
sides of the system.
The restrictors in the deploy supply tube to the DCV delay
the pressure build up to the deploy side so the pressure
present on the stow side can push the locking actuators
towards the stow. This releases the pressure acting on the
tine locking mechanism.
Stow
To stow the thrust reverser the select levers are moved to
the down position.
• The EEC has selection control over the HIV.
• The EIU has selection control over the DCV.
The DCV control solenoid is de-energised. The pilot valve
moves to the right due to spring pressure alone. This
leaves high pressure hydraulic fluid present in the stow
side of the system.
When the thrust reverser system has fully stowed the EEC
will sense this by a feedback signal coming from the
unlock sensors. The EEC will then de-energise the HIV
control solenoid.
The control solenoid pilot valve will move to the right due
to spring pressure alone and the high pressure hydraulic
fluid is cut off from the thrust reverser system.
HIV Deactivation
The deactivating lever prevents the HIV pilot valve from
moving. This prevents high pressure hydraulic fluid from
entering the thrust reverser system.

Lower Locking Actuators
Purpose
The hydraulic actuators in general control the movement of
the translating sleeves.
The lower locking actuators incorporate a locking
mechanism which gives;
• A feedback signal to the EEC of actuator locked or
unlocked.
• A means of preventing the translating sleeves from un-
commanded movement.
Location
The lower locking actuators are located in the thrust
reverser C duct units at the lower positions.
Description
The actuators in general control the movement of the
translating sleeves in a linear motion.
The lower actuators on either thrust reverser C duct have
a locking mechanism incorporated in the design. The
locking mechanism adds to the safety of the system.
The lower locking actuators incorporate the following
features;
• Hydraulically operated linear motion actuators.
• A locking mechanism called a tine lock. This can only
be unlocked when hydraulic pressure is present in the
system.
• A linear motion actuator with pressure surfaces on
either side of the pressure plate.
• An acme screw thread that rotates a worm gear.
• A worm wheel that rotates the flex shafts.
• A manual unlocking feature for maintenance purposes.

Lower Locking Actuators Operation
Deploy
The EEC will energise the HIV and DCV control solenoids
which allows high pressure hydraulic fluid to be present in
the thrust reverser system. This is on the stow and deploy
sides.
The restrictor in the pressure feed tube to the deploy side
delays the deploy pressure build up enough to allow the
stow pressure to initially push the actuator piston towards
the stow direction. This releases the lock pressure on the
tine locking mechanism.
The unlock sleeve is then pushed towards the right of the
tine lock. With the locking sleeve clear of the tine lock the
tine lock flexible spring type fingers are free to flex.
When the locking sleeve moves a lever assembly also
moves. The lever assembly is linked to the external area of
the actuator. The lever has a target attached to it.
• When the actuator is at stow the target is in line with
the proximity sensor.
• When the actuator is at deploy the target is away from
the proximity sensor.
The EEC detects these conditions.
High pressure fluid being present in both sides of the
system forces the actuator to move towards the deploy
direction. This bias of movement exists because the
surface area of the deploy side of the pressure plate is
greater that the surface area of the stow side.
Stow
To stow the thrust reverser the high pressure fluid present
on the deploy side of the system must be reduced to tank
pressure and ported back to the hydraulic reservoir of the
aircraft.
The EEC will de-energise the DCV control solenoid and
this will leave high pressure hydraulic fluid present on the
stow side of the system.
The actuator will now move in the stow direction. The head
end of the actuator engages into the tine lock. The locking
sleeve will move into position to immobilise the tine lock by
spring pressure.
As the locking sleeve moves to the lock position the target
on the unlock indicator moves in line with the proximity
sensor. The EEC detects this and sees that the thrust
reverser system is stowed.

Upper Non Locking Actuators
Description
The hydraulic actuators in general control the movement of
the translating sleeves.
The upper non locking actuators incorporate a linear
variable displacement transducer (LVDT) that feeds back
translating sleeve status of translation and deployment.
The actuators operate in the same manner as the locking
actuators.
• For deployment the high pressure hydraulic fluid is
present on both sides of the system.
• For stowing the high pressure hydraulic fluid is present
on the stow side of the system only.
The LVDT monitors the movement of the translating
sleeves and feeds back the signals to the EEC. This is
done to tell the EEC that the translating sleeves are in
transit and when 78% of travel towards the deploy has
been achieved.

