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:
----------------------------------------------------------------------------
Revision 1 Page 1-1

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.
Revision 1 Page 1-2

Revision 1 Page 1-3

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.
Revision 1 Page 1-4

Revision 1 Page 1-5

THIS PAGE IS LEFT INTENTIONALLY BLANK
Revision 1 Page 1-6

Revision 1 Page 1-7

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.
Revision 1 Page 1-8

Revision 1 Page 1-9

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.
Revision 1 Page 1-10

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.
Revision 1 Page 2-1

Revision 1 Page 2-2

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.
Revision 1 Page 2-3

Revision 1 Page 2-4

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.
Revision 1 Page 2-5

Revision 1 Page 2-6

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.
Revision 1 Page 2-7

Revision 1 Page 2-8

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).
Revision 1 Page 2-9

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.
Revision 1 Page 3-1

Revision 1 Page 3-2

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.
Revision 1 Page 3-3

Revision 1 Page 3-4

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.
Revision 1 Page 3-5

Revision 1 Page 3-6

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.
Revision 1 Page 3-7

Revision 1 Page 3-8

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.
Revision 1 Page 3-9

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.
Revision 1 Page 4-1

Revision 1 Page 4-2

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
Revision 1 Page 4-3

Revision 1 Page 4.4

.
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
Revision 1 Page 4-5

Revision 1 Page 4.6

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.
Revision 1 Page 4-7

Revision 1 Page 4.8

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.
Revision 1 Page 4-9

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
Revision 1 Page 5-1

Revision 1 Page 5-2

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.
Revision 1 Page 5-3

Revision 1 Page 5.4

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.
Revision 1 Page 5-5

Revision 1 Page 5.6

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.
Revision 1 Page 5-7

Revision 1 Page 5-8

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.
Revision 1 Page 5-9

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.
Revision 1 Page 6-1

Revision 1 Page 6-2

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.
Revision 1 Page 6-3

Revision 1 Page 6-4

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.
Revision 1 Page 6-5

Revision 1 Page 6-6

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.
Revision 1 Page 6-7

Revision 1 Page 6-8

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.
Revision 1 Page 6-9

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.
Revision 1 Page 7-1

Revision 1 Page 7-2

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
Revision 1 Page 7-3

Revision 1 Page 7-4

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.
Revision 1 Page 7-5

Revision 1 Page 7-6

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.
Revision 1 Page 7-7

Revision 1 Page 7-8

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.
Revision 1 Page 7-9

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.

V2500 gf issue 01

Recommended

Recommended

More Related Content

What's hot

What's hot (20)

Similar to V2500 gf issue 01

Similar to V2500 gf issue 01 (20)

Recently uploaded

Recently uploaded (20)

V2500 gf issue 01