DESCRIPTION
AND OPERATION
OF A320 ENGINE
(IAE V2500)
MANUFACTURERS:
Pratt & Whitney of U.S.A,
Rolls Royce of UK,
Japanese Aero Engine,
Fiat Aviazone of Italy
Motor Turbine Union (MTU) of Germany.
BY
M PINAKINI-06951A2125
B ADITYA REDDY-06951A2101
CERTIFICATES
ACKNOWLEDGEMENT
We earnestly take the responsibility to acknowledge the following distinguished
personalities who graciously allowed us to carry out our project work successfully.
we express our gratitude to the esteemed organization Air India (National
Aviation Company of India Limited (I)) , Hyderabad for providing means of attaining
of our most cherished goals.
we are grateful and indebted to Mr. V.M.M. Rao, Engineer, Indian Airlines
Ltd, for his valuable support and guidance in executing this project work. It was his
qualified commitment that led to the completion of the project work.
we take this opportunity to express our acknowledgement and record our sincere
thanks to our respected parents, family members, friends, all faculty members and all
other well wishers whose casual and informal encouragement and motivations at
various stages resulted in the accomplishment of project.
ABSTRACT
The concept of Jet propulsion Engine has started in 150-BC by Hero through his
Aeolipile based on Newton’s laws. After various inventions and developments
numerous types of jet engines have been introduced in aviation for commercial, military
and space research applications. Now a days, engines are being designed which are
more reliable and efficient. Further developments are being made to travel faster,
higher, quite, and smoother.
Aeronautics is the one of the science of flight which involves method of
designing an airplane or other flying machine. There are four basic areas that
aeronautical engineers must understand in order to be able to design planes. Mainly
aerodynamics and propulsion. Aerodynamics is the study of how air flows around the
airplane. By studying the way air flows around the plane the engineers can define the
shapes of the plane. Propulsion is the study of how to design an engine that will provide
the thrust that is needed for a plane to take off and fly through the air. The engine
provides the power for the airplane. Our work involves with case study of turboprop
engine.
A turboprop engine includes an engine nacelle and at least one bleed air line on
the low-pressure compressor and at least one ejector formed by a cooling air duct and a
nozzle to create a cooling air flow within the engine nacelle during critical ground idle
operation (controlled or uncontrolled), and without undesirably increasing fuel
consumption or disturbing the work cycle of the engine. The ejector is arranged within
the engine nacelle in the forward part of the turboprop engine , with the cooling air duct
appertaining to the ejector connecting at least one air intake disposed on the periphery
of the engine nacelle with the interior of the engine nacelle , and with the at least one
nozzle being arranged in the cooling air duct .
INDEX
1.
POWERPLANT
2. ENGINE
3. ENGINE FUEL AND CONTROLLING
4. IGNITION
5. AIR
6. ENGINE CONTROL
7. ENGINE INDICATING
8. EXHAUST
9. OIL
10.
STARTING
INTRODUCTION
Most of the IAE V2500 engine's advanced features are the result of its
heritage. The V2500 traces its pedigree through partner company engines such as the
Rolls-Royce RB211 and the Pratt and Whitney PW4000.
One of the most noticeable features of the V2500 are its unique fan blades, a
good example of the advanced, proven technology contributed by the IAE partners.
The V2500 uses wide chord, shroudless, hollow blades designed and
developed by Rolls-Royce. These are manufactured by placing a 3D-machined piece of
honey-comb material between two sheets of pre-machined titanium.
At high
temperatures a diffusion bond is formed between these three piece of materials such that
the finished blade is effectively a single piece, hollow structure. This lightweight blade
is then extremely strong with a leading edge that is robust and can resist damage due to
foreign object impact. In addition, the wide chord nature of the blade centrifuges
runway debris and dust into the bypass duct reducing engine removals due to Foreign
Object Damage (FOD) by a factor of four when measured against conventional narrow
blades.
When the V2500 entered service this unique blade had accumulated five
years of in-service experience of Rolls-Royce RB211 series engines. To date, hollow
wide-chord fan blades have achieved over 50 million hours in-service experience
worldwide.
Pratt and Whitney's "float wall" combustor is another example of V2500
technology which fits this category. The combustor liner consists of sheet metal shells
to which turbine alloy segments are attached. These segments "float" on the cooling air
between the segments and the outer shell. The design improves cooling effectiveness,
eliminates stresses and the segment can be replaced individually lowering maintenance
cost.
Fuel efficiency is another factor, the V2500 burns up to 4% less fuel than the
competition, which is equivalent to 5500 barrels less per aircraft per year.
This
reduction in overall cost is achieved by a number of component efficiencies: the
upgraded wide chord fan gives the highest flow/unit area and lowest drag; the
utilization of the ten-stage high-pressure compressor creates extra efficiency as does the
two-stage high-pressure turbine and the five stage low-pressure turbine.
ROLLS ROYCE decided to team up with Pratt & Whitney, along with MTU
engines of Germany, Fiat of Italy & Japan Aero engines to form "I.A.E." and build a
mid-range high by-pass two-stage turbofan, the "V2500" series.
The initial "V2500-AI" engine variant was introduced in 1989 for airbus
A320 twin jet airliner and featured a single-fan and four-associated Low pressure
compressor stages; a 10 stage High pressure compressors and an annular combustor
with 20 fuel injectors, a 2-stage air cooled High pressure turbine; a 5-stage uncooled
Low pressure turbine. The engine features a FADEC and the turbines have active
clearance control. It provides a maximum take off thrust of 111 kN (11,335 kgf or
25,000 lbf).
The fan and compressor are derived from ROLLS ROYCE technology, while
the turbines, gearbox and FADEC (Full Authority Digital Engine Control ) are derived
from Pratt and Whitney technology. The V2500 series has been used on the Airbus
A321 and A320 twin jet airliners as well as the McDonnel Douglas MD-90 (now
Boeing 717) twin jet airliners.
International aero engine V2500-AI Type
SPECIFICATIONS
Fan Diameter
Length
Dry weight
Maximum takeoff thrust
Pressure ratio
Bypass ratio
METRIC SYSTEM
1.6 m
2.96 m
2,242 kg
111 kN, 11,335 kgf
36:1
5.7:1
ENGLISH SYSTEM
5 ft 3 inches
10 ft 2.1 inches
4,943 lbs
25000 lbf
-------------
Tsfc (thrust specific
fuelconsumption)
15.81 mg/ N-s
0.560 lb/lb-hr
Airflow
Twr(max takeoff thrust)
358 kg/s
5.06
789 lb/s
Different stages with blade specifications
FAN-TYPE
LOW PRESSUE COMPRESSOR
HIGH PRESSURE COMPRESSOR
COMBUSTOR
HIGH PRESSURE TURBINE
LOW PRESSURE TURBINE
SINGLE-STAGE
3 STAGES
10 STAGES
ANNULAR, 20 FUEL INJECTORS
2 STAGES, air cooled
ACC 5 STAGES, air cooled
GENERAL:
Two IAE V2500 turbofan engines designed for subsonic commercial
airline service power the A320 aircraft. Each engine is housed in a nacelle
suspended from a pylon attached below the wing. The right and left
power plants are inter-changeable except for the thrust reverser ‘C’duct.
1
POWERPLANT
0.1 POWER PLANT - GENERAL DESCRIPTION AND OPERATION
0.1.1 General
The aircraft is powered by two IAE V2500 turbo fan engines designed
for subsonic commercial airline service. Each engine is housed in a
nacelle suspended from a pylon attached below the wing. The right
and left power plants are interchangeable except for the thrust
reverser C ducts.
0.1.2 Description
A. Engine
The V2500 is a two spool, axial flow, high by-pass ratio turbo fan
engine. The design and configuration of the engine are based on
obtaining long life high reliability and easy access for line
maintenance. The V2500 incorporates a Full Authority Digital Engine
Control (FADEC).The control system governs all engine control
functions including power plant management.
The main modules of the engine are: the Low Pressure (LP) compressor (fan and
booster) assembly, the LP compressor/intermediate case, the No. 4 bearing and
combustion section, the High Pressure (HP) compressor, the HP turbine section, the LP
turbine section and the accessory drives(gearbox).
B. Cowling
The cowling assembly consists of:
•
the air intake cowl,
•
the fan cowl 437AL, 438AR, 447AL, 448AR,
•
the thrust reverser 451AL, 452AR, 461AL, 462AR,
•
the Common Nozzle Assembly (CNA).
The hinged fan reverser and fan cowls are attached to the pylon. The
fan cowls are hinged at the upper part by four hinges. They are held
open by hold-open rods providing access to the engine for:
Fig.1.1 Component Location
•
maintenance,
•
rigging,
•
trouble shooting.
The composite components of the nacelle incorporate a lightning
protection system consisting of:
•
a conductive graphite material within the skin.
Mechanical attachments and fittings conduct the current into the
pylon/engine structure.
C. Mounts
The engine is attached to the pylon by two damage tolerant mounts.
The mounts are the main attaching structures (forward and aft).The
forward mount comprises five attach bolts that connect the engine
directly to the pylon. The aft mount consists of a link and beam
system attaching the engine to the pylon.
D. Attachment Fittings
The attachment fittings and support brackets ensure the attachment
on the
•
engine of:
•
components,
•
ducts,
•
pipes,
•
electrical cables.
The attachment fittings and support brackets are attached on:
•
the engine core,
•
the fan case,
•
the accessories,
•
the accessory gearbox.
E. Fire Seals
The fire walls and fire seals provide fire protection (to a fire proof
standard) between the power plant designated fire zones (fan and
core compartments).
F. Electrical Harness
The engine electrical harness:
•
distributes the power required by the aircraft electrical system,
•
supplies the 115VAC and 28VAC power to the engine and
nacelle systems.
•
transmits signals for:
-
nacelle sub-systems
-
engine control
-
monitoring functions
G. Engine Drains
The drain and vent system consists of lines collecting and carrying
waste fluids and vapors overboard through the system drain mast.
This system drains the gearbox mounted accessories and engine
components
0.2 COWLING - DESCRIPTION AND OPERATION
0.2.1 General
The cowls enclose the periphery of the engine so as to form the
engine Nacelle. The nacelle ensures airflow around the engine during
its operation and also provides protection for the engine and
accessories. This section is a description of the following cowls:
•
engine air intake,
•
fan cowl,
•
the thrust reversers and the common nozzle assembly
NOTE : Fan cowls and thrust reversers are not removed for an engine
change since they are hinged to the pylon.
Fig.1.2 Nacelle Component
0.3 COWL - AIR INTAKE - DESCRIPTION AND OPERATION
0.3.1 General
The air intake cowl structure is an interchangeable aerodynamically
faired assembly which is mounted on the front of the engine fan case.
The assembly is composed of:
•
an inner and outer barrel,
•
a nose lip,
•
a forward and aft bulkhead.
The assembly also includes installation of:
•
the anti-icing ducting,
•
the phone jack,
•
the P2/T2 probe,
•
hoisting provisions,
•
a drainage provision,
•
air intake collecting atmospheric air to ventilate the fan case
compartment
0.3.2 Description
A. Air Intake Cowl Configuration
•
The outer skin assembly of the intake cowl is constructed of a
carbon fiber composite solid laminate.
•
The inner barrel consists of acoustically treated carbon
fibercomposite/Nomex honeycomb which is bolted to the
engine fan casing front flange.
The aft bulkhead and nose lip/forward bulkhead assembly connect the
outer
barrel to the inner barrel.
B. Air Intake Cowl Anti-Icing
A piccolo tube is mounted in the air intake cowl lip and distributes
anti-icing air into the lip inner surface. The anti-icing air is supplied to
the piccolo tube by a supply tube that penetrates the forward and aft
bulkheads
Fig1.3 Air Intake Assembly Details
C. Air Intake Cowl Structure
The majority of the internal pressure loads and internal air loads are
taken
in
hoop
tension
through
the
inner
barrel
skins.
The
longititudinal and transverse loads are distributed into the fan case
forward flange through a bolted joint. The acoustic panels are
structural and carry air intake cowl loads. The air intake cowl aft
bulkhead forms a land on which the fan cowl doors leading edges are
supported when closed.
D. Air Intake Cowl Materials
The intake cowl aft bulkhead and the rear of the inner barrel are constructed of carbon
composite sandwich and together provide a firewall barrier to the fan case
compartment. The intake cowl aft bulkhead is constructed of titanium and the rear of
the inner barrel is constructed of carbon composite sandwich, and together provide a
firewall barrier to the fan case compartment.
0.4 COWL - FAN - DESCRIPTION AND OPERATION
0.4.1 General
The left and right fan cowl assemblies enclose the engine fan case
between the air intake cowl and the thrust reverser. Each door is
interchangeable from one engine to the other and is attached to the
pylon by three hinges. A fourth hinge at the forward end of each door
connects to a common tie link between each door. The doors are
latched together along the bottom centre line by four adjustable
tension hook latches. Each inboard fan cowl has a strake attached to
it. The strake helps give smooth airflow between the cowl and the
fuselage to decrease turbulence. The strakes are attached to the cowl
with Hi-Loks and jointing compound.
0.4.2 Description
The fan cowl doors are constructed from a sandwich of carbon fibre
composite skins and an aluminium honeycomb core. Both doors are
supported on land formed in the air intake cowl at the front and in
the thrust reverser at the rear. The forward land incorporates
alignment fittings.
Two hold-open rods engage into brackets on the engine fan case to support the fan cowl
doors in the 55 degrees open position for ground maintenance only. A pressure relief
door is located in the right fan cowl door to limit fan case compartment pressure.
0.5 MOUNTS - DESCRIPTION AND OPERATION
0.5.1 General
The engine is attached to the aircraft pylon by two mount assemblies,
one at the front and one at the rear of the engine. The mount
assemblies transmit loads from the engine to the aircraft structure.
Spherical bearings in each mount permit thermal expansion and some
movement between the engine and the pylon. Both mounts are made
to be fail-safe and have a tolerance to damage.
0.5.2 Component Location
The front mount is installed at the top centre of the Low Pressure
Compressor (LPC) case. The rear mount is installed at the top centre
of the Low Pressure Turbine (LPT) case.
0.5.3 System Description
The engine mount system has these components:
•
A front mount 1500KM,
•
A rear mount 1550KM.
0.5.4 Operation
A. The Front Mount 1500KM
•
The thrust of the engine is transmitted through the thrust links,
the cross beam assembly and the beam assembly to the
aircraft pylon.
•
Vertical and side loads are transmitted through the support
bearing to the beam assembly and then to the aircraft pylon.
•
The support bearing permits the engine to turn so that torsional
loads are not transmitted to the aircraft structure.
•
The front mount is made to be fail-safe. If one of the two thrust
links or the cross beam should fail, then thrust loads are
transmitted through the ball stop and into the beam assembly.
The thrust is then transmitted to the pylon structure.
B. The Aft Mount 1550KM
•
Vertical and side loads are transmitted through the side links
and beam assembly and into the pylon.
•
Torsional loads are transmitted by the center link to the beam
assembly and in to the pylon.
•
The mount is made to be fail-safe. The side links are each made up of two parts
which are attached together to make one unit. If one part of the link should fail,
the remaining part will transmit the loads to the beam assembly.
0.6
ELECTRICAL HARNESS - DESCRIPTION AND OPERATION
0.6.1 General
The engine electrical harnesses connect the electrical components
installed in the nacelle to the aircraft electrical systems. The
harnesses have two primary assemblies. The fan zone harness and
the core zone harness. Each of these primary assemblies has smaller
harness assemblies .The core zone harnesses are connected to the
fan zone harnesses at the bifurcation panel. The harnesses are
installed around the engine and go up to the pylon break points. Here
they interface with the aircraft electrical systems. The harnesses are
attached to the engine with brackets, raceways, clips and clamps.
0.6.2 Description
A. The harness assemblies
•
•
that are part of the fan zone harness are:
-
The Electronic Engine Control (EEC) harness.
-
The EEC and ignition supply harness.
-
The general supply harness.
These harness assemblies are connected to the following
engine components:
-
The EDA.
-
The ACOC modulating valve.
-
The stage 10 solenoid valve.
-
The pneumatic starter valve.
-
The stage 7 solenoid valves.
-
The ACOC oil temperature thermocouple.
-
The relay box.
-
The EEC.
-
The Fuel Metering Unit (FMU).
-
The fuel flow transmitter.
-
The Integrated Drive Generator (IDG) fuel temperature
thermocouple.
•
-
The FCOC fuel temperature thermocouple.
-
The fuel diverter and return valve.
that are part of the core zone harness are:
-
The fire detection (System A), cabin services and nacelle
over temperature harness.
•
-
The fire detection (System B) harness.
-
The EEC harness.
-
The EEC link harness.
-
The Exhaust Gas Temperature (EGT) harness.
-
The ignition harness.
These harness assemblies are connected to the following
engine components:
-
The Active Clearance Control (ACC) actuator.