Shut Off Valve
Purpose
To give additional safety to un-commanded deployments
of the thrust reverser. This is known as the third lock of
safety.
The shut off valve has a filter assembly installed in line
with the valve.
Location
The shut off valve is located in the aircraft strut at the front.
It is behind the HCU.
Description
The shut off valve has the following features;
Control solenoid.
Two position valve assembly.
Operation
The shut off valve operation for opening and closing relies
upon the signals from the spoilers and elevators computer
(SEC).
Pulling up the thrust reverse select levers will signal the
SEC to open the shut off valve.
The SEC sends a signal to the shut off valve relay. The
relay energises the solenoid and opens the shut off valve.
The high pressure hydraulic fluid now flows towards the
HCU.
To close the shut off valve the selection of thrust reverse
levers must be in the forward thrust position.
Manual Bypass Non Return valve
Purpose
For normal thrust reverser operation provides a one way
directional flow for the hydraulic fluid.
For maintenance purposes it allows the flow of hydraulic
fluid easier when manual deployment and stow of the
thrust reverser is required.
Location
The manual bypass valve is located in the aircraft strut just
behind the shut off valve.
Description
The manual bypass non return valve allows the flow of
hydraulic fluid in one direction only.
There is a very hard spring loaded valve inside that is
difficult to unseat when carrying out manual operations of
the thrust reverser system.
The bypass handle allows the valve to become unseated
for maintenance operations only.
Access to the bypass valve is through an access panel
located on the left hand side of the aircraft strut.

Thrust Reverser Flex (Synchronisation) Shaft
Purpose
The flex shaft purpose is to maintain synchronous
movement of the actuators. This prevents any one
actuator from moving faster than the others.
Location
The flex shaft is located within the deploy tube system.
Description
The flex shaft system comprises of the following;
T piece housing assembly. This allows the distribution of
high pressure hydraulic fluid to both sides of the thrust
reverser system.
Two flexible tubes. This allows the crossover shaft to link
the reverser halves together while allowing the C ducts to
be opened.
Two rigid tubes. These are found between the upper and
lower actuators. They carry hydraulic fluid to the deploy
side of the system.
Three flexible shafts. These link all the actuators together.
Note;
The two deploy tubes have a telescopic coupling at one
end to permit simple removal and installation without
disturbing the actuators.

Thrust Reverser Deflector Boxes (Cascades)
Purpose
The cascades are designed to direct the fan air to provide
the reverse thrust for the engine.
Location
The cascades are located between the inner and outer
translating sleeve sleeves. They are mounted on the fixed
section of the C ducts.
Description
There are 16 cascades fitted to the thrust reverser system.
The cascades are designed to direct the fan air forwards
thus providing for the function of the thrust reverse system.
They are designed to direct the fan air away from the
ground thus reducing the risk of debris from being blown
up and ingested into the engine.
They are designed to direct the fan air away from the
airframe thus not inducing any unnecessary stress upon
the airframe itself.

Operation
Thrust reverse is selected from the flight deck by pulling up
on the thrust reverse select levers.
• The EEC has main control of the thrust reverse system
operation for deploy and stow.
• The EIU has a inhibit relay that controls the power
supply signal from the EEC to the DCV.
• The flight crew have control of reverse thrust power
selection.
Deploy
Pulling up on the thrust reverse select lever in the flight
deck will send a signal for thrust reverse select to the EEC
and EIU. This will also put the main throttles in the reverse
thrust quadrant.
The EEC will look for the following conditions before thrust
reverse will be allowed;
• The EEC will check that the aircraft is on the ground by
checking the LGCU signal of the aircraft computers.
• The EEC will check that the engine is running by means
of a N2 signal.
• The EIU will look for the signal from the throttle control
unit for energising of the inhibit relay.
• The SEC signal for opening the shut off valve.
The hydraulic isolation valve solenoid and the directional
control valve solenoid will both be energised for a deploy
condition. This will admit high pressure hydraulic fluid to
the stow and deploy sides of the thrust reverse system.
The lower locking actuators will unlock and the EEC sees
a signal from the proximity sensor of reverser system
unlocked.
In the flight deck this unlocked condition is identified as an
amber coloured REV caption on the EPR indicating gauge.
When the translating sleeves have moved to 78% of the
full deploy position the amber REV indication will change
to a green REV indication.
When green REV is indicated the full reverse thrust power
is available to the flight crew.
Stow
To stow the thrust reverse system the flight crew will return
the throttles to the idle detent position and select levers to
the down position. This will put the throttles back to the
forward thrust quadrant.
The DCV solenoid will be de-energised as commanded by
the EEC via the EIU inhibit relay. This will leave high
pressure hydraulic fluid present on the stow side of the
reverser system only.
As the translating sleeves move from deploy back to stow
the flight deck indication will change from green to amber
on the REV indication.
When the thrust reverser has reached the fully stowed
position the amber REV will go and the EPR gauge will
return to normal indication.
This will indicate that the thrust reverser is fully stowed and
locked.

