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.

ABSTRACTThe concept of Jet propulsion Engine has started in 150-BC by Hero through his Aeolipile based on Newtons 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. OIL10. STARTING

INTRODUCTIONMost 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 358 kg/s 5.06

0.560 lb/lb-hr 789 lb/s

Airflow Twr(max takeoff thrust)

Different stages with blade specificationsFAN-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 Cduct.

1

POWERPLANT

0.1 POWER PLANT - GENERAL DESCRIPTION AND OPERATION0.1.1 GeneralThe 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 DescriptionA. 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 engine core, the fan case, the accessories, the accessory gearbox.

The attachment fittings and support brackets are attached on:

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 GeneralThe 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 GeneralThe 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 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

The assembly also includes installation of:

0.3.2 DescriptionA. 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 GeneralThe 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 DescriptionThe 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 GeneralThe 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 LocationThe 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 DescriptionThe engine mount system has these components: A front mount 1500KM, A rear mount 1550KM.

0.5.4 OperationA. 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 GeneralThe 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 DescriptionA. 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.

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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. 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.

that are part of the core zone harness are:

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 OperationThe 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 GeneralThe 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 OperationA. 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 Variable