-
The Variable Stator Vane (VSV) actuator.
-
The Low Pressure Control (LPC) bleed master actuator.
-
The T2.5 CM terminal.
-
The terminal block.
-
The igniter boxes.
-
The 10th stage valve.
-
The T3 sensor.
-
The EGT thermocouples.
0.6.3 Operation
The engine electrical harness supply the power that is necessary for
the electrical systems. They also transmit the signals for the nacelle
sub-systems and the engine control and monitoring functions.
0.7 POWER PLANT DRAINS - DESCRIPTION AND OPERATION
0.7.1 General
The power plant drain system collects fluids that may leak from some
of the
engine accessories and drives. The fluids collected from the power
plant are
discharged overboard through the drain mast.
0.7.2 System Operation
A. General
•
•
The drain system comprises two sub-systems:
-
fuel drains,
-
oil, hydraulic and water drains.
The two sub-systems come together at the same drain mast.
This drain mast is installed below the engine accessory
gearbox.
B. Fuel Drains
The fuel drain lines come from engine accessories on the engine core,
the engine fan case and gearbox. The engine core drains go through
the bifurcation panel. The fuel drain system is connected to these
engine accessories:
•
Booster bleed master actuator
)
•
Booster bleed slave actuator
) Engine
•
Variable Stator Vane Actuator
) Core
•
Active Clearance Control Actuator )
•
Fuel diverter valve
) Engine fan case
•
Air Cooled Oil Cooler actuator
) Engine fan case
•
Fuel metering unit
) Gearbox
•
LP/HP fuel pumps
)
Fuel that can leak from these accessories is removed by a steel tube
connected to the accessory or gearbox mounting pad. The tubes are
shaped to go around the engine and go in the side of the drain mast.
Each tube is connected by a union with a packing seal at the
accessory or gearbox pad.
C. Oil, Hydraulic and Water Drains
The oil, hydraulic and water drains system comes from engine
accessories on the engine fan case and gearbox. The drain system is
connected to these engine accessories:
•
Integrated Drive Generator )
•
Air starter
) Gearbox
•
Hydraulic Pump
)
•
Oil tank scupper
) Oil tank
The only hydraulic fluid drain is from the hydraulic pump. The other
drains are for engine oil or accessory lubricant. Each tube is
connected
by a union with a packing seal at the accessory or gearbox pad
1
ENGINE
1.1 ENGINE – GENERAL – DESCRIPTION AND OPERATION
1.1.1 General
Fig.2.1 Engine Bearing
The engine is a two spool, axial flow, high bypass ratio turbofan engine. Its
compression system features a single stage fan, a three stage booster, and a ten stage
High Pressure Compressor (HPC). The Low Pressure Compressor (LPC) is driven by a
five stage Low Pressure Turbine (LPT) and the HPC by a two stage High Pressure
Turbine (HPT). The HPT also drives a gearbox which, in turn, drives the engine and
aircraft mounted accessories. The two shafts are supported by five main bearings. The
engine incorporates a full authority digital Electronic Engine Control (EEC). The
control system governs all engine functions, including power management. Reverse
thrust for braking the aircraft after landing is supplied by an integrated system which
acts on the fan discharge airflow.
1.1.2 Component location
Fig.2.2Component Location
1.1.3 Engine Modules
The engine modules are
•
the fan module
•
the intercase module
•
the HPC
•
the diffuser/combustor module
•
the HPT
•
the LPT
•
the accessory drive gearbox.
A. Fan Module
It consists of a single stage, wide-chord, shroudless fan and hub.
B. Intercase Module
It consists of the fan containment case, fan Exit Guide Vanes (EGV),
intermediate case, booster, low spool stub shaft, the accessory
gearbox tower shaft drive assembly, high spool stub shaft and the
station 2.5 bleed valve (BSBV). The booster consists of inlet stators,
rotor assembly, and outlet stators. The No. 1, 2 and 3 (front) bearing
compartment is built into the module and contains the support
bearings for the low spool and high spool stubshafts.In conjunction
with the inner fan section, the booster increases the pressure at the
entrance to the HPC and provides an even pressure profile to improve
efficiency. The station 2.5 bleed is used for engine handling by
controlling airflow to the high compressor entrance.
C. High Pressure Compressor
The HPC is a ten stage, axial flow module. It is comprised of the drum
rotor assembly, the front casing which houses the variable geometry
vanes and the rear casing which contains the fixed geometry stators
and forms the bleed manifolds.
D. Diffuser/Combustor Module
The combustion section consists primarily of the diffuser case,
combustor, fuel injector and ignitors. The high compressor exit guide
vanes and the No. 4 bearing compartment are also part of the
module. The main features of the module include a close-coupled
prediffuser and combustor that provide low velocity shroud air to feed
the combustor liners and to minimize performance losses.
E. High Pressure Turbine
The HPT is a two stage turbine and drives the HPC and the accessory
gearbox.
F. Low Pressure Turbine
The LPT is a five stage module. The elliptical leading edge airfoils
improve the aerodynamic efficiency. Module efficiency is further
enhanced by incorporation of rim seals and clustered vanes which
results in reduced losses due to leakage. Active clearance control is
used to control seal clearances and to provide structural cooling.
G. Accessory Drive Gearbox
The accessory drive gearbox provides shaft horse power to drive engine and aircraft
accessories. These include fuel, oil and hydraulic pressure pumps and electrical power
generators for the EEC and for the aircraft. The gearbox also includes provision for a
starter which is used to drive the N2 shaft for engine starting.
1.2 COMPRESSOR SECTION – DESCRIPTION AND OPERATION
1.2.1 General
The compressor section consists of three modules:
•
LP compressor (fan) module assembly,
•
LP compressor/intermediate case module,
•
HP compressor.
Rotation of the fan rotor causes air to be ingested into the front of the
engine and to be compressed. Compressed air is then divided into
two separate airflows:
•
a large portion is delivered to the exhaust nozzle,
•
the remainder is compressed in the booster before being again
compressed by the HP compressor.
1.2.2 LP Compressor (Fan) Module Assembly
The LP compressor (fan) module is a rotor assembly which includes
twenty two titanium blades and a titanium disk. Rotation of the rotor
causes air to be ingested into the front of the engine and
compressed. A larger proportion of the compressed air is delivered
through the fan discharge duct to the exhaust nozzle to provide the
majority of engine thrust. The remainder of the compressed air
passes into the booster section for further compression by the
booster.
1.2.3 LP Compressor / Intermediate case module
The LP compressor intermediate case module consists of booster
section, a fan case section and an internal gearbox and drive section.
1.2.4 HP Compressor Module
The HP compressor is a 10 stage axial flow module. It comprises the
HP compressor rotor, blades, the front casing and variable vanes, the
rear casing which contains the fixed stators and forms the bleed
manifolds. Mounted on the front casing is the linkage system
associated with the variable inlet guide vanes and stators. Attached
to the rear of the compressor rotor is the rear thrust balance seal
rotating member. Power to drive the HP compressor is provided
through the rear shaft from the HP turbine system.
1.3 COMBUSTION SECTION – DESCRIPTION AND OPERATION
1.3.1 General
The combustion section includes the diffuser and combustion group,
the No. 4 bearing section and the turbine nozzle assembly. The
combustion section has four primary functions:
•
straighten the flow of air from the HP compressor,
•
change the flow of air characteristics to get the best speed and
pressure for combustion,
•
mix fuel with the air and supply ignition to make the fuel burn,
•
hold the No. 4 bearing in position.
1.3.2 Diffuser Case And Combustor Assemblies
A. Diffuser Case
Fig.2.3 Diffuser Case And Combustor Assemblies
The diffuser case is a main structural part of the engine. The diameter
of the diffuser section is larger at the rear than at the front. This
diametral difference decreases the speed of the air and changes the
energy of the speed into pressure. The diffuser case has 20 mounting
pads, where the fuel injectors are installed, two mounting pads where
the ignitor plugs are installed and five borescope bosses located
around the case.
B. Combustor
The combustor is an annular type which consists of an outer liner
assembly. The liner is fitted with mechanically attached segments
which form the inner wall of the combustion chamber. Air which
surrounds the combustor is used in the combustion process for
dilution and exit temperature profile control of the combustion gases
and for cooling of the combustor walls. The front of the combustor
outer liner is secured to the diffuser case outer wall by five
combustion chamber retaining bolts. A seal is provided at the
transition from the rear of the combustor outer liner to the turbine
nozzle guide vanes. The rear of the combustor inner liner is bolted to
the stage 1 HP turbine blade cooling duct. The cooling duct directs
and meters HP compressor exit air to the stage 1 HP turbine blades.
The front lip of the combustor inner liner forms a seal with the inner
diameter of the combustor hood.
1.3.3 Turbine Nozzle Assembly
Fig.2.4 Turbine Nozzle Assembly
The stage on the turbine nozzle assembly consists of an outer ring, 40
cobalt alloy vanes welded together in pairs, inner combustor linerassembly and the stage 1 HP turbine blade cooling duct assembly.
The vane pairs are retained at the outer end by a support ring which
is bolted to the rear outer flange of the diffuser case. The outer ends
of the vane pairs are retained such that the vanes may slide radially
under thermal growth. The vane pairs are bolted to the stage 1 HP
Turbine Blade Cooling Duct Assembly at the inner end. Each vane is
cooled by air which enters at the outer and inner ends and exits
through airfoil holes into
the primary gas path.
1.3.4 Bearing Compartment Assembly
The No. 4 bearing compartment consists of front and rear walls which attach to the No.
4 bearing support assembly. The bearing support assembly, in turn, is bolted to the
diffuser case rear inner flange. The compartment walls are surrounded by a cooling duct
which is itself insulated by a heat shield. The compartment walls also provide support
for the carbon seals. The No. 4 bearing compartment service tubes connect to the front
wall and supply the compartment with oil and cooling air. 12th stage compressor air is
directed through an external air-to-air heat exchanger and carried by service tubes to the
bearing compartment cooling duct.
1.4
TURBINE SECTION – DESCRIPTION AND OPERATION
1.4.1 General
Fig.2.5 Turbine Section Assembly
The turbine section consists of the HP and LP turbine modules.The HP
turbine uses combustion gases to drive the HP compressor and the
accessory gearbox, and provides a gas stream to the LP turbine in
order to drive the LP compressor and the fan through the LPT shaft.
1.4.2 HP Turbine Section
A. General
The HP Turbine Rotor and Stator Assembly provides the rotational
driving force for the HP compressor and accessory gearbox by
extracting energy from the hot combustion gases. It consists of a
Stage 1 Turbine Rotor Assembly ; a HP Turbine Case and Vane
Assembly ; a Stage 2 HPT Airseal ; and a Stage 2 Turbine Rotor
Assembly.
B. Cooling
All of the HPT airfoils are cooled by secondary air flow.The first stage
HPT blades are cooled by the HPC discharge air which flows through
the first stage HPT duct assembly. The velocity of the air increases to
the outside between the turbine front hub and the first stage HPT
(front outer) air seal into the blade root, thus providing(once the
speed is converted back into pressure) the pressure differential
required to ensure cooling air flow.The second stage vane clusters are
cooled by tenth stage compressor air mixed with thrust balance seal
vent air supplied externally. Air flows into the case and through the
center of each vane and then outward into the turbine area and the
gaspath. Some of this air is used for cooling of the second stage HPT
air seal.Second stage HPT cooling air is a mixture of HPC discharge air
and tenth stage compressor air. This air moves through holes in the
first stage HPT(front inner) air seal and the turbine front hub into the
area between the hubs. The air then goes into the second blade root
and out thecooling holes.
C. Clearance Control
The abradable duct segments and abrasive blade tips, along with
active clearance control, keep tight blade tip clearances for better
performance.The abrasive/abradable system makes tight clearances
by letting the parts rub. The abrasive decreases blade tip wear during
rub.Active clearance control tubes around the turbine case supply fan
discharge air to cool the surface of the case during climb and cruisepower operation.Cooling results in shrinkage of the case and
decreased blade tip clearances.
1.4.3 Low Pressure Turbine Section
The five stage Low Pressure Turbine (LPT) extracts energy from the
gas stream delivered from the HP Turbine in order to provide a
mechanical drive through the LPT shaft to the LP Compressor and the
Fan. Exhaust gas from the LPT passes through a nozzle to provide
propulsive thrust.Seal clearance and LPT case heat expansion are
controlled by an external Active Clearance Control (ACC) System. Fan
discharge air is directed externally to the LPT case via the ACC tubes.
This controls the heat expansion of the LPT case and optimizes the
seal clearances.
1.5 ACCESSORY DRIVES (EXTERNAL GEARBOX) – DESCRIPTION
AND OPERATION
1.5.1 General
The external gearbox has an angle gearbox assembly, a main gear
assembly and external components.The external gearbox is installed
at the bottom of the fan case. Four articulated support links attach
the gearbox to the fan case. The links have spherical bearings at each
end
to
allow
for
any
necessary
mount
articulation.
Fig.2.6 Accessory Drives
1.5.2 Angle Gearbox
A. General
The angle gearbox transmits power from the engine to the main
gearbox and from the starter on the main gearbox to the engine.
B. Description
The angle gearbox has a bevel gear set held by a cast aluminium
housing. The bevel gear set transmits the power to and from the
engine through a power shaft engaged to the high pressure
compressor rotor. The bevel gear set has a spiral gear mesh which
drives a horizontal input gear shaft in the main gearbox. Two metered
jets supply pressure oil to the bearings and gears in the angle
gearbox.
1.5.3 Main Gearbox
A. General
The main gearbox is installed forward of the angle gearbox. The main
gearbox transmits power from the engine to the accessories installed
on the gearbox and from the starter to the engine. The main gearbox
supplies speed torques necessary for the accessories to perform their
various functions.
B. Description
The main gearbox has a cast aluminium housing that has a rear train
and mounting pads for the airframe and engine accessories. An
external de-oiler is installed on the front face of the main gearbox.
Each of the accessories drive geartrain sections is individually
replaceable. The metered oil nozzles are installed on a gearbox
housing and supply pressure oil to the bearings and gears in the
gearbox. An external oil tank is attached at the LH flange of the main
gearbox.
3.ENGINE FUEL AND CONTROLLING
3.1 ENGINE FUEL AND CONTROL – GENERAL – DESCRIPTION
AND OPERATION
3.1.1 General
The fuel system enables delivery of a fuel flow corresponding to the
power required and compatible with engine limits.The system
consists of:
•
the two stage fuel pump with low pressure and high pressure
elements,
•
the Fuel Metering Unit (FMU),
•
the engine Fuel Cooled Oil Cooler (FCOC),
•
the Integrated Drive Generator (IDG) Fuel Cooled Oil Cooler
(FCOC),
•
the fuel filter,
•
the fuel distribution valve,
•
the fuel flow meter,
•
20 fuel nozzles,
•
the fuel diverter and return (to tank) valve.
3.1.2 Distribution
The fuel supplied from aircraft tanks flows through a centrifugal pump
Low Pressure (LP) stage then through the FCOC and then through a
filter and a gear pump High Pressure (HP) stage. The fuel from the HP
pump is delivered to the FMU which controls the fuel flow supplied to
the fuel nozzles (through the fuel flow-meter and the fuel distribution
valve). The FMU also provides hydraulic pressure to all hydraulic
system external actuators. These include the booster stage bleed
valve actuators, stator vane actuator, Air Cooler Oil Cooler (ACOC) air
modulating valve, High Pressure Turbine Active Clearance Control
(HPTACC) and Low Pressure Turbine Active Clearance Control
(LPTACC) valve. Low pressure return fuel from the actuators is routed
back into the fuel diverter and return valve. The fuel diverter and
return valve enables the selection of one of the four basic
configurations between which the flow paths of the fuel in the engine
are varied to maintain the IDG oil, engine oil and fuel temperatures
within specified limits. The transfer between configurations is
determined by a software logic contained in the Electronic Engine
Control (EEC).
3.1.3 Controlling
The Full Authority Digital Electronic Control (FADEC) system provides
full range control of the engine to achieve steady state and transient
performance when operated in combination with aircraft subsystems.
The FADEC is a dual channel EEC with crosstalk and failure detection
capability. In case of failure detection, the FADEC switches from one
channel to the other.
3.1.4 Indicating
The engine fuel system is monitored from:
•
the ECAM display,
•
the warning and caution lights.
The indications cover all the main engine parameters through the
FADEC. The warnings and cautions reflect:
•
the engine health and status through the FADEC,
•
the FADEC health & status,
•
the fuel filter condition through a dedicated hardwired pressure
switch.