IAE V2500 General Familiarisation Troubleshooting
IAE V2500 Troubleshooting
Introduction
In order to locate the source of an engine problem both
quickly and efficiently, it is essential that the aircraft
maintenance engineer is aware of the fundamental
approach to troubleshooting required on the Airbus
A319/320/321.
Having acquired the knowledge of various engine systems
functionality and operation during this course, we are now
in a position to take the course the natural step forward
and discuss the all-important methodology of isolating and
identifying the source of a problem.
An important tool available to the engineer is the
A319/A320/A321 ‘Computer Assisted Aircraft Trouble
Shooting’ (CAATS) CD-ROM. This valuable aid provides
the user with an enormous amount of detail and
information.
This manual is revised and issued every three months.
Upon receipt and installation of the up-dated version the
previous version is automatically overwritten and the
previous disc is now no longer valid or useable.
The CAATS CD-ROM is password protected for each
airline, as it is tailored specifically for each operator’s
requirements. Therefore it is essential that the user always
access the procedures for their own particular airline.
After inserting the correct password the user is presented
with the screen shot below.
The most common and straightforward menu selection is
‘Trouble Shooting Procedures’. If the user were to select
‘Trouble Shooting Manual’, this would require the user to
insert a known trouble shooting task reference number in
order to progress.
Unless a procedural task has been already identified
during previous investigation activity it will not be practical
to use, as its selection would be dependent knowing which
system has the fault.
Below is a screen shot showing the opening menu options
and for the purpose of these training notes we are using
‘British Midland’ as the ‘Log on Airline’.

Introduction continued:
Upon selecting ‘Trouble Shooting Procedures’ the user is then
presented with the screen shot (fig. 1) shown below.
From this menu it is possible to enter into the trouble shooting
process with information derived from a variety of sources:
1. ECAM WARNINGS
These are the messages that appeared on the Upper ECAM
Screen during operation and show the symptom or system,
which has been degraded by a fault.
These are generally of a ‘Class 1’ level which would prevent
the aircraft from being dispatched unless the problem and
source of the message had been rectified. Check Minimum
Equipment List (MEL).
2. ECAM STATUS (Inoperative Systems and
Maintenance Status).
The presence of an ECAM Status Message ‘STS’ is
automatically displayed on the Upper ECAM Screen during
Flight phase 1 (Electrical Power ‘on’ before first engine start)
and Flight Phase 10 (When the second engine has been shut
down after the flight). It is used to highlight a problem or
degradation in the built in redundancy facility of the FADEC
System. This feature prevents un-wanted distractions of
system degradation being shown to the pilot during the flight.
A fault of this nature is dispatchable and the fault can be left
un-rectified for up to ten days.
Check Minimum Equipment List (MEL) the Status Page can
then be selected by pressing the STS button on the Systems
Page Select Panel. This will then provide information under
the ‘Maintenance’ heading regarding the failure, for example
ENG 1(2) FADEC or ENG 1 (2) EIU.
Fig.1 3. LOCAL WARNINGS (Panel Lights and Standby
Indicators).
Lists the entry into the Trouble Shooting procedure for given
indicated engine related faults. (This has limited use).
Information is restricted to ‘Start Air Valves’, ‘Fuel Start
Valves’ and Anti-Ice Valve problems.

STATUS
INOP SYSTEM
G+B HYD
CAT 3
G RSVR
L+R AIL
SPLR 1+3+5
L ELEV
AP 1+2
ENG 1 REV
NORM BREAK
NW STEER
MAINTENANCE
APU
AIR COND
ENG 1 FADEC
MAX SPD…………………….250/.85
APPR PROC DUAL HYD LO PR
-IF BLUE OVHT OUT:
-BLUE ELEC PUMP….. ON
-LG………………………..GRVTY EXTN
-LDG SPD INCRMT………10 KT
SLATS SLOW
CAT 1 ONLY
CANCELLED CAUTION
NAV IR 2 FAULT
PSI 35 Class 2 Failure of the
Engine Number One’s
FADEC System
TAT -5 C
SAT -30 C
G.W. 60300KG23H56
STATUS PAGE (LOWER ECAM)

4. FLAGS and ADVISORIES (On ECAM and EFIS
System Pages)
Selecting this provides the user with the screen shown
Below, see (fig. 2).
By selecting the appropriate system, the user will be
presented a complete listing of Flags and Advisories
available, related to problems with that particular system.
In this example the Engine System. See (fig 3) below.
(fig. 2) (fig.3)

5. CREW and MAINTENANCE OBSERVATIONS
By selecting this option, the user can relate to what conditions
they have seen during the engines operation and link and
match the symptoms that to the list provides. Fig 4 below
illustrates that if the user types in the main heading for that
system, a complete list of all possible observations of faults
are produced.
Fig. 5 illustrates the complete listing of, in this example (ATA
73) referenced observations
Fig 5Fig 4

6. CFDS FAULT MESSAGE
Centralised Fault Display System (CFDS) This menu selection
is one of the most common methods of entering into the
Trouble Shooting process. By interpreting the information
provided on the Post Flight Report (PFR) and completing the
necessary data field boxes, the user can quickly locate the
appropriate Trouble Shooting task for this particular systems
problem.
In this example we have a problem with the Number 2
Engine’s Fuel Heat Management System. This message
appeared on the upper ECAM as an ‘ECAM WARNING’ this is
a Class 1 failure and is not dispatchable.
The CFDS Fault Message is the text contained under the
heading ‘FAILURE MESSAGES’ on the Post Flight Report.
Again, in this example the Failure Message that is linked to
the Upper ECAM ‘ENG 2 FUEL HEAT SYS’ is:
‘FUL DIV RET VLV/HC/EEC2’
This text along with the ATA reference number:
‘73-13-42’
and the Source:
‘EIU2FAD’
is copied into the text boxes as shown on (fig. 6) opposite.
Completing the Class of Failure data can make further
refinement of identifying the task. In this case we had an
upper ECAM warning message, so in this example we can
identify it as a ‘Class 1’ fault.
Fig. 6

7. None
In selecting the menu option of ‘None’ the user is presented
with the screen shown below (fig. 7)
This requires the insertion of known information in order to
refine the search. If in the example shown the user simply
types the first two ‘ATA’ digits for engine related problems,
which are ‘77’ and then selects ‘Enter’. Then the complete list
of failures and associated warnings is produced (fig. 8)
opposite.
Fig 7 Fig 8

Beginning of PFR recording, first engine start +
3 minutes =18:27
End of PFR Recording. 80 knots + 30 seconds
= 21:17
GMT (Greenwich Mean Time) =Time when the
cockpit warning was displayed.
PH = Flight Phase.
ATA = Air Transport Association
Note time of ECAM Warning and CFDS Failure
Message is the same. (Although there can be
up to two minutes difference)
Source = System detecting the fault
POST FLIGHT REPORT

E V2500 Troubleshooting
Centralised Fault Display System (CFDS)
The purpose of the CFDS is to give the maintenance
engineers a central maintenance aid to intervene at system or
sub-system level from a Multipurpose Centralised Display Unit
(MCDU) located on the flight deck.
The MCDU allows the engineer to;
• Interrogate a variety of systems using Built in Test
Equipment (BITES) for maintenance information.
• To initiate system return to service tests.
The detection of the failures, processing and formatting of the
failure messages to be displayed is carried out in each
systems individual systems BITE.
There are two MCDU’s and either the Captain’s or First
Officer’s MCDU can be used.

Revision 1
Alpha Keys
Line Select Keys
Numeric Keys
Brightness Adjust
Annunciators
Function and Mode Keys
Line Select Keys
Page 15-14
MULTIPURPOSE CENTRALISED DISPLAY UNIT (MCDU)