Fig.3.1 Fuel System Schematic
3.2 DISTRIBUTION – DISCRIPTION AND OPERATION
3.2.1 General
The engine fuel supply distribution system mainly consists of:
•
a fuel supply line
•
an engine 2-stage pump High Pressure/Low Pressure (HP/LP),
•
a fuel filter,
•
an engine Fuel Cooled Oil Cooler (FCOC),
•
a Fuel Metering Unit (FMU),
•
an Integrated Drive Generator (IDG) Fuel Cooled Oil Cooler
(FCOC),
•
a fuel diverter and return (to tank) valve,
•
a fuel flow-meter,
•
a Fuel Distribution Valve (FDV),
•
20 fuel nozzles.
3.2.2 Fuel Manifold And Fuel Tubes
The fuel manifold and fuel tubes consist of several single wall tubes
which carry fuel between components in the fuel system. Fuel
supplied to the fuel nozzles is carried by a large tube from the FMU to
the fuel distribution valve. At the fuel distribution valve the fuel
supply is split and carried to twenty fuel nozzles by ten manifolds.
Each fuel manifold feeds two fuel nozzles. Fuel pressure for actuating
various valves is supplied by small tubes from the FMU mounted on
the fuel pump. All the brackets and tubing are fireproof.
3.2.3 Fuel Pump
The Low Pressure/High Pressure (LP/HP) fuel pumps are housed in a
single pump unit which is driven by a common gearbox output shaft.
A LP stage and a HP stage provide fuel at the flows and pressures
required for the operation of the hydromechanical components and
for combustion in the burner. The unit consists of a LP centrifugal
boost stage which feeds an HP single stage, two gear pumps. The
housing has a provision for the installation of the FMU. Fuel from the
aircraft tanks flows to the LP stage of the engine fuel pumps, through
the aircraft fuel pumps. The LP pump is designed to provide fuel to
the HP gear stage with the aircraft pumps inoperative. After passing
through the LP boost stage, the fuel flows through the fuel filter to the
HP gear stage. A coarse mesh strainer is provided at the inlet to the
HP gear stage. This stage is protected from overpressure by a relief
valve. Excess flow from the gear stage pump is recirculated through
the FMU bypass loop to the low pressure side of the pump.
3.2.4 Fuel Filter
A. Description
•
The fuel filter element is a low pressure filter which removes all
contamination from fuel to go through it.
•
The filter element is installed in the lower housing of a Fuel
Cooled Oil Cooler (FCOC). The FCOC includes the following
components:
-
A filter cap which has a pressure plate to keep the filter
element in position once installed.
-
A filter bypass valve to let the fuel go around the filter
element when it becomes clogged.
The filter cap of the FCOC also includes a fuel drain plug to drain the
fuel for maintenance purposes.
B. Operation
•
The fuel from the FCOC goes through the filter element into the
high pressure gear element of an LP/HP fuel pump.
•
The filter bypass valve keeps a pressure drop across the filter
element to a maximum of 17 psi (1.17 bar) differential. If the
pressure drop is higher than the maximum limit, the bypass
valve will start to open and let the fuel go around the filter
element.
3.2.5 Fuel Nozzle
A. General
The fuel nozzles receive fuel from the fuel manifolds. The fuel nozzles
mix the fuel with air, and send the mixture into the combustion
chamber in a controlled pattern.
B. Description/Operation
There are 20 fuel nozzles equally spaced around the diffuser case
assembly. The fuel nozzles are installed through the wall of the case,
and each nozzle is held in position by three bolts. The fuel nozzles
carry the fuel through a single orifice. The fuel is vaporized by highvelocity air as it enters the combustion chamber. The fuel nozzle
forms the atomized mixture of fuel and air into the correct pattern for
satisfactory combustion. The design of the fuel nozzle results in fast
vaporization of the fuel through the full range of operation. The highvelocity flow of fuel prevents formation of coke on areas where fuel
touches metal. Heat shields installed internally and externally also
prevent formation of coke.
3.2.6 Fuel Diverter And Return Valve
A. General
The Fuel Diverter and Return Valve (FD and RV) is a primary unit in
the Heat Management System (HMS) of the engine. The FD and RV
has two valves in one body. They are a Fuel Diverter Valve (FDV) and
a Fuel Return Valve (FRV).The FDV operates to change the direction
of the FMU spill flow to:
•
The Fuel Cooled Oil Cooler (FCOC) or,
•
the fuel filter (element) inlet or,
•
the IDG FCOC.
The FRV operates to control fuel flow which goes back to the aircraft
fuel tank acting as a fuel cooler.
B. Description
The fuel diverter and return valve is installed on the FCOC. The FDV is
a two-position selector valve which has two pistons in a sleeve. The
two pistons are mechanically connected and make two valve areas
which are referred to as valve A and valve B. The FRV has a main
valve and a pushing piston in a sleeve. This main valve is a half-area
piston-type valve which moves the valve to change the metering port
area. The main valve has two valve functions that are referred to as
valve C and valve D The EEC gives the electrical signal to the FDRV to
change the position of the valves. The FDRV gives a feedback signal
to the EEC to transmit the position of valves in the unit. The fuel flow
changes with the position of the valves. Thus, the fuel flow can be
controlled through the FDRV and the EEC.
C. Operation
(1) General
The FDRV configuration allows four modes of operation according to
the electrical signals generated from the EEC (based on fuel and oil
temperature measurements transmitted by means of thermocouples).
(2) Fuel return valve
•
The EEC operates the dual-wound torque motor to control the
servo pressure. This servo fuel pushes the main valve.
•
The pressure balance between two sides of the main valve
(Valves C and D) gives the direction and the speed of the valve
movement.
Then the valve changes the direction of the fuel flow and controls the
metering port area.
(3) Fuel diverter valve
•
The EEC energizes the solenoid valve to let the servo fuel flow.
This servo fuel goes into one side of the piston face in the valve
B.
•
The servo pressure pushes two pistons (which are the valves A
and B) in the same direction. Then these valves change the
direction of the fuel flow, and one of these pistons compresses
the spring.
•
When the solenoid is de-energized, this spring pushes back two
pistons.
(There is an orifice to release the servo fuel to the FMU spill port).
And the other one of two pistons pushes the switch assemblies. The
switch assemblies transmit the EEC the valve position when the
solenoid is de-energized.
(4) Constant Pressure Valve (CPV)
The CPV makes the servo pressure constant between the HP port and
LP port of the FDRV. This servo flow moves each valve in the FDRV.
(5) Failure mode
•
When the servo pressure becomes zero:
-
The pushing piston comes up to hold the main valve at a
mode 5 position.
-
The spring extends to hold the FDV pistons at a mode 5
position.
•
Other than in (a), the FDRV keeps the mode 5 position in these
conditions:
-
The failure of the electrical signal.
-
During the engine stops.
•
If the IDG FCOC port is clogged in this valve position, relief
valve releases the FMU spill flow. This relief valve is in the valve
A and it can release the unwanted pressure to the FCOC port.
1.5.4 Fuel Distribution Valve
A. General
The FDV subdivides scheduled engine fuel flow from the FMU equally
to ten fuel manifolds, each of which in turn feeds two nozzles.
B. Description
The fuel distribution valve is installed at the 4:00 o_clock location, at
the front flange of the diffuser case. The fuel distribution valve
receives fuel through a fuel line from the FMU. The fuel goes through
a 200 micron strainer, and then into ten internal discharge ports. The
ten discharge ports are connected to the
ten fuel manifolds. Eight of the ten internal discharge ports in the
valve are connected after an engine shutdown. This lets fuel drain
from eight of the fuel
1.6 CONTROLLING – DESCRIPTION – AND OPERATION
1.6.1 General
The Full Authority Digital Electronic Control (FADEC) system provides
full range control of the engine to achieve steady state and transient
performance when operated in combination with aircraft subsystems.
The FADEC system consists of
•
a dual-channel FADEC unit;
•
Fuel Metering Unit (FMU);
•
dedicated Permanent Magnetic Alternator (PMA);
•
actuation systems for stator vanes,
•
engine bleeds,
•
active clearance control,
•
10th stage cooling air,
•
engine and Integrated Drive Generator (IDG) heat management
control;
•
sensors;
•
electrical harness;
•
and start system components.
The FADEC Electronic Engine Control (EEC) is a vibration-isolated, aircooled unit mounted on the engine fan case. Its vibration isolation
and cooling systems are specifically designed to provide a protected
and controlled internal environment that is completely compatible
with the electronic components.
1.6.2 System Description
A. FADEC
(1) FADEC Functions
The FADEC system operates compatibly with applicable aircraft
systems
to perform the following functions
•
GAS
generator
control
for
steady
state
and
transient
engineoperation within safe limits.
-
Fuel flow control
-
Acceleration and deceleration schedules
-
Variable Stator Vane (VSV) and Booster Stage Bleed
Valve (BSBV)schedules
-
Turbine clearance control (High Pressure/Low Pressure)
(HP/LP)
•
-
10th stage cooling air control
-
Idle setting.
Engine limits protection
-
Fan and core over speed protection to prevent engine
running over certified red lines
-
Engine turbine outlet gas temperature monitoring.
Power management
-
Automatic engine thrust rating control
-
Thrust parameter limits computation
-
manual power management through constant ratings versus
throttle lever relation
-
Automatic power management through direct engine power
adjustment to the auto thrust system demand.
Automatic engine start sequencing
-
Control of the starter valve ON/OFF
-
Control of HP fuel shutoff valve (ON/OFF on ground, ON in flight)
-
Control of the fuel schedule
-
Control of the ignition (ON/OFF)
-
Engine
Pressure
Ratio
(EPR),
GasTemperature (EGT) monitoring
-
Abort/Recycle capability on ground.
Thrust reverser control
-
Control of thrust reverser actuation
N1,
N2,
WF,
Exhaust
-
Control of engine power during reverser operation
-
engine idle setting during reverser transient
Fig.3.2 Full Authority Digital Elecetronic Controll – Schematic
-
Control of maximum reverser power at full rearward throttle
control lever position.
-
Restow command in case of non commanded deployment.
-
Redeploy command in case of non commanded stowage.
Engine parameters transmission for cockpit indication
-
Engine primary parameters
-
Starting system status
-
Thrust reverser system status
-
FADEC system status.
Engine condition monitoring parameters transmission.
Detection, isolation, accommodation and memorization of its internal
system failures.
Fuel return valve control. The FADEC controls the ON/OFF return to
the aircraft tank in relationship with:
-
Engine oil, IDG oil and fuel temperatures
-
Aircraft fuel system configuration
-
Flight phases.
B. Gas Generator Control
(1) Fuel Control
The EEC produces a fuel flow request using the control laws relevant
to engine operation. The request is transmitted through the torque
motor in the fuel metering unit. Setting steady state power, idle
speed and accel/decel transients requires different control laws. The
primary mode of setting steady state power is provided by controlling
fuel flow to set EPR as illustrated in An EPR Reference (EPR REF) is
calculated as a function of the Throttle Resolver Angle (TRA), ambient
temperature (T2), Mach number and altitude. The EPR reference is
compared to sensed EPR and dynamic compensation is then applied
to this EPR error. The result is that fuel flow is modulated until the
EPR error is eliminated. The rotor speed reference (N1 REF) will be
scheduled as a function of TRA and T2.
(2) Variable Stator Vane (VSV) Control
The VSV position is controlled by the EEC as a function of N2/square
root of theta T2.6. The EEC uses the VSV feedback signal from the
Linear Variable Differential Transducer (LVDT) to adjust the actual
VSV position.
(3) Booster Stage Bleed Valve (BSBV) Control
The BSBV position is controlled by the EEC. The EEC uses the BSBV
feedback signal from the LVDT to adjust the actual BSBV position.
(4) HPT/LPT Active Clearance Control (HPT/LPT ACC)
The HPT/LPT ACC valve modulates fan air flow to the HP and LP
turbine cases. The EEC controls the valve position as a function of the
thrust level. The LVDTs transmit the valve position to the EEC.
(5) HP Turbine (10th Stage) Cooling Air Control
The HP turbine cooling air valve supplies supplemental air (from HP
compressor 10th stage) to cool various parts of the HP and LP
turbines. The valve operates as a function of high rotor speed and
altitude and incorporates a 2-position switch to provide a feedback
signal to the EEC (channels A and B)
(6) Oil/Fuel Temperature Control
Heating and cooling of fuel, engine oil and IDG oil is accomplished by
the Fuel Cooled Oil Cooler (FCOC), the Air Cooled Oil Cooler (ACOC)
and the IDG cooler under the management of the EEC. Devices used
by the EEC include the fuel diverter valve, the ACOC modulating air
valve and the return to tank valve. Fuel, engine oil and IDG oil
temperatures are transmitted to the EEC by thermocouples. The fuel
temperature is measured at the exit of the filter. The engine oil
temperature is measured upstream of the ACOC. The IDG oil
temperature is measured at IDG oil cooler exit.
C. Engine Limits Protection
The FADEC prevents inadvertent over boosting of the expected
rating(EPR limit and EPR target) during power setting. It also prevents
exceedance of rotor speeds (N1 and N2) and burner pressure limits.
In addition, the FADEC unit monitors EGT and sends an appropriate
indication to the cockpit display in case of exceedance of the limit.
The FADEC unit also provides surge recovery.
D. Power Management
The FADEC unit contains all the engine thrust setting curves to
provide automatic engine thrust ratings control in Engine Pressure
Ratio, (EPR) (in normal mode) and N1 (in reversionary mode).The
FADEC unit computes power management LIMIT and COMMAND
parameters in EPR mode, except during reverse operation (N1 mode).
These parameters are available for the following engine thrust
modes:
•
Maximum Take-Off and Go-Around
•
Flexible Take-Off
•
Maximum Continuous
•
Maximum Climb
•
Idle (no limit parameter)
•
Reverse (N1 mode operation)
E. Engine Starting/Ignition Control
There are two modes of starting control associated with two different
procedures and corresponding to two engine starting logics in the EEC
(a) Automatic starting logic under the full authority of the FADEC
system. The FADEC initiates the automatic sequence of command to:
•
pneumatic starter valve opening and closing
•
HP fuel valve
•
igniters.
The FADEC provides:
•
engine limits protection N1, N2, EGT
•
on ground start abort in case of detected incident (hot start,
stall, failure to light, hung start)
•
in flight start, only fault indication, without automatic start
abort
•
specific fault message transmission.
(b) Alternate start logic with authority of the FADEC limited to:
•
actuation of the pneumatic starter valve, through the activation
of the MAN START pushbutton switch and the setting of the
ENG/MODE/CRANK/NORM/IGN START to IGN START.
•
energization of the spark igniter and setting of the ENG/MASTER
control switch to ON to energize the HP fuel shut off valve.
•
Stop of the ignition and starter air valve.
•
generation of warning indications.
F. Engine Parameter Transmission for Cockpit Display
The FADEC provides the necessary engine parameters for cockpit
display through the ARINC 429 data bus outputs.
G. Engine Condition Parameter Transmission
Engine Condition monitoring is provided by the ability of the FADEC to
transmit the engine parameters through the ARINC 429 data bus
output. The basic engine parameters available are:
•
WF, N1, N2, P5, PB, Pamb T4.9 (EGT), P2, T2, P3 and T3.
•
VSV, BSBV, 7th and 10th stage bleed commanded positions
HPT/LPT ACC,HPT cooling, WF valve or actuator position
•
status and maintenance words, engine serial number and
position.In order to perform a better analysis of the engine
condition, some additional parameters are optionally available.
These are P12.5, P2.5 and T2.
H. FADEC System Maintenance
The FADEC maintenance is facilitated by internal extensive Built in
Test Equipment (BITE) providing efficient fault detection. The results
of this fault detection are contained in status and maintenance words
according to ARINC 429 specification and are available on the output
data bus.
1.6.3 Component Description
A. Engine Sensors
•
T4.9 (EGT) Sensor
•
N1 Sensor
•
N2 Sensor
•
Engine Oil Temperature Sensor
•
P2/T2 Sensor
•
P3/T3 Sensor
•
P5 (4.9) Sensor
•
Fuel Temperature Sensor
B. Dedicated Permanent Magnet Alternator (PMA)
The alternator functions as the primary power source for the EEC and
transmits an N2 signal to the EEC and the cockpit. It comprises two
stators, a rotor and an N2 winding. The rotor is mounted directly on
the gearbox output shaft and the stator is bolted to the gearbox
housing. The alternator provides two identical and independent power
outputs, one for each channel of the EEC. The EEC obtains its N2
speed input by sensing the frequency of the input provided by the
alternator. A separate stator winding provides a dedicated frequency
signal to the Engine Vibration Monitoring Unit (EVMU), and is designed
to tolerate indefinite short circuit conditions. The stator and rotor are
sealed from the gearbox by a shaft seal. If a shaft seal failure occurs
and the alternator fills with engine oil, the alternator will continue to
function normally.
C. Engine Electronic Control (EEC)
The Electronic Engine Control is the main component of the engine
fuel and control system. The EEC receives data input from the other
aircraft systems and generates control signals to the engine systems
and components. The EEC also monitors the systems and components
to make sure they operate properly.
D. Fuel Metering Unit (FMU)
The Fuel Metering Unit (FMU) provides fuel flow control for all
operating conditions. Variable fuel metering is provided by the FMU
through EEC commands, by a torque motor controlled servo drive.
Position resolves provide feedback to the EEC. The FMU has provision
to route excess fuel above engine requirements to the fuel diverter
valve through the bypass loop.
E. Ignition Boxes
They are powered with aircraft 115VAC - 400Hz through the EIV and
the FADEC. The igniter A is powered from the emergency bus and the
igniter B is
Powered from the normal bus.
F. Pneumatic Starter Valve
The FADEC controls the opening/closing of the valve and receives the
open/not open signal from the valve.
G. VSV Feedback Signal
The FADEC receives a VSV position feedback signal from the VSV actuator.
1.7 FUNCTIONAL INTERFACES – DESCRIPTION AND OPERATION
1.7.1 General
The FADEC unit interfaces with the following aircraft functional
elements:
•
Air data computer which transmits air data signals to the
engine control system.
•
Engine Interface Unit which: concentrates airframe signals and
transmits them to the FADEC, receives information from the
engine and dispatches them to other systems.
•
Cockpit system display which furnishes engine parameters
indication and warnings to the crew.
•
Throttle control system which translates the crew commands
for engine power level into a command signal to the FADEC.
•
The thrust reverser control system.
•
The AIDS interface which records engine data for maintenance
purposes.
•
Electrical power supply from airframe to power FADEC while
engines are not running.
1.7.2 System Description
A. FADEC Inputs/Outputs
Digital inputs/outputs of the FADEC conform to ARINC 429-7
specification.
(1) Digital Inputs
Inputs to each channel are isolated from each other in order to
prevent failure propagation between channel A and B and/or between
both engines. Each channel of the EEC has input ports for both ADIRU
and EIU plus an input port for a spare. The FADEC makes fault
detection on its inputs by performing the following:
•
range and rate tests
•
status Matrix check
•
Source/Destination Identifier (SDI) check (except for data from
the EIU)
•
data parity test.
Faults detected by the FADEC are annunciated and recorded for
maintenance or crew action if required.
(2) Digital Outputs
Bit transmission rate is nominally 12.5 kHz. Each channel of the EEC
has two /output ports and each bus has separated line driver.Outputs
are isolated in such a way that propagation of failuresis prevented.
Information contained on FADEC output buses includes the following
general items:
•
Engine Rating Parameter Information
•
Parameters used for Engine Control
•
FADEC System Maintenance Data
•
Engine Condition Monitoring Parameters
•
ECU Status and Fault Indication.
1.7.3 Interfrace
A. ADIRU/EEC Interface
Air Data Inertial Reference Unit (ADIRU) sends air data parameters to
the FADEC through ARINC 429 buses.Each channel of the EEC
receives a digital data stream from both ADIRU which contains total
temperature, total pressure and altitude pressure signals from the
airframe sensing system
B. EIU/EEC Interface
1 EIU input from the EEC
The EIU acquires two ARINC 429 output data buses from the
associated EEC (one from each channel) and it reads data from the
channel in control. When some data are not available on the channel
in control, data from the other channel are used.In the case where EIU
is not able to identify the channel in control, it will assume Channel A
as in control.The EIU looks at particular engine data on the EEC digital
data flow to interface them with other aircraft computers and with
engine cockpit panels.
2 EIU output to the EEC
Through its output ARINC 429 data bus, the EIU transmits data
coming from all the A/C computers which have to communicate with
the EEC, except from ADCs and throttle which communicate directly
with the EEC.There is no data flow during EIU internal test or
initialization.
C. EIU/A/C Interface
The EIU concentrates data from cockpit panels and different aircraft
systems to send them to the FADEC and gives selected FADEC
information to the A/C systems. The EIU communicates with a lot of
A/C systems through analog and digital interfaces.
D. Cockpit System Display/EEC Interface
The aircraft system which processes the engine data and messages
for cockpit display on the cathode ray tubes consists in three display
management
computers
(DMC)
and
two
flight
warning
computers.Each DMC receives 4 engines data buses, one from each
channel of EEC and two from each engine. All the 8 buses from EEC
engine 1 and EEC engine 2 are acquired by the 3 DMCs. Each FWC
recives 4 data buses one from each channel of EEC General
Arrangement.
E. Throttle Control System/EEC Interfaces
The throttle control system is fully electrical and each throttle lever
drives two resolvers ; located in the cockpit center pedestal, these
resolvers are dedicated to the FADEC, one for each engine.The FADEC
excites and demodulates these resolvers.Each throttle lever is fitted
with one pushbutton which is used to generate the autothrust
disconnect discrete signal to the EEC.
F. Thrust Reverser Interface
The EEC controls the deployment and stowing sequence of the thrust
reverser. The logic which is implemented in the EEC is based on TLA
signals, flight ground signals, thrust reverser position feedback.
1.7.4 ECU / Engine Interface
A. 28VDC Power Supply
The EEC is designed to operate with the engine not running, the EEC
is operational five seconds after it is electrically powered by aircraft
28VDC. The EEC is electrically powered by the aircraft through the
EIU. A320 28VDC permits:
•
automatic ground check of the FADEC before engine running
•
engine starting
•
powering the EEC (while engine is running below 10 % N2).
As soon as the engine is running at and above 10 % N2 rpm, the
dedicated alternator provides electrical power for the FADEC system.
B. 115 VAC Power Supply
The 115 VAC power supply is dedicated to the ignitors and to the
P2/T2 probe heating.
1.7.5 Additional Engine Sensor
All interface between ECC and engine sensors, LVDTS feedback and
FMU are additional engine sensor. These additional engine sensors
are optional and dedicated to the engine condition monitoring
through the AIDS. These engine parameters (P12.5,P2.5, T2.5) are
available on the EEC data bus output if installed on the engine.
P12.5 Sensor
P12.5 sensor provides air pressure from the fan exit.
P2.5 Sensor
P2.5 sensor provides air pressure from the LP compressor exit.
T2.5 Sensor
T2.5 sensor provides air pressure at the LP compressor exit.
1.8 INDICATING – DESCRIPTION AND OPERATION
1.8.1 General
The fuel system is monitored from:
•
The fuel flow indication on the upper ECAM display unit
permanently displayed in green and under numerical form.
•
The fuel filter clogging caution (amber) on the lower ECAM
display unit.
1.8.2 Fuel Flow Indication
A. Fuel Flow Indication
The fuel flow transmitter signal is fed to the FADEC which processes it
and transmits the information to the ECAM system for display through
the digital FADEC data bus
B. Fuel Flow Transmitter
(1) Description
The fuel flow transmitter is installed in the fuel line between the fuel
metering unit and the fuel distribution valve. It is mounted on the
lower left-hand side of the fan case, rearward of the LP/HP fuel pump.
The fuel flow transmitter is made of these primary assemblies:
•
the transmitter body,
•
the inlet bearing assembly,
•
the turbine assembly,
•
the measurement assembly.
(2) Operation
Fuel goes into the transmitter and drives the turbine. At low fuel flow
rates all the fuel goes through the turbine. As the fuel flow increase,
the bypass valve starts to open to let some of the fuel go around the
turbine. This prevents the turbine from turning too quickly.
The fuel leaves the turbine and the bypass valve and then all of it
flows through the straightening vanes. These vanes straighten the
fuel flow before it goes into the measurement assembly. When the
fuel flow transmitter is stopped, the
Fig.3.3 Fuel Flow Indication – Diagram
magnets on the impeller align axially with the magnets on the drum.
As the measurement assembly is turned, a pulse is generated each
time a magnet passes its related pick-off coil. When the fuel goes into
the measurement assembly its flow, through the impeller vanes,
resists the movement of the impeller. The spring permits the impeller
to move in relation to the drum. The magnets are then not aligned
and there will be a time difference between the pulses generated in
the drum and impeller pick-off coils. This time difference is directly
proportional to the fuel mass flow rate and is used to calculate the
fuel flow.
1.8.3 Fuel Filter Clogging Indication
A. Fuel Filter Clogging Indication
The fuel filter clog indication is provided on the lower ECAM display
unit. When the pressure loss in the fuel filter exceeds 5 ± 2 psi, the
pressure switch is energized. This causes:
•
The engine page to come on the lower ECAM DU with the
caution signal FUEL CLOG.
•
The associated caution message to come on the upper ECAM
DU.
When the pressure loss in the filter decreases between 0 and -1.5
psid from the filter clog energizing pressure, the pressure switch is
de-energized which causes the caution to go off. The differential
pressure switch signal is fed directly to the SDAC through the
hardware
B. Fuel Filter Differential Pressure Switch
(1) Description
The pressure switch is bolted to the fuel filter housing and connects to
ports in it. The switch is in two housing held together with screws.
One housing contains a bellows and the other a switch. The bellows
housing is connected to the fuel supply with two ports. The bellows is
connected to the filter inlet side and the housing (vent side) is
connected to the filter outlet side. A lever connects the bellows to the
switch. The switch housing contains the switch and an electrical
connector. The switch is lever operated by the bellows lever.
(2) Operation
The bellows and bellows housing filled with fuel at the pressure of the
system. The pressure in the bellows and bellows housing is thus the
same so the bellows do not move. If the filter element gets clogged or
not fully clogged, the filter inlet pressure will increase. This will cause
the bellows to extend. At the pressure set point the bellows will
extend sufficiently far to push the lever and close the switch contacts.
The switch will then transmit a message signal of a clogged filter to
the cockpit.
2
IGNITION
2.1 IGNITION – GENERAL – DESCRIPTION AND OPERATION
2.1.1 General
Fig.4.1 Component Location
The ignition system consists of 2 independent subsystems : 2 high
energy ignition exciters for which the energization is controlled by the
EEC, 2 igniter plugs, and 2 coaxial shielded ignition leads. The
purpose of the system is:
•
to produce an electrical spark to ignite the fuel air/mixture in
the engine combustion chamber during the starting cycle on
ground and in flight.
•
to provide continuous ignition (manual or automatic selection)
during take off, landing and operation in adverse weather
conditions. Continuous ignition will also be automatically
selected when the EIU has failed.
The ignition systems are selected alternately by the EEC (for auto
start only) in order to have no failure on ignition channel for more
than one flight and to increase overall system life. Both ignition
systems are used for manual starts (alternate mode).
2.1.2 Description
The 2 ignition exciters are mounted on the outer surface of the HP
compressor. Each unit has one power input circuit, (4 joules stored)
and a high voltage output circuit to the igniter. The two igniter plugs
are installed on the diffuser case. The igniter plug has 3 sections:
•
the Sparking end with the surface gap
•
the Main body with the plug thread.
•
the Connector to connect the ignition lead
The ignition leads have two conducting paths, one carrying the
current between the exciter and the center electrode of the igniter.
The other providing the return path from the igniter body to the case
of the exciter.
2.1.3 Operation
The ignition exciters operate with 115 V - 400 Hz input. The power is
transformed, rectified and discharged in the form of capacitor
discharge pulses through the ignition leads to the igniter plugs. Each
ignition system can operate independently. The selection of the
system (A or B) is made by the EEC in auto start only. Ignition system
A comprises the upper ignition exciter and its associated cable and
igniter, ignition system B comprises the lower ignition exciter and its
associated cable and igniter.
2.2 ELECTRICAL POWER SUPPLY DESCRIPTION AND
OPERATION
2.2.1 Description
The ignition system requires 115 VAC – 400 HZ power supply. The
power is delivered via the EIU to the FADEC system and associated
relay box. The igniter A is powered from the 115VAC ESS bus 401XP
or from the 115VAC STAT INV BUS 901XP in the emergency
configuration, and the igniter B is powered from normal bus (103XP
for engine 1 and 204XP for engine 2). The availability of this 115VAC
power to relay box is controlled by the EIU according to:
•
The MASTER control switch position: no power is supplied to the
relay box when the MASTER switch is set to OFF.
•
ENG FIRE pushbutton switch position: no power is supplied to
the relay box when the ENG FIRE pushbutton switch is pushed.
A. Ignition Exciters
(1) General
There are 2 ignition exciters which are mounted on the right hand
side of the high pressure compressor front casing. The exciters
provide starting and continuous duty ignition on demand.
(2) Description
The ignition exciter is a capacitor discharge type exciter requiring an input of 115V
(106 to 120 volts AC) at 400Hz (370 to 430 Hz). The output voltage is 22 to 26KV.
2.3 DISTRIBUTION – DESCRIPTION AND OPERATION
2.3.1 Description
The ignition for each engine is carried out by means of one or both
ignition exciters which transform(s) the 115 V - 400 Hz power supply
into high voltage pulsating current. The high voltage flows through
the ignition lead (shielded and ventilated) and delivers to the igniter
plug the power required to initiate the fuel/air mixture combustion by
a series of sparks.
A. Ignition Leads
The air-cooled ignition lead is a part of an ignition system having two
separate channels. The approximate length of the leads is three feet
(‘0’, ‘90’ meter). The air-cooled ignition lead has two conductive
paths: one connects the exciter and the centre electrode of the
igniter plugs, the other is the return path from the igniter body to the
case of the exciter. The air-cooled ignition lead is connected to the
output end of the exciter and to the input end of the igniter plug.
When the exciter discharges the stored energy, the energy goes
through the ignition lead to the igniter plug. Ignition system A
comprises the upper ignition exciter and its associated cable and
igniter, ignition system B comprises the lower ignition exciter and its
associated cable and igniter.
B. Igniter plug
(1) General
The igniter plug is one of the components of the ignition system. Two
igniter plugs are installed on the diffuser case. The igniter plug has
three sections:
•
The sparking end with the surface gap.
•
The main body with the plug thread.
•
The connector to connect the ignition lead.
(2) Operation
When a high voltage pulse from the ignition exciter is delivered to the
igniter plug, the surface gap is ionized and becomes conductive. The
capacitor (in the ignition exciter) discharges the stored energy across
the surface gap. This gives off a spark with high energy at the
sparking end.
2.4 SWITCHING – DESCRIPTION AND OPERATION
2.4.1 General
The ignition system is controlled by: the EEC upon commanded
signals from ENG START panel 115VU through the EIU.
2.4.2 Description
A. ENG Panel (115VU)
Engine start panel is located on the centre pedestal in the cockpit. It
is composed of:
•
ENG/MODE selector switch
It is common to both engines and can be placed in any of the three
positions CRANK, NORM, and IGN/START.
-
CRANK position - In CRANK position no ignition system is
supplied but an engine dry motoring is allowed.
-
NORM position - This position is selected by the pilot at
the end of starting sequence or after engine shutdown on
ground. In this position the EEC selects automatically the
continuous ignition in some specific configurations:
1. The engine is running and the air intake cowl antiicing is selected to ON or the EIU controlled ignition
is failed or during take-off or during flexible take off
or when the approach idle has been selected.
2. In flight, when there is an engine flameout or stall.
-
IGN/START position, This position is selected for:
1. normal starting procedure (automatic)
2. alternate starting procedure (manual)
3. continuous ignition after starting sequence
•
Two ENG/MASTER control switches
There is a MASTER control switch for each engine.
With the MASTER control switch in OFF position, the HP fuel shut off
valve is closed, the engine is stopped. This position of the MASTER
control switch overrides any EECfunction.
•
The MASTER control switch in ON position enables:
-
normal starting procedure (automatic)
-
alternate starting procedure (manual)
-
wet crank procedure
-
Normal operation.
•
Two FAULT legends on the ENG1 and 2 annunciators one amber
FAULT legend dedicated to each engine is supplied by the EIU
when:
-
a starting failure is detected in AUTO MODE
-
Or a disagreement occurs between the HP fuel shut off valve
actual position and the commanded position.
•
Two FIRE legends on the ENG1 and 2 annunciators one red FIRE
legend dedicated to each engine is supplied by the FDU (Fire
Detection Unit) when a fire occurs.
B. ENG/MAN START Section of Panel 22VU
It is located on the overhead panel in the cockpit. There are two
pushbutton switches, one for each engine. Each pushbutton switch
serves for the manual starting procedure only.
2.5 IGNITION STARTING AND CONTINOUS RELIGHT –
DESCRIPTION AND OPERATION
2.5.1 Description
The ignition circuit is supplied with 115 VAC – 400 Hz. The electrical
power is supplied via the EEC and EIU which control the ignition of the
igniter plugs. A dormant failure of an ignition exciter is not possible
for more than one flight because:
•
the two ignition systems are independent
•
The EEC selects alternately ignition system A or B.
Ignition system A comprises the upper ignition exciter and its
associated cable and igniter, ignition system B comprises the lower
ignition exciter and its associated cable and igniter.
2.5.2 Operation
Fig.4.3 Ignition and Starting System
A. Automatic Start Sequence
When an automatic start sequence has been activated by the EEC
(ENG/MODE selector switch in IGN/START position and MASTER
control switch to ON) an automatic dry crank sequence of 50 seconds
is performed; then, the EEC energizes automatically the appropriate
ignition exciter and keeps it energized until N2 reaches 43%. For in-
flight restart the EEC selects simultaneously both ignition exciters. On
the ground, after engine start, the selector must be placed in NORM
position, then back to IGN/START to select continuous ignition. In
flight after engine restart, if the selector is maintained in IGN/START
position, the EEC selects the continuous ignition on the corresponding
engine. In case of incident during an automatic starting on the
ground, the EEC aborts automatically the sequence by closing the
starter shut-off valve and the HP fuel shut-off valve and deenergizing
the igniters.
B. Alternate Start Sequence
When a manual start sequence has been activated by the EEC
(ENG/MODE selector switch in IGN/START position and the ENG/MAN
START pushbutton switch selected to ON) the EEC opens the starter
valve and a dry crank sequence of 50s is performed. Then the
MASTER control switch is placed in ON position (the EEC opens the HP
fuel shut off valve and energizes the ignition exciters). The
deenergization of the ignition exciters is automatically commanded
by the EEC. Positioning of the MASTER control switch to ‘OFF’ during
that starting sequence, results in ignition exciter deenergization.
C. Continuous Ignition
(1) Manual Selection
When the engines are running on the ground or in flight the
continuous ignition is obtained by positioning the ENG/MODE selector
switch in IGN/START position.
(2) Automatic selection
The EEC automatically selects continuous ignition when:
The ENG/MODE selector switch is turned to IGN/START while the
•
engines are at or above idle,
•
flameout is detected,
•
takeoff (determined by TRA) or flex takeoff is performed,
•
engine anti-ice is selected,
•
EIU fails (except during cranking),
•
approach idle is selected,
•
Master Lever is inadvertently cycled from ON to OFF then back
to ON position.
D. Igniter Plug Test
The operation of the igniter plugs can be checked on the ground,
engine not running, through the maintenance MENU mode of the
FADEC. The test will be performed by selecting the corresponding
IGNITOR TEST page in the MENU and positioning the MASTER control
switch to ON to have the 115VAC power supply on the relevant
engine.
AIR
3
3.1 AIR – GENERAL – DESCRIPTION AND OPERATION
3.1.1 General
The air system covers primary, secondary (bypass) and parasitic
(cooling and pressurizing) airflows and the systems used to control
the airflow. It is composed in 2 major sections.
A. Engine Section
The airstream flowing through the IAEV2500 turbofan engine supplies
2 majors systems:
•
The internal air system, which consists of the following
subsystems:
-
Propulsion airflow (secondary and primary flows).
-
Bearing compartments pressurizing air.
-
Cooling air.
•
The external air system, which consists of the following
subsystems:
-
HP/LP turbine active clearance control.
-
High-energy igniter harness cooling air.
-
Engine bleed air.
B. Nacelle Section
The nacelle installation is designed to provide cooling and ventilation
air for engine accessories mounted along the fan and core casing. The
distribution and circulation of the air in the components is such that
the temperature limit for specific components is not exceeded.
3.1.2 Description
All engine air enters the front mounted fan through the engine air
intake cowl. After being compressed by the fan, the airflow is divided
into primary and secondary (bypass) airflows by the flow splitter in
the fan frame.
A. Propulsion Airflow System
(1) Secondary flow
Fan air passes through the Outlet Guide Vanes (OGV) and the fan
frame struts. Bypass air is discharged through the Common Nozzle
Assembly (CNA) during normal engine functioning and provides the
major portion of engine thrust (approximately 4/5 of the total airflow
of the engine). When the thrust reverser is deployed, the bypass air is
directed outward through the thrust reverser cascades to provide
reverse thrust.A small portion of the bypass air is used for HP/LP
turbine active clearance control and environmental control system
cooling through the precooler.
(2) Primary flow
A portion of fan air passes into the 3-stage booster and enters the
core by a converging duct formed by the fan frame. This duct is
provided with Booster Stage Bleed Valves (BSBV). The air then enters
the HP compressor which is provided with 7th and 10th stage bleed
valves (required for engine stability during starting and transient
conditions).The compressed air enters the combustion chamber and
is ignited with the fuel. The exhaust gases flow through the high
pressure turbine (HPT) and the low pressure turbine (LPT) and are
discharged through the Common Nozzle Assembly (CNA).
B. Bearing Compartments Pressurizing Air
(1) The bearing compartment pressure, to ensure satisfactory sealing,
is obtained from the 6th stage compressor manifold for the N⁰_s 1, 2
and 3 bearings.
C. Engine Internal Cooling Air System
The engine internal cooling air system consists of two subsystems:
(1) The HP turbine cooling controlled air system which uses 10th
stage HP compressor bleed air to cool the stage 2 turbine blades,
both HP turbine disk bores and the LP turbine cavity, in response to
EEC command.
(2) The HP turbine cooling air system which uses 10th stage HP
compressor bleed air to cool the HP turbine case, the LP turbine
support rails for the diffuser duct outer segments and the stage 2
vanes.
D. HP/LP Turbine Active Clearance Control
The HP and LP turbines are cooled by fan air drawn from a common
HP/LP ACC air scoop in the fan duct and distributed to both turbine
casings.
E. High Energy Igniter Harness Cooling Air
The ignition system is cooled by fan air which is directed to the
exciter, the lead and igniter plug.
F. Engine Bleed Air
Two customer bleeds are available at stages 7 and 10 of the HPC.
(1) The air intake cowl anti-icing system consists of ducting routing
from a 7th stage engine dedicated bleed port to the air intake cowl.
An on-off valve controls the air supply to the air intake lip.
(2) The environmental control system (ECS) pneumatic installation
collects bleed air from either the engine 7th stage manifold or the
engine 10th stage manifold and delivers bleed air through a pressure
regulating valve to the pylon/nacelle assembly interface
(3) ECS air cooling is provided through the precooler by air taken in
the fan discharge.
3.1.3 Nacelle Temperature indication
This system enables the nacelle core zone ambient temperature
indication to
be displayed.
3.2 COOLING – DESCRIPTION AND OPERATION
3.2.1 General
The power plant cooling system consists of:
•
cooling of the nacelle compartments
•
cooling of the aircraft and engine accessories
•
cooling of engine parts (HPT, LPT)
A. Nacelle Compartment Cooling
The nacelle is divided in three major areas
•
the engine air inlet
•
fan compartment
•
core compartment.
The last two compartments only are cooled to fulfil the following
functions:
•
Sufficient airflow to offset the effects of engine case heat
rejection and engine flange air leakage thereby maintaining an
acceptable compartment temperature level.
•
Cooling of temperature critical components.
•
Cowling pressure load limiting in the event of pneumatic duct
failures.
•
Ventilation of compartment during engine shutdown.
•
Ventilation of combustible fluid vapors to preclude fires.
B. Aircraft and Engine Accessories Cooling
The nacelle installation is designed to provide cooling and ventilation
air for engine accessories mounted on the fan and core casing. The
distribution and circulation of the air in the compartments is such that
the limit for specific components is not exceeded.
C. Cooling of Engine Parts
Differents parts of the IAE V2500 engine are cooled by air bled in the
primary flow (HP turbine cooling), and secondary flow (HPT/LPT ACC).
3.2.2 Component Location
A. Location of HP Turbine Cooling System
Fig.5.3 HP Turbine Cooling System Location
B. Location of HP/LP Turbine Active Clearance Control System
Fig.5.4 HPT / LPT Active Clearance Control System Location
3.3 NACELLE COMPARTMENT AND ACCESSORY COOLING –
DESCRIPTION AND O[ERATION
3.3.1 General
The nacelle compartment and engine accessories are air cooled .The
cooling air is taken from the air flowing in and around the nacelle
cowls. There are three cooling systems:
•
Fan and core compartments.
•
Engine gearbox breather vent.
•
Air cooled oil cooler vent.
The nacelle cooling and ventilation systems provide the following
functions:
•
Sufficient airflow to offset the effects of engine case heat
rejection and engine flange air leakage, thereby maintaining an
acceptable compartment temperature level.
•
Cooling of temperature critical components.
•
Ventilation of compartment during engine shutdown.
•
Ventilation of combustible fluid vapours to preclude fire.
3.3.2 Description
A. Fan and Core Compartments Cooling
(1) Fan case compartment accessories are cooled by air which enters
through a scoop in the air intake cowl. A duct from this scoop goes to
a Y-shaped outlet duct on the cowl aft bulkhead. The air comes out of
this duct into the fan compartment.
(2) The air in the fan compartment is vented overboard through two
outlet vents in the bottom of the fan cowl (one in each fan cowl door).
(3) Core compartment ventilation is provided by fan air through holes
in the inner wall of the
c
ducts. Air circulates through the core
compartment and exits through the exhaust orifice located in the
lower bifurcation of the C ducts. This is supplemented by air
exhausting from Active Clearance Control System around the turbine
area.
Fig.5.5 Fan Case compartments cooling
(4) During ground running local pockets of natural convection provide
some ventilation of the fan case zone.
(5) The fan compartment is sealed to keep the air in. Seals are
installed at the following locations:
•
Air intake cowl to fan cowl junction.
•
Fan cowl door split line.
•
Fan cowl to thrust reverser C-ducts junction.
•
Gearbox breather to right fan cowl junction.
•
ACOC outlet vent to right fan cowl junction.
(6) The maximum permitted air temperatures are:
•
230 degrees F (110 degrees C) at the top of the fan
compartment
•
255 degrees F (124 degrees C) at the bottom of the fan
compartment.
•
797 degrees F (425 degrees C) in the core compartment.
B. Engine Gearbox Breather Vent System
The engine gearbox breather lets the gas from the gearbox go
overboard from the nacelle. The gas goes through a breather duct on
the gearbox and then a duct in the right hand side fan cowl. A seal is
installed at duct junction.
C. Air Cooled Oil Cooler Vent System
The air cooled oil cooler (ACOC) is cooled with air from the fan
duct. The hot air goes overboard from the cooler through an outlet
grille in the right fan cowl.
3.4 BEARING COMPARTMENT COOLING AND SEALING –
DESCRIPTION
3.4.1 General
A. The engine main bearings are contained in three bearings
compartments:
•
The front bearing compartment.
•
The N⁰4 bearing compartment.
•
The rear bearing compartment.
B. Each compartment has seals installed to prevent oil leakage. The
seals are pressurized by air taken from stage 6 and 8 of the
compressor. Compressor air is also used to keep the N⁰4 bearing
compartment cool.
3.4.2 Description
A. Description
The front bearing compartment contains three bearings, No. 1 ball
bearing and No. 2 roller bearing for the low spool shaft and No. 3 ball
bearing for the high spool shaft. The wall of the front bearing
compartment is made up of the inner wall of intermediate case and
the No. 1 bearing support. The compartment is sealed against the
high spool shaft and the low spool shaft with two brush seals, two
carbon seals and a hydraulic seal. The combination of a brush seal
and a carbon seal in front of No. 1 bearing seals the compartment
against the low spool shaft. The hydraulic inter-shaft seal is used to
seal the compartment against 8th stage compressor bleed air in the
annulus between the high and low rotor shaft. The combination of a
brush seal and a carbon seal rear of No. 3 bearing seals the
compartment against the high spool shaft.
B. Operation
Pressurizing air from 6th stage compressor goes through two tubes to
the space in front of No. 1 bearing and between the brush seal and
the carbon seal. Pressurizing air to the space rear of No. 3 bearing
and between the brush seal and the carbon seal is supplied through a
routing in case casting from 6th stage compressor. Air in the
compartment is vented to a de-oiler to keep proper seal differential
pressure. A restrictor in the venting line controls the air flow to
prevent air leak at the seals
3.4.3 No. 4 Bearing compartment cooling and sealing
A. General
The No. 4 bearing compartment is cooled by 12th stage air. An
external plumbing carries this air from a single diffuser case port
through an air cooled air cooler and back to the diffuser case at three
locations. Internal diffuser case plumbing carries the cooled air to the
No. 4 bearing compartment where it is distributed between the
compartment walls forming a thermal barrier. The air exhausts
through holes into the front and rear annuli formed by the HP rotor
shaft and the bearing compartment. This air in the annuli flows in two
directions. Some air flows past the front and rear carbon seals into
the bearing compartment. The remainder flows into the diffuser case
inner cavity. This system prevents ingestion of hot 12th stage
compressor air in the diffuser case inner cavity from entering the
bearing compartment should a carbon seal fail. Before entering the
No. 4 bearing compartment the 12th stage air passes through an air
cooled air cooler to be cooled.
3.4.4 Air Cooled Air Cooler
A. Description
The No. 4 bearing compartment air cooler is installed on the turbine
casing. The matrix of cooler is made of 283 dimpled hairpin stainless
steel tubes, rounded to match the fan air routing radius, vacuumfurnace brazed to a stainless steel tube sheet and contained within a
stainless steel casing. Upper and lower casing side plates are
strengthened by five (5) attached baffles, through which the tubes
may move freely to agree with thermal expansion. The exchanger is
held by its coolant air duct flanges.
B. Operation
Fan air goes into the No. 4 bearing compartment air cooler and makes
a single pass over the tubes. Bleed air from 12th stage compressor
goes through tubes
3.5 HP TURBINE COOLING – DESCRIPTION AND
OPERATION
3.5.1 General
A. HP Turbine Cooling Make-up Air System
Fig.5.8 HP Turbine Cooling Controlled Air System
A cooling system is provided to supply supplemental air to cool the
stage 2 turbine blades and HP 1 turbine disk bores during all power
settings except cruise. The source of this air is 10th stage compressor
bleed. This air is supplied to the diffuser case through the 10th stage
to HP turbine air valve mounted on the HP compressor case and two
external pipes which connect to the valve. The pipes pass through
two diffuser case struts into the diffuser case internal cavity. Here the
air mixes with air from various sources and continues through internal
cavities to the stage two blades and HP 1 turbine disk bores. At cruise
power settings the valve in this system is shut off to improve engine
performance.
3.5.2 Description
(1) General
The HP turbine cooling controlled air system consists of:
•
a control valve
•
a valve solenoid (controlled by EEC)
•
2 tubes
(2) Control Valve
The external piping incorporates a cooling air control valve which is
either fully closed or fully open. The valve is operated by a solenoid
valve controlled in response to EEC signals. The control valve is
normally open. A visual position indicator on the valve is provided for
maintenance purposes. It also incorporates 2 position indication
switches to provide a signal to the EEC (channels A and B) for fault
detection purposes.
(3) Solenoid valve
The solenoid valve is mounted on the fan case. It is energized/deenergized in response to EEC signals. When energized the solenoid
supplies high pressure compressor discharge air to the control valve
which shuts-off the flow of cooling air. De-energizing the solenoid
opens the control valve allowing full cooling air flow. Scheduling by
the EEC is a function of high rotor speed and altitude.
3.6 HP / LP TURBINE ACTIVE CLEARENCE CONTROL (ACC)
SYSTEM – DESCRIPTION AND OPERATION
3.6.1 General
The HP/LP Turbine Active Clearance Control (ACC) system uses fan air
to cool the HP and LP cases for blade tip clearance control in order to
improve engine performance and maximize the turbine cases life
time. Fan air is drawn from a common HP/LP turbine ACC air scoop in
the fan duct. This air is divided into HP and LP cooling air and passes
through individual short ducts to the Active Clearance Control Valves
which direct air for both HP and LP turbine case cooling.
3.6.2 Description
The LP and HP Turbine ACC Systems function in the same manner.
Cooling air passes through the tube perforations, located on the inner
surfaces of the tubes, and impinges on the case outer surface. The
cool air reduces the temperature and the thermal expansion of the
cases under some operating conditions. The LPT/HPT ACC system
consists of the following components:
•
one HP turbine ACC manifold
•
two LP turbine ACC manifolds
•
Four HPT ACC perforated tubes
•
height LPT ACC perforated tube
•
one LP turbine supply pipe
•
one LP/HP turbine valve (composed of a HPT control valve and a
LPT control valve)
•
one LP/HP turbine ACC actuator
3.6.3 Operation
A. Normal Operation
The torque motor of the ACC actuator has springs to hold the jet pipe
at neutral bias position. The torque motor current from the EEC
changes the jet pipe direction. This direction change of the jet pipe
makes a higher pressure on one side of the servo valve. This higher
pressure moves the servo valve and it changes the fuel passage to
the actuator piston. The pressure difference between two sides of the
actuator piston and the spring load moves the actuator piston to the
required position. The actuator piston drives the butterfly valves in
the ACC valve. The butterfly valves thus control the airflows for the
HP/LP turbine ACC manifolds. The LVDT_s transmit the valve position
to the EEC.
B. Fail Safe Operation
When there is no torque motor current or no fuel servo pressure, the
actuator piston moves to the point A. The actuator piston remains at
this point at all defective actions.
C. Mode of Operation
The butterfly valves of the ACC valve are driven by the same actuator
in the modes given below
(1) Engine stop
With the engine stop, the position of the actuator piston is at point A
At this point: The butterfly valve for the HP turbine ACC is closed. The
butterfly valve for the LP turbine ACC is not less than 45 per cent
opened.
(2) Engine operation
During engine operation, the EEC controls the position of the actuator
piston between point B and point E.
(3) Take-off
During take-off, the position of the actuator piston is at point C At this
point: The butterfly valve for the HP for the HP turbine ACC is closed
The butterfly valve for the LP turbine ACC is not less than 70 per cent
opened.
3.7 IGNITION SYSTEM COOLING – DESCRIPTION AND
OPERATION
3.7.1 General
The ignition cooling system uses air from the fan stream to cool the
ignition exciter, the lead and igniter plug. Fan air is directed to each
exciter through a flexible conduit where it enters the exciter cooling
box. Cooling air passes over the exciter and exits through a connector
cooling jacket surrounding the ignition lead connector and enters a
flexible conduit which surrounds the lead. Cooling air then flows
around the lead to the igniter plug and excits through the igniter
cooling jacket.
3.8 COMPRESSOR CONTROL – DESCRIPTION AND OPERATION
3.8.1 General
The booster stage bleed valve, the variable stator vane and HP
compressor bleed valves systems are controlled by the EEC. The
booster stage bleed valve controls the LP compressor airflow. The
variable stator vane and the 7th and 10th stage bleed valves control
the HP compressor airflow.
3.8.2 Variable Stator Vane System (VSV)
The VSV system maintains a satisfactory compressor performance
over a wide range of operating conditions. The system varies the
angle of the inlet guide vanes and stator vanes to aerodynamically
match the low pressure stages of compression with the high pressure
stages. This variation of vane position changes the effective angle at
which the air flows across the compressor blades and vanes. The VSV
angle determines the compression characteristics (direction and
velocity) for any particular stage at compression.
3.8.3 Booster Stage Bleed Valve
At low LP spool speeds the booster provides more air than the core
engine can utilize. To match the booster discharge airflow to the core
engine requirements at low speed, excess air is bled off through
booster stage bleed valves (BSBV) into the fan discharge air stream.
At higher engine speeds the BSBV are closed so that all the booster
discharge (primary air flow) enters the core engine.
3.8.4 HP Compressor Bleed Valves
The 7th and 10th stage bleed valves maintain a more stable
operation of the
Compressor .
3.9 LP COMPRESSOR AIRFLOW CONTROL SYSTEM –
DESCRIPTION AND OPERATION
3.9.1 General
The primary function of the LP compressor airflow control system is to
control the airflow thus ensuring compressor stable operation during:
•
Engine start.
•
Engine transient operation.
3.9.2 Description
A. General
the airflow control system includes:
•
An LPC bleed-slave actuator
•
An LPC bleed-master actuator
•
Two bleed-valve actuating rods
•
A booster bleed valve and actuating mechanism The airflow
control system automatically operates to control the air bled
from the LP compressor.
The two actuators are mechanically attached to each actuating rod
and, the bleed-valve and actuating mechanism. The two actuators are
connected hydraulically and operate together by command and
feedback signals from/ to the EEC.
B. LPC Bleed-Master and Slave Actuators
Description
•
The two actuators are hydraulically connected units which
operate together. Each of them is attached to a bracket on the
intermediate structure.
•
The two bleed-valve actuating rods connect each actuator fork
end to the synchronizing ring. The force to move the
synchronizing and its linkage through the actuating rods, is
supplied from the two actuators.
•
The master actuator gives an accurate control in relation to an
electrical input signal (command signal) from the EEC. The
master actuator also gives an electrical output signal (the
positions of the actuator piston fork end) to an electrical input
signal (command signal) from the EEC. The master actuator
also gives an electrical output signal (the positions of the
actuator piston fork end) to the EEC.
Operation
the two actuators open or close in relation to the input signal from the
EEC.
C. Booster Bleed Valve and Actuating Mechanism
Description
•
The bleed valve and actuating mechanism operates to make
each bleed valve synchronized, in relation to the positions of
the two actuators.
•
The bleed valve and actuating mechanism is a sub-assembly
which includes:
•
27 bleed valves which are attached to the support ring.
•
27 links which connect the bleed valves to the 25 link arms and
the two power arms.
•
the synchronizing ring which is connected to the 25 link arms
and the two power arms.
The two actuating rods connect the two power arms to the two
actuators.
Operation
The 27 bleed valves open during the engine start and other scheduled
flight operations. This permits some of the LP compressor airflow
(which normally go into the HP compressor) to go through the fan
airstream. The bleed valves are closed during takeoff and cruise
operations.
3.10 HP COMPRESSOR AIRFLOW CONTROL SYSTEM –
DESCRIPTION AND OPERATION
3.10.1General
This system consists of the Variable Stator Vane Actuator, the HPC
bleed air tubes and the HPC stage 7 and 10 solenoid valves and bleed
valves. The variable stator vanes and the 7th and 10 t stage HP
compressor bleeds are used to improve engine performance and
stability during engine starting, acceleration and deceleration. The
variable stator vanes are modulated open or closed by a hydraulic
actuator controlled by the Electronic Engine Control (EEC). The 7th
and 10th HP compressor bleed valves are pneumatically actuated by
two position air shut-off valves which are actuated by EEC solenoids.
3.10.2Description
A. Variable Stator Vane System
(1) General
The purpose of this system is to position the Inlet Guide Vanes (IGV)
and stator vanes, using a fuel driven hydraulic actuator, in response
to electrical signals provided by the EEC.
(2) Description
The stator vane actuator accurately controls vane movement with
respect to a torque motor current supplied by the EEC. Operation of
the stator vanes in regulated by accurate control of high pressure fuel
flow to one or other side of a differential area piston. The piston has
an externally adjustable low speed stop at the extended end of its
travel. The high speed stop is formed by a collar which limits piston
retraction. Provision is made to lock the piston with a rigging pin for
setting purposes. A control servo valve and piston type pressure drop
regulator regulate the flow to either side of the piston.
B. HP Compressor Bleed Valves System
(1) General
HP compressor bleed valves are required for stability of engine
operation during starting and transient conditions. Three bleed valves
for the seventh stage and one for the tenth stage of the HP
compressor redirect a portion of HP compressor air into the fan
stream.
(2) Description/Operation
The bleed valves are mounted onto the outside of the HP compressor
case. Each valve has a spring loaded open position. The piston has
holes to permit a flow of air through the valve when it is open. When
the valve is closed the piston is held against a plate which is part of
the valve body. This shuts off the air bleed from the compressor. The
piston is held in the open position by spring and compressor delivery
(P3) pressures. When the P3 pressure is released by the controlling
solenoid valve, the bleed valve is closed by compressor air pressure.
3.11 NACELLE TEMPERATURE INDICATING – DESCRIPTION
AND OPERATION
3.11.1General
A nacelle temperature probe provides indication of nacelle core zone
ventilation air exit temperature. It indicates over temperature
resulting from loose or broken air ducts or from loose flanges, worn
VSV bushings etc. In order to allow early maintenance corrective
action.
3.11.2Description
The nacelle temperature indicating system is composed of a probe
and an indicator on the ECAM. The nacelle temperature probe has a
measurement range of -54ºC to 330 ºC (-65.2 ºF to 626 ºF). The signal
is fed to the EIU which transforms the analog information into digital
form. Then the EIU transmits the data to the ECAM system. When the
value reaches 320 ºC the indication is displayed and flashes (greenadvisory). Below 320 ºC the nacelle temperature indication is not
displayed.
4
ENGINE CONTROL
4.1 ENGINE CONTROL – DESCRIPTION AND OPERATION
4.1.1 General
The engine control system consist of:
•
the throttle control
•
the HP fuel shut-off valve control
•
the LP fuel shut-off valve control
A. Throttle Control Lever
The throttle control system is fully electrical. The throttle control lever
drives several position detectors. The position detectors are located
under the cockpit center pedestal. Two of them are resolvers
dedicated to the FADEC system. Each channel of the Electronic
Engine Control receives the position signal from one resolver in the
analog form.
B. ENG/MASTER Control Switch
The MASTER control switch interfaces with the HP fuel shut off valve
and the Engine Interface Unit (EIU). It controls directly the closure of
the fuel shut off valve when it energizes the valve solenoid. In parallel
an analog signal informs the EIU of the position of the MASTER control
switch. The EIU transmits then the information to the EEC in digital
form. The EEC uses this information in addition to the ENG/MODE
selector switch position and manual start pushbutton switch to control
the starting sequence
C. LP Fuel Shut-Off Valve Control
The LP fuel shut-off valve is electrically controlled:
•
either by the ENG/MASTER control switch on the ENG panel
•
or by the ENG FIRE pushbutton switch
4.2 POWER CONTROL – DESCRIPTION AND OPERATION
4.2.1 General
A. EPR Mode
(1) Manual thrust setting mode (the autothrust function is not active)
The throttle control lever in the cockpit center pedestal controls:
•
the actual engine thrust setting
•
the selection of the thrust limit mode
The EEC:
•
receives the throttle lever angle
•
computes the engine thrust limit parameter and
•
adjusts the actual engine thrust parameter accordingly.
(2) Automatic thrust setting mode (the autothrust function is active)
The throttle control lever controls:
•
the limitation of the actual thrust
•
the selection of the thrust limit mode.
The FMGC computes the target thrust parameter and sends it to the
EEC.
The EEC adjusts the actual engine thrust parameter according to this
Value
The top limit value of the target thrust parameter is the actual
position of the throttle control lever (this is applicable when the alpha
floor protection is not activated). When the alpha floor protection is
activated the throttle control lever represents the lower limit of the
autothrust demand.
B. N1 Reversionary Mode
When N1 reversionary mode is selected, the automatic thrust setting
mode (autothrust active) cannot be activated. The throttle control
lever in the cockpit center pedestal controls the actual thrust
setting.The EEC receives the throttle lever angle and adjusts the
actual engine thrust parameter according to ambient conditions.
4.3 THROTTLE CONTROL – DESCRIPTION AND OPERATION
4.3.1 General
The throttle control system consist of:
•
the throttle control lever
•
the throttle control artificial feel unit
•
the throttle control unit
•
the electrical harness
In addition each throttle control lever is fitted with an instinctive
disconnect pushbutton switch.This pushbutton switch is used for the
de-activation and disengagement of the autothrust function. The
design of the throttle control is based upon a fixed throttle concept:
this means that the throttle control levers are not servo motorized.
4.3.2 Throttle Control Mechanical Parts
A. Throttle Control Lever
The throttle control handle comprises:
•
a throttle control lever which incorporates stop devices
•
a graduated fixed sector
•
and autothrust instinctive disconnect pushbutton switch
•
a reverse latching lever.
The throttle control lever moves over a range of 65 º from:
•
20 º TLA (Full Reverser Throttle stop) to 45 º TLA (Full Forward
Throttle stop).
An intermediate mechanical stop is fitted at 0 ºTLA. This stop is
overriden when the reverse latching lever is pulled up for selection of
the reverse power.
This idle stop is reset as soon as the lever is selected back to forward
thrust area. In the forward thrust area, there are two detent points:
•
MAX CLIMB detent point set at 25 º TLA
•
MAX CONTINUOUS/FLEX TAKE OFF detent point set at 35 º TLA.
In the reverse thrust throttle range, there is one detent point set at -6
º TLA : this corresponds to the selection of the thrust reverser
command and the Reverse Idle setting From -20 deg. TLA to 0 deg.
TLA, the autothrust function cannot be activated. In the middle
throttle range (0 deg. to 35 deg. TLA), the autothrust function can be
active if engaged. This range corresponds to the selection of MAX
CLIMB or MAX CONTINUOUS thrust limit mode, except the FLEX TAKE
OFF mode. If the auto thrust is not engaged, the engine control is
manual. In the forward range (35 deg. to 45 deg. TLA), the auto thrust
function cannot be activated (except in alpha floor condition). The
range over 35 deg. TLA corresponds to the selection of MAX TAKE
OFF/ GO AROUND mode. The throttle control lever is linked to a
mechanical rod. This rod drives the input lever of the throttle control
artificial feel unit.
B. Throttle Control Artificial Feel Unit
The Throttle control artificial feel unit is located below the cockpit
center pedestal. This artificial feel unit is connected to engine 1/2
throttle control lever and to the engine 1/2 throttle control unit by
means of rods. The artificial feel unit is a friction system which
provides a load feedback to the throttle control lever. This artificial
feel unit comprises two symmetrical casings, one left and one right.
Each casing contains an identical and independent mechanism. Each
mechanism is composed of:
•
a friction brake assembly
•
a gear assembly
•
a lever assembly
•
a bell crank assembly
Throttle lever travel is transmitted to the to the artificial feel unit and
to the throttle control unit. The linear movement of the throttle levers
is transformed into a rotary movement at the bell crank which turns
about the friction brake assembly shaft. This movement rotates a
toothed quadrant integral with the shaft. This toothed quadrant
causes inverse rotation of a gear equipped with a disk which has four
detent notches. Each notch corresponds to a throttle lever setting and
is felt as a friction point at the throttle levers.
C. Throttle Control Unit
A mechanical rod transmits the throttle control lever movement. It
connects the throttle control artificial feel unit to the input lever of the
throttle control unit. The throttle control unit comprises:
•
an input lever
•
mechanical stops which limit the angular range
•
2 resolvers whose signals are dedicated to the EEC (one
resolver per channel of the EEC)
•
6 potentiometers fitted three by three. Their signals are used
by the flight control system
•
a device which drives the resolver and the potentiometer
•
a pin device for rigging the resolvers and potentiometers
•
a safety device which leads the resolvers outside the normal
operating range in case of failure of the driving device
•
two output electrical connectors.
D. Electrical Harness
There is one independent electrical harness dedicated to each TLA
signal with its associated output connector. The electrical harness is a
triplex twisted shielded pair of wires:
•
one pair for sine
•
one pair for cosine
•
one pair for excitation.
4.4 ENGINE MASTER CONTROL – DESCRIPTION AND
OPERATION
4.4.1 HP Fuel Shut OFF Valve Control
A. General
The HP fuel shut off valve control is fully electrical. It is performed
from the engine panel in the cockpit as follows:
•
opening of the HP fuel SOV: This is controlled by the EEC : the
EEC receives the commands from the MASTER control switch
and ignition selector switch.
•
closure of the HP fuel SOV: This is controlled directly from the
MASTER control switch in OFF position. The MASTER control
switches are part of the panel 115VU.
B. Valve Operation
The aircraft provides 28 Volt DC power to the EEC for operation of the
engine HP fuel shut off valve latching torque motor. Two coils of the
latching torque motor are operated by the EEC (1 per channel). Each
coil provides the ON and OFF function depending upon the current
polarity sent from the EEC. One additional coil is controlled by the
hardwired fuel-off input from the cockpit.
4.4.2 LP Fuel Shut Off Valve Control
A. General
The function of the LP fuel shut-off valve is to control the fuel supply
at engine-to-pylon interface. The valve is located on the engine fuel
supply system in the wing leading edge.
B. Valve Operation
The LP fuel shut off valve is controlled:
•
From the cockpit overhead panel 20VU by means of the ENG
FIRE pushbutton switch
•
From the cockpit center pedestal 115VU by means of the
MASTER control switch on, the ENG panel.
5
ENGINE INDICATING
5.1 ENGINE INDICATING – GENERAL – DESCRIPTION AND
OPERATION
5.1.1 General
The engine power management indicating is performed by means of:
•
the ECAM system (upper and lower display units)
•
the warning and caution systems.
The engine monitoring is carried out by means of:
•
the EEC and the ECAM
•
the vibration monitoring system with a display on the ECAM.
This system description note deals only with:
•
the engine management main parameters (N1, EPR LIMIT, EPR
THROTTLE, EPR MAX, EPR REF)
•
the engine vibration monitoring system.
5.1.2 Indicating
A. Upper ECAM Display Unit
•
N1
•
EPR LIMIT
•
EPR THROTTLE
•
EPR MAX
•
EPR REF
•
EGT
•
N2
B. Lower ECAM Display Unit
The VIB indication is normally green. If the advisory level is reached,
the indication flashes (0.6 s bright and 0.3 s normal).
5.2 POWER – DESCRIPTION AND OPERATION
5.2.1 General
This chapter deals with the N1, N2 and EPR indications. The other
engine indications related to the power management parameters
5.2.2 N1 Indicating System
The measurement channel for the low-pressure rotor speed is
designed as follows:
•
the speed sensor on the engine sends a signal to the EEC,
•
the EEC uses this signal in engine control computation and also
transmits it to the ECAM through the ARINC 429 data bus
5.2.3 N2 Indicating System
The N2 indicating system provides the signals proportional to the
High Pressure (HP) shaft rotational speed to the EEC for use in engine
control computation, to the ECAM for visual display in the cockpit and
to the Engine Vibration Monitoring Unit for use in processing engine
vibration data.
5.2.4 EPR (Engine Pressure Ratio) Indicating System
The Engine Pressure Ratio (EPR) indicating system consists of one
combined P2/T2 sensor and eight pressure ports located in three LP
turbine exhaust case struts (24 total). The pressures from these
sensors are routed to the EEC pressure transducers. The signals are
then converted into a digital format. The Electronic Engine Control
(EEC) processes the pressures to form actual EPR (P4.9/P2) and
transmits the EPR value to the ECAM system through it digital data
bus. Each of the two EEC channels performs this operation
independently.
5.3 TEMPERATURE – DESCRIPTION AND OPERATION
5.3.1 General
The measurement channel for the exhaust gas temperature is
designed as follows:
•
Four probe assemblies each comprising 2 thermocouples.
•
Four thermocouples (one from each probe assembly) are used
to form an averaged signal sent to the channel A of the EEC.
•
The remaining four thermocouples (one from each probe
assembly are used to form an averaged signal sent to the
channel B of the EEC.
•
The EEC uses the Exhaust Gas Temperature in the engine start
control logic and also transmits the EGT signal to the ECAM
through the ARINC bus.
5.3.2 EGT Indication
Four thermocouple probe assemblies are located at station 4.95 (LP
turbine exhaust case strut) at 9.5, 7.5, 4.5 and 2 o’clock viewed from
the rear.
5.3.3 Indication
The EGT indication appears on the upper ECAM display unit.
The ECAM provides the EGT indication:
•
in analog form with a pointer which deflects in front of a dial,
•
in digital form, in the lower section of the dial.
The indication is normally green. When : 610 deg.C < EGT < 635
deg.C:
•
the indication becomes amber
•
the MASTER CAUT light comes on accompanied by the single
chime
•
the following message appears on the ECAM : EGT > 610⁰
When : EGT > 635 deg.C:
•
the indication becomes red
•
the MASTER WARN light comes on accompanied by the
repetitive chime
5.4 ANALYZERS – DESCRIPTION AND OPERATION
5.4.1 General
The EVMU located in the electronics bay provides the following
functions by means of accelerometer and tachometer signals and
ARINC 429 lines.
•
To give data on the unbalance conditions of each rotor of each
engine (with narrow-band tracking filters) and to transmit them
for display in the cockpit with advisory levels based on fixed
maximun limits and in some conditions on initial levels.
•
To give data to rebalance the low-pressure rotor.
•
To give broadband filtered data with threshold indications for
each engine for maintenance purposes.
•
To give automatically or upon request a frequency analysis for
maintenance.
•
To communicate with the CFDS for maintenance.
5.4.2 Engine Vibration Transducer
The system monitors engine vibration for engine 1 and engine 2.
Monitoring is performed 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 cockpit. Two channels come from each
engine.
5.4.3 Engine Vibration Monitoring Unit (EVMU)
The Engine Vibration Monitoring Unit (EVMU) is located on the
avionics compartment shelf 86VU. The EVMU receives analog signals
from:
•
the 4 engine accelerometers 2 per engine (A and B).
•
the N1 speed sensor (TRIM balance probe signal) and the N2
speed sensors of each engine.
It also receives digital input from the CFDS through the ARINC 429
data bus.
The EVMU sends signals through the digital ARINC 429 data bus to:
•
the SDAC1 and 2 for cockpit indication
6
EXHAUST
6.1 EXHAUST - GENERAL – DESCRIPTION AND
OPERATION
6.1.1 General
The air entering by the fan is divided into two parts, the cold and hot
airflows, which are mixed through the common nozzle assembly
(CNA) before exhausting to the atmosphere. The engine exhaust
section directs fan discharge air to either normal or reverse thrust
operation. In the forward thrust mode, the fan airflow and burnt gases
are evacuated directly at the back. Each engine is equipped with a
reverser system which reverses cold fan air by means of translating
sleeves and cascades, turning the engine airflow forward and
providing a braking effect for the aircraft on the ground. Thrust
reversers can be operated on the ground only.
6.1.2 Description
The exhaust system consists of the common nozzle assembly and the
thrust reverser system.
A. Common Nozzle Assembly (CNA)
The common nozzle assembly forms the rear part of the engine. It
consist of a common nozzle exhaust collector and the engine exhaust
cone.
B. Thrust Reverser System
The thrust reverser system uses part of the engine exhaust power to
provide additional aerodynamic braking during aircraft landing. The
thrust reverser system is hydraulically operated by hydraulic pressure
from the aircraft hydraulic system. The aircraft hydraulic system is
pressurized by pumps mounted on each engine. It is controlled
through the FADEC (Full Authority Digital Engine Control) from the
cockpit by a lever hinged to the corresponding throttle control lever.
6.2 MIXED EXHAUST SYSTEM – DESCRIPTION AND OPERATION
6.2.1 General
The mixed exhaust system collects two flows of air. The first is the
cold airflow, which is the fan bypass air. The second is the hot airflow
which comes from the engine core. The mixed exhaust system is
made up of the common nozzle exhaust collector and the engine
exhaust cone.
•
The common exhaust collector admits the hot and cold gas
outflows. These gas outflows then go out to the atmosphere
through the common nozzle. The nozzle forms a convergent
duct which increases the speed of the mixed gas to give
forward thrust.
•
The engine exhaust cone forms the inner contour of the
common nozzle exhaust collector. It is made of a welded
inconel panel, an attachment ring and a closure cone.
6.2.2 Description
(1) The common nozzle exhaust collector is made up of an inner duct
with four radial struts. These struts support an outer duct. The unit is
installed on the outer rear flange of the engine turbine case.
(2) The inner duct forms a ring around the core outlet nozzle. The
front flange of the duct is used to install the exhaust collector to the
outer rear turbine case flange.
(3) Four support struts, positioned radially around the inner duct, hold
the exhaust collector outer duct assembly. Two side struts are
positioned horizontally and are riveted on either side of the inner
duct. Each of the struts is attached to the outer duct by two bolts and
captive nuts.
(4) The lower center strut lines up with the vertical center line. It is
riveted to the inner duct. The strut is attached to the outer duct by
eight bolts with captive self-locking nuts.
(5) At the top center position, an upper strut assembly straddles the
inner duct across the vertical center line. It is riveted to the inner duct
and to the outer duct.
(6) The collector outer duct is made up of a conical duct to which a
sheet metal forward fairing is riveted. The leading edge of the fairing
is against the rear of the cold airflow thrust reverser case. The
translating cowl trailing edge overlaps the outer duct leading edge.
(7) A common nozzle fairing is attached to the rear of the upper
support strut and to the inside of the outer duct along its upper edge.
The fairing tapers to a pen nib profile at its trailing edge. This gives a
streamlined profile for the cold stream air which enters the collector.
(8) Interface seals provide sealing between the exhaust collector, the
thrust reverser and the pylon.
6.2.3 Interface Sealing
Seals are installed around the front of the inner duct and upper and
lower support struts. They seal the inner duct with the thrust reverser
C-ducts. These seals are made in sections to suit the directional
changes in the seal run.
A circumferential seal is installed around the outside of the forward
section of the exhaust collector. This seal gives an interface seal
between the collector and the rear edge of the translating cowl. It
prevents leakage of fan air to ambient during forward thrust
operation. Block seals are attached to the left and right front corners
of the upper support strut. They complete the seal between the
aircraft pylon, the C-ducts, the upper strut and the outer duct seals.
6.3 THRUST REVERSER SYSTEM – DESCRIPTION AND
OPERATION
6.3.1 General
The thrust reverser system for the V2500 engine is designed for use
on the ground to reduce aircraft landing roll. The thrust reverser is
designed to be used at two extreme positions, stowed and deployed.
In the stowed position, the thrust reverser provides an aerodynamic
fairing between the fan cowl and the core nozzle. When deployed, the
translating sleeves are hydraulically moved rearward exposing the
cascades ; the blocker doors are positioned to close the fan duct
downstream and the fan air exhaust flow is diverted forward through
the cascades providing a braking effort for the aircraft. The thrust
reverser system is synchronized by means of a flexible synchronizing
shaft connecting all four translating sleeve hydraulic actuators. Thrust
reverser shutoff valve (TRSOV) is designed to isolate the Thrust
reverser shutoff valve (TRSOV) is designed to isolate the thrust
reverser from the aircraft hydraulic system.
6.3.2 Description
The thrust reverser system comprises:
•
a hydraulic control unit (HCU) including:
-
an isolation valve
-
an isolation valve solenoid
-
a directional control valve
-
a directional control valve solenoid
-
a pressure switch
-
a filter and clogging indicator
-
a lockout lever
•
four actuators with internal lock for lower actuators
•
three flexible shafts
•
two linear variable differential transformers located on each
upper actuator
•
two proximity switches located on each lower actuator
•
two thrust reverser cowls comprising a fixed structure and 2
translating sleeves latched together. one thrust reverser shutoff
valve.
6.3.3 Operation
The thrust reverser is actuated in response to signals from the Engine
Electronic Control (EEC). Selection of either stow or deploy from the
cockpit generates a signal to the engine EEC which in turn, supplies
signals to the thrust reverser hydraulic control unit.
6.3.4 Thrust Reverser Manual Opening/Closing
Thrust reverser system can be operated manually on the ground for
maintenance purposes using a manually operated drive shaft. A
manually bypassable non return valve is used to allow movement of
the translating sleeves without any hydraulic blockage.
6.4 THRUST REVERSER CONTROL AND INDICATING –
DESCRIPTION AND OPERATION
6.4.1 General
The thrust reverser is controlled from the cockpit by means of the throttle control lever.
A thrust reverser latching lever, when actuated, allows the throttle control lever to be
moved rearward in the reverse thrust selection area and the thrust reverser to be
controlled by the Engine Electronic Control (EEC).The EEC incorporates a thrust
reverser feedback position and ground/flight configuration which generates a command
signal to the isolation valve and the directional control valve. The signal from the EEC
to the directional valve is fed to the avionics compartment where it passes through an
inhibition relay (also called permission switch) controlled by the Engine Interface Unit
(EIU) and the spoiler/elevator computer (SEC) according to the throttle control lever
position.
6.4.2 Component Location
6.5 THRUST REVERSER HALVES – DESCRIPTION AND
OPERATION
6.5.1 General
The thrust reverser is attached to the pylon (immediately aft of the
engine fan case) and to the fan case by the vee blade-vee groove
system. The reverser is constructed in two halves which are hinged at
the top to the pylon and latched together along the bottom
centerline.
6.5.2 Description
Each thrust reverser half is supported from the pylon by four hinges.
These hinges are integral with a machined aluminium beam which
runs along the upper section of the thrust reverser. The thrust
reverser consists of four parts : the C-duct fixed structure, the
translating sleeve, the blocker doors and the cascades.
6.5.3 Material
The C-duct inner barrel is constructed of an aluminium face sheet
bonded to an aluminium honeycomb core. The cascades are
constructed of moulded graphite. The blocker doors are machined
aluminium. The blocker door drag links and cascade aft rings are
made of aluminium. Translating sleeve outer skin is constructed of a
graphite panel composed of four layers of graphite skin with
aluminium core
frames. The translating sleeve inner skin is
constructed of Dyna Rohr acoustic material which consist of 3/8 inch
titanium core all with graphite skins ; two plies of back skin and three
plies of perforated skin.
6.5.4 Removal / Installation
The thrust reverser incorporates hoisting points for removal and
installation of the thrust reverser C-duct assemblies with the use of a
three point handling sling.
6.6 THRUST REVERSER OPENING MECHANISM –
DESCRIPTION AND OPERATION
6.6.1 General
The purpose of the thrust reverser opening mechanism is to open the
C-duct doors to permit access to the engine for maintenance. The
opening system is manually operated by a GSE hydraulic hand pump.
6.6.2 Cowl Latch
A. General
The thrust reverser opening mechanism consists of two parts for each
C-duct door : the opening actuator and the hydraulic tube.
B. Opening Actuator
The opening actuator acts as a piston with the base attached to the
top of the C-duct door and the end to the pylon. As the hydraulic
pressure builds up in the actuator, the piston extends to move the
door outward. The actuator is not designed to hold the weight of the
door for a long time. The door hold open rods must be positioned to
support the door.
C. Hydraulic Tubing
The hydraulic tubing is made up of three parts : a flexible hose, a
hard tubing and a manifold connector.
•
The flexible hose connects to the opening actuator to absorb
the rocking movement of the actuator as the C-duct door opens
and closes.The flexible hydraulic hose is inside a larger conduit.
Fan air from the precooler is brought into the area between the
hose and the cover to cool the hydraulic fluid. The two ends of
the cover are half-shells made of firm fiber-glass. The half-shells
at the opening actuator can be separated to permit access to
the hydraulic tube connection. The single half-shell at the lower
end can be raised to permit access to the hydraulic tube
connection.
•
The hard hydraulic tubing connects to the flexible hose and
runs down the forward edge of the C-duct door. Near the
bottom of the door the tubing connects to the manifold.
D. Cowl Latch
A total of six latches are used to secure the two _C_ ducts to each
other.
Access to front latch is through the left and right fan cowls. Three
latches are located under a hinged access panel. Rear latch is a
double
latch
but
the
two
latches
must
be
released/latched
individually.
6.6.3 Operation
The thrust reverser cowls can be opened to a 45 degree position for
engine removal or other maintenance purposes.
7
OIL
7.1 OIL – GENERAL – DESCRIPTION AND OPERATION
7.1.1 General
The lubrication system provides lubrication and cooling to the engine
components. The oil system includes separate pressure and scavenge
pumps, a fuel cooled oil cooler, an air cooled oil cooler, an oil tank
and associated valves and filters.
7.1.2 Operation
The oil leaves the tank and is routed to a one-stage pressure pump.
The oil is piped through a pressure filter to the air cooled oil cooler
(ACOC) and then to the fuel cooled oil cooler (FCOC). After leaving the
FCOC, the oil is delivered to the No. 1, 2, 3 bearing compartment, the
No. 4 bearing compartment, the No. 5 bearing compartment and to
both the main and angle gearboxes. The scavenge oil is then piped,
either directly or through the deoiler, to the scavenge pumps. From
the scavenge pumps all the oil is pumped in a common line to a fine
scavenge filter and sent to the oil tank where it is de-aerated.
7.1.3 Lubrication System Components
The lubrication system consits of four subsystems, each of which
accomplishes a specific function to ensure adequate lubrication for
the bearings and gears of the engine:
•
the lubrication supply system
•
the lubrication scavenge system
•
the oil seal pressurization system
•
the sump venting system.
7.2 STORAGE – DESCRIPTION AND OPERATION
7.2.1 General
The oil tank contains the oil to be delivered under pressure to the
engine and external gearbox. All the scavenged oil returns to the tank
to maintain the oil level during the time the engine is operating. The
oil tank capacity allows a maximum of 40 flight hours without any
replenishment.
7.2.2 Description
•
The oil tank is attached to the LH side of the external gearbox
at the 7 o
•
clock position on the fan case.
The oil tank has:
(1) A mounting boss for the quantity transmitter
(2) Oil servicing:
•
A gravity fill port with a scupper drain
•
Remote fill and overflow ports
•
A prismatic oil level indicator
•
A usable oil capacity of 25 US quarts
(3) A tank pressurization valve:
•
This valve supplies adequate inlet pressure to the inlet of the oil
pump lubrication stage.
(4) A filler neck valve:
•
This valve prevents rapid oil loss through the oil filler neck.
(5) An internal de-aerator:
•
This de-aerates the oil which comes from the scavenge filter.
7.3 DISTRIBUTION – DISCRIPTION AND OPERATION
7.3.1 General
The function of the oil distribution system is to provide lubrication and
cooling to the essential components of the engine.
7.3.2 Description
The oil distribution system includes the following components:
•
an air cooled oil cooler
•
a fuel cooled oil cooler
•
a pressure oil filter element
•
a scavenge oil filter element
•
a one-stage pressure pump
•
a five-stage scavenge pump
•
a deoiler
•
a two-position scavenge valve
•
a filter bypass valve
•
magnetic chip detectors
7.3.3 Supply
The oil leaves the tank through a strainer and is routed into a onestage pressure pump. It is pumped through a pressure filter to
remove any large debris. The oil is then piped to the air/oil heat
exchanger before entering a fuel cooled oil cooler. After leaving the
FCOC, the oil flows to the No. 1, 2 and 3 bearing compartment, to the
No. 4 bearing compartment and to the No. 5 bearing compartment,
and to both the main and angle gearboxes.
7.3.4 Scavenge
A. General
The oil (which has lubricated the engine bearings and the angle and
main gearbox) is piped either directly or through the deoiler to the
scavenge pumps.
The line to each scavenge pump incorporates a strainer and a
magnetic chip detector. The chip detector for the No. 4 bearing
compartment is in the deoiler
scavenge outlet. The combined flow from all the scavenge pump
outlets is returned to the oil tank via a cyclone type deaerator
mounted within the tank.
B. Description
The scavenge circuit main components are:
•
five scavenge filters with chip detectors,
•
five scavenge pumps,
•
one common scavenge filter.
•
A 2-positions scavenge valve.
7.4 OIL QUANTITY INDICATING SYSTEM – DESCRIPTION AND
OPERATION
7.4.1 General
A. Oil Quantity Transmitter
•
The oil quantity transmitter is located in the oil tank.
•
The system is supplied with 28VDC from busbar 101PP through
circuit breaker 1EN1 (1EN2).
B. Description :
The oil quantity tranmitter is a tank probe with a capacitor (tube
portion) and an electronic module (on the top of the transmitter) for
probe enrgizing and signal output.
C. Specification:
•
Output voltage : 1VDC to 9VDC varying linearly with the usable
oil quantity from 0 to 25.8 quarts.
•
33 K OHMS less than charge impedance less than 100 k OHMS
•
Accuracy : plus or minus 0.5 to plus or minus 0.6 US quarts.
D. Oil Quantity Indicating
•
The analog signal from the oil quantity transmitter is sent to:
-
the SDAC1
-
the SDAC2
-
the EIU which transforms the analog signal into a digital
signal.
•
The DMC’s process the information received as a priority order
from the EIU’s through FWC 1 and 2, SDAC1, SDAC2.
•
Scale:The oil quantity displayed in green on the ECAM display
unit is graduated from:
-
0 to 27.5 quarts in analog form (the normal max-usable
oil quantity in the tank is 25 US quarts, the maximum oil
tank capacity is 30.5 US quarts).
-
0 to 99.9 in digital form.
•
Advisory level:The oil quantity flashes green when the oil level
quantity is lower than or equal to 5 quarts (decreasing) and
stop pulsing at 7 quarts (increasing).
•
Overfilling of the oil tank can be generate an out of range
feedback to the EIU. This may result in amber XX’s being
displayed in place of the oil quantity.
7.5 OIL TEMPERATURE INDICATING SYSTEM –
DESCRIPTION AND OPERATION
7.5.1 Scavenge Oil Temperature Thermocouple
The scavenge oil temperature thermocouple is located in the
combined scavenge line between the master magnetic chip detector
and the scavenge filter for indication in the cockpit. The oil
temperature is sensed by a dual resistor unit. The unit consists of a
sealed, wire-wound resistance element. This element causes a linear
change in the DC resistance when exposed to a temperature change.
Temperature measurement range: - 60 deg. C to 250 deg. C.
7.5.2 Oil Temperature Indicating
The analog signal from the scavenge oil temperature thermocouple is
transmitted to the EIU. The EIU transforms this signal into a digital
signal. This digital signal is then transmitted to the lower ECAM
display unit through the FWCs and the DMC. Scale : the ECAM oil
temperature indication is graduated from 0 deg.C to 999 deg. C.
Advisory level : the oil temperature indication flashes green when the
temperature is higher or equal to 156 deg. C and becomes steady
amber beyond 165 deg. C. The warning message ENG 1(2) OIL HI
TEMP is displayed on the upper ECAM display unit associated with the
amber MASTER CAUT light and the single chime when:
•
the oil temperature is higher than 156 deg. C for 15 mm,
•
the oil temperature is higher than 165 deg. C.
The warning message ENG 1(2) OIL LO TEMP is displayed on the
upper ECAM display unit associated with the amber MASTER CAUT
light and the single chime when the oil temperature is lower than -10
deg. C (with the throttle above idle and the engine running).
7.6 OIL PRESSURE INDICATIND SYSTEM – DESCRIPTION AND
OPERATION
7.6.1 Oil Pressure Transmitter
A. General
The oil pressure indicating system gives a cockpit in indication of the
engine oil system working pressure. The indication of this pressure
comes electrically from an oil pressure transmitter on each engine.
B. Operation
•
The oil pressure transmitter is bolted to a bracket on the top
left side of the engine fan case.
•
The oil pressure transmitter is connected to the engine oil
system by two steel tube. One tube connects to the oil supply
tube (to the engine and gearbox bearings). The other tube
connects to the No. 4 bearing oil scavenge tube (to the oil
scavenge pump).
•
Power supply : 28VDC from busbar 101PP (202PP).
•
Pressure range : 0 to 400 psid.
•
Output voltage : 0.7VDC to 9.1VDC varying lineearly with
pressure from 0 to 400 psid.
•
Accuracy : plus or minus 1%, for 0 to 400 psid.
C. Operation
The 28VDC supply goes through the electromagnetic interference
(EMI) filter to the DC/DC converter which changes the voltage to
15VDC. This decreased voltage goes to the strain gauge and amplifier
regulator. In the strain gauge, the voltage causes an excitation of the
circuit. Each diaphragm moves by oil pressure from the inlet port. As
the diaphragm moves it changes the condition of the attached strain
gauge. This causes a change in the associated excitation voltage.
Thus the voltage is in proportion to the oil pressure at the diaphragm.
The excitation voltage goes from the strain gauge to the amplifier
regulator. The amplifier increases the voltage to between 0 and 9VDC
(changes as the oil pressure changes). This is the output signal
voltage which goes through the EMI filter to the cockpit. The signal
voltage of 0 to 9VDC gives a pressure range of between 0 and 400 psi
(2720 kPa).
7.6.2 Oil Pressure Indicating
•
The analog signal from the oil pressure transmitter is sent to
the SDAC1, SDAC2 and the EIU which transforms the analog
signal into a digital signal.
•
The digital signal is then transmitted to the ECAM through the
FWCs and the DMC.
•
The order of priority has been defined as follows:
SDAC1
SDAC2
EIU.
7.7 LOW OIL PRESSURE WARNING SYSTEM – DESCRIPTION AND
OPERATION
7.7.1 Low Oil Pressure Switch
The low oil pressure switch is installed on a bracket at the top left
side of the engine fan case. The low oil pressure switch is connected
between the oil supply tube and the No. 4 bearing scavenge tube.
Actuation of the low pressure switch is at 60 psid decreasing
pressure. The set point range is between 45 psi and 75 psi.
7.7.2 Low Oil Pressure Indication (Warning)
When the oil pressure drops below 60 psid plus or minus psid
(decreasing) the low oil pressure switch closes. As a result of this:
•
The red MASTER WARN light located on the glareshield comes
on.
•
The audio warning is activated.
•
The ENG page appears on the lower ECAM display unit: The oil
pressure indication flashes red.
•
These warning messages appear on the upper ECAM display
unit: ENG1 (2) OIL LOW PRESS*
THROTTLE 1 (2) IDLE*
•
Activation on the ground of the following circuits:
WN (Emergency Escape Slide Release Warning)*
CC (Engagement and Internal Monitoring)*
CA (Electrical Power Supply).
When the oil pressure reaches 60 psid increasing, the low oil
pressure switch opens. As a result of this:
•
The MASTER WARN light goes off.
•
The audio warning is deactivated.
•
The oil pressure is indicated in green.
•
The warning messages disappear on the upper ECAM display
unit.
7.8 OIL FILTER CLOGGING WARNING SYSTEM –
DESCRIPTION AND OPERATION
7.8.1 Scavenge Filter Differential Pressure Switch
The scavenge filter differential pressure switch is installed on a
bracket at the top left side of the engine fan case. The switch is
connected to ports on the filter housing. The differential pressure
switch is set to operate when the differential pressure increases and
gets to 12 psi plus or minus 2 psi
7.8.2 Oil Filter Clogging Indication
When the differential pressure through the scavenge filter is higher
than 12 psi plus or minus 2 psi, the switch closes. As a result:
•
On the ENG page on the lower ECAM display unit: The amber
OIL FILTER CLOG indication appears.
•
On the upper ECAM display unit: The warning message ENG
1(2) OIL FILTER CLOG appears. When the differential pressure
through the scavenge filter drops below 12 psi, the switch
opens. As a result:
•
The OIL FILTER CLOG indication on the ENG page on the lower
ECAM display unit disappears.
•
The warning message on the upper ECAM display unit
disappears.
7.9 NO 4 BEARING INDICATING SYSTEM – DESCRIPTION
AND OPERATION
7.9.1 General
The purpose of the No.4 bearing indicating system is to monitor the
correct operation of the No.4 bearing 2-position scavenge valve and
to detect a No.4 bearing carbon-seal failure.
7.9.2 Description
A. No.4 Bearing Pressure Transducer
The No.4 bearing pressure transducer is installed on the right side of
the deoiler and senses pressure at the No.4 bearing outlet line.
Accuracy : plus or minus 2.5 percent above 100 PSID, plus or minus 1
percent below 100 PSID
Linear output 1VDC to 9 VDC (0 To 300 PSIG), Impedance 33k ohms
to 100k ohms.
B. No.4 Bearing Scavenge Valve Position Indicator.
The position indicator is integral with the No.4 bearing scavenge
valve which is located in the No.4 bearing compartment scavenge
line. At engine low power, the valve is open and the switch closes
providing a ground signal. At engine high power, the valve closes and
the switch opens.
7.9.3 No 4 Bearing Indication
The EIU incorporates three logics allowing the monitoring of the
scavenge valve operation as well as a No.4 bearing carbon-seal
failure:
•
When a No.4 bearing scavenge valve failure (stuck close or
stuck open) occurs:
•
an ECAM warning is displayed.
•
an amber indication appears showing the position of the failed
valve.
•
The EIU elaborates a label used to set a message on the CFDS.
•
When a No.4 bearing carbon-seal failure or a scavenge valve
restrictor clogging occurs:
•
A message appears on the maintenance part of the ECAM
status page (activated at landing).
•
The EIU elaborates a label used to set a message on the CFDS.
8
STARTING
8.1 STARTING – GENERAL – DESCRIPTION AND
OPERATION
8.1.1 General
The starting system of the engine utilizes pressurized air to drive a
turbine at high speed. The turbine turning through a reduction gear,
applies a torque to the HP shaft, thus driving the engine and the
aircraft accessories. The air necessary for the starting is bled from the
duct connecting engine bleed and the precooler via a 4 in.dia. duct.
This duct is attached to the engine before the precooler inlet. The air
necessary for the starter is supplied by either:
•
the other engine through the crossbleed system
•
the APU and in that case, all the air bled from the APU is used
for starting
•
an external source able to supply a pressure between 30 and
40 psi.
8.2 CRACKING DESCRIPTION AND OPERATION
8.2.1 Air Supply
The air necessary for the starting comes from the duct connecting
engine bleed and the precooler via a 4 in.dia. duct. The air necessary
for the starter is supplied by either:
•
the other engine through the crossbleed system
•
the APU and in that case, all the air bled from the APU is used
for starting
•
an external source able to supply a pressure between 30 and
40 psig.
8.2.2 Dry Cracking
A. Requirement
(1) A dry motoring of the engine will be needed when:
•
it is necessary to eliminate any fuel accumulated in the
combustion chamber
•
a leak ckeck of engine systems is needed.
To perform this operation, the starter is engaged and the engine is
motored but the HP fuel shut off valve remains closed and both
ignition systems are OFF. An engine dry motoring can be performed
for a maximum of three consecutive cycles (2 of 2 minutes and 1 of 1
minute with a cooling period of 15 seconds between each cycles).
After three cycles or 4 minutes of continuous cranking, stop for a
cooling period of 30 minutes.
B. Dry cranking control
(1) A selector switch is located on ENG panel 115VU.
C. Automatic Dry Cranking
An automatic selection of dry cranking is accomplished during the
engine start procedure (when the ENG/MASTER control switch is set to
ON) and when the starting sequence is aborted by the FADEC. This
can be interrupted at any time by placing the MASTER control switch
in OFF position.
8.2.3 Wet Cracking
A. Requirement
A wet motoring will be needed when the integrity of the fuel system
has to be checked. If such a test is performed, both ignition systems
are off and the starter is engaged to raise N2 up to the required
speed of 20%. The MASTER control switch is moved to ON and the
exhaust nozzle of the engine carefully monitored to detect any trace
of fuel. The wet motoring can be performed for a maximum of three
consecutive cycles (2 of 2 minutes and 1 of 1 minute with a cooling
period of 15 seconds between each cycles). After three cycles or 4
miutes of continuous cranking, stop for a cooling period of 30
minutes. In all cases, the MASTER control switch will be returned to
OFF and the engine motored for at least 30 seconds to eliminate
entrapped fuel or vapor.
B. Wet Cranking Control
(1) A selector switch is located on the ENG panel 115VU.
8.3 CONTROL AND INDICATING SYSTEM – DESCRIPTION AND
OPERATION
8.3.1 Starter
The starting control signals are acquired by the FADEC Engine
Electronic Control (EEC) from the ENG panel (115VU), and the
overhead panel via the EIU.
8.3.2 Operation
A. The ENG panel (115VU) initial configuration is:
•
ENG/MODE selector switch in NORM position
•
both MASTER control switches in OFF position.
-
Engine dry cranking.
-
Engine wet motoring.
-
Engine starting (manual and automatic)
B. Start interruption
The interruption of automatic starting can be achieved if:
(1) You place the MASTER control switch back to OFF. This causes:
•
the direct closure of the HP fuel SOV
•
the starter pneumatic valve closure via the EEC
•
the ignition exciters de-energization.
C. Alternate Engine Start Sequence
The alternate start procedure is achieved by a dedicated logic module
in the EEC which allows separate control of:
•
the pneumatic starter valve through ENG/MODE selector switch
positioning and MAN START pushbutton switch
•
the HP fuel SOV through the MASTER control switch.
D. Engine Relight
The automatic and alternate start procedures are both available for
engine relight with the same procedure as on the ground.
•
When the automatic start is selected in flight, the EEC identifies
windmilling or starter-assisted relight conditions according to
flight environmental parameters or engine parameters.
•
When the alternate start is selected in flight the EEC always
commands a starter assisted relight.
•
Continuous ignition is selected whenever the selector switch is
placed in IGN/START position.
E. Engine Flameout
In case of engine flameout the EEC is designed to:
•
Detect unscheduled sub-idle engine operation.
•
Transmit a message for cockpit display indicating a probable
flameout condition.
•
Activate the continuous ignition.
8.4 ENGINE STARTING SYSTEM – DESCRIPTION AND
OPERATION
8.4.1 Starter
The starter is attached to the forward face of the gearbox using a cast
aluminium adapter and Quick-Attach-Detach (QAD) clamp. It is
designed to rotate and accelerate the HP rotor of the engine to allow
the starting.
The starter consists of a drive turbine assembly housed in steel/
containment housing, a gear reduction system, a clutch, output shaft
assembly, a mounting adapter, Quick Attach Detach (QAD) clamp.
Steel inlet housing provides the starter air inlet flange and provides
containment for the turbine rotor. Steel nozzles which direct air into
the rotor, are integral with the turbine inlet housing. Flow paths are
shaped to minimise inlet and outlet losses. The drive turbine rotative
assembly consists of the rotor, shaft, pinion gear
and bearing
cartridge. The turbine blades, hub, and shaft are integral and
machined from a titanium forging. The shaft is supported by the
bearing cartridge which utilises two preloaded angular contact ball
bearings. A carbon seal assembly
is used to prevent external oil
leakage from the rotor/bearing system.
8.4.2 Pneumatic Starter Valve
The pneumatic starter valve is installed on the air duct on the lower
right side of the LP compressor. It is provided to control the supply of
the starter with air, from an order given by the EEC.
(1) The starter valve is a butterfly type valve, pneumatically operated
and electrically controlled. The valve is installed on the air duct on the
lower right hand side of the LP compressor case.
(2) The starter valve controls the flow of air from the air duct to the
engine starter.
(3) The steel butterfly valve is installed on a stainless steel shaft,
which is held in PTFE bushes.
(4) An extension on the valve shaft can be used to manually operate
the valve. This can be done with a 0.375 in. square socket. The
extension includes a switch and a visual position indication device. As
the valve opens more than 7 degrees, the switch operates and
transmits a position indication to the EEC.
(5) Pneumatic control is given by a two-function diaphragm type
actuator and a solenoid valve. The actuator has two pistons of
different areas which are installed on the same shaft. Actuator
movement is linked mechanically to the butterfly valve shaft.
(6) The solenoid valve has an electrically operated solenoid, a ball
valve and a plunger loaded by a spring.
BIBILIOGRAPHY
1. AIRCRAFT MAINTENANCE MANUAL
- AIR INDIA
2. AIRCRAFT GAS TUBINE ENGINE TECHNOLOGY
- TREAGER
3. AIRCRAFT POWERPLANTS
- KROES – WILD
4. GAS TURBINE ENGINE
BRATON WHITNEY
5. TRAINING MANUAL ENGINE (V2500) A320
- AIR INDIA
6. 15 A TRAINING MANUAL
- AIR INDIA