Failure Classification and Master Minimum Equipment
List (MMEL)
The MMEL cannot be used as a Minimum Equipment List
(MEL) due to the fact that it is not related to operational
requirements, specific operations or airlines particular
definitions. The MMEL can be used as a basis for
particular operators own MEL. The MEL should be used to
establish dispatchability for a particular operation.
The MEL does not include those items that are obviously
required for aircraft safety, such as wings, engines etc.
The MEL does not include those items that do not affect
the airworthiness of the aircraft, such as galley equipment,
entertainment system etc.
Note;
All items, which are related to the airworthiness of the
aircraft and not included in the list, are automatically
required to be operational for each flight.
MEL Preamble
The MEL is intended to permit operation with inoperative
items of equipment for a period of time, until repairs can
be accomplished at the earliest opportunity. In order to
maintain
acceptable levels of safety and reliability the MEL
establishes limitations on the duration of and conditions for
operation with inoperative equipment.
When an item of equipment is discovered to be
inoperative, it is reported by making an entry into the
Aircraft Maintenance Record/Logbook as prescribed by the
Federal Aviation Regulations (FAR).
The item is then either repaired or may be deferred as per
the MEL or other approved means acceptable to the
Administrator prior to further operation.
MEL conditions and limitations do not relieve the operator
from determining that the aircraft is in a condition for safe
operation with items of equipment inoperative.
Federal Aviation Authority (FAA) Repair Intervals
All users of an approved MEL, must effect repairs of
inoperative systems or components, deferred in
accordance with the MEL, at or prior to the repair times
established by the following letter designators:
• Category A: To be repaired within the time interval
specified in the remarks column of the operator’s
approved MEL.
• Category B: To be repaired within three (3) consecutive
calendar days (72 hours), excluding the day the
malfunction occurred.
• Category C: To be repaired within ten (10) consecutive
calendar days), excluding the day the malfunction
occurred.
• Category D: To be repaired within one hundred and
twenty (120) days), excluding the day the malfunction
occurred.

Printer
MCDU’s
LOCATION OF MCDU’S AND PRINTER

Failure Classifications
There are three (3) levels of ‘Failure Classifications’ and
these are signified by the method of notification of their
existence to the Flight Crew or to the Maintenance
Engineer during ground operation and testing.
Class 1
Failures are indicated, by means of the upper ECAM
display or local warnings. Procedures to be followed by the
operator to help to ameliorate the problem may also be
displayed.
Class 2
The operator is informed of a Class 2 failure on the ECAM
STATUS page which only shows the system affected by
the Class 2 failure. A white ‘STS’ symbol appears on the
upper ECAM.
Class 3
The operator is not informed of Class 3 failures. Class 3
failures are only accessible through the Centralised Fault
Display System (CFDS) via the MCDU in ‘menu mode’

CURRENT
GROUND
FAULTS
FADEC Self Test
Reverser Test
Ignitor Test
Start Valve Test
P2 T2 Heater Test
500 HOURS
Class 3 Faults
UNLIMITED
DESPACTH
DATE
TIME
ATA CHAPTER
CELL NUMBER (1 - 60)
Trouble Shooting Data
from Stored Faults
Faults stored
during the last leg
Faults stored
during previous
63 legs
FAULT ACRONYM
CELL NUMBER
FLIGHT PHASE
FLIGHT LEG
ENGINE PARAMETERS
WHEN FAULT RECORDED
'FLIGHT' OPERATION IS DEFINED AS
ENGINE AT IDLE (PLUS 3 MINUTES)
LAST LEG
REPORT
(Screen 7)
PREVIOUS LEG
REPORT
GROUND
DATA
(Cells 46 - 60)
FLIGHT
DATA
(Cells 1- 45)
(Screen 9 &10)
TROUBLESHOOTING
(Screen 8)
GROUND
SCANNING
SYSTEM TEST SCHEDULED
MAINT REPORT
CLASS 3
(Cells 61-69)
FADEC XX
MAIN MENU
(Screen 6)
CFDS
(Screen 2)
FADEC XX
(Screen 6)
ENG
(Screen
SYSTEM REPORT
TEST
MCDU
Main Menu
(Screen 1)
Clear Language Message
MCDU SCREEN ROUTEMAP

FAULT CLASS FLIGHT CREW ALERT DISPATCH
CONDITIONS
ACTION REQUIRED
Class 1 Visual and Audible Warning
(Upper ECAM)
NO GO
or
GO IF
or
GO
Refer to MEL for details
Class 2 Visual Indication
(‘STS’ appears on Upper
ECAM)
Specific Details
(Lower ECAM)
GO Fault must be recorded and
repaired as per MEL
SMR
Scheduled Maintenance
Report
No Indication
(CFDS must be interrogated
for details)
No Conditions Repair at next
'A' Check / 500 hours
Class 3 No Indication
(CFDS must be interrogated
for details)
No Conditions Time Unlimited Fault
(Should be repaired at
earliest convenient
opportunity)
FAULT CLASSIFICATION TABLE AND REQUIREMENTS

V2500 gf issue 01

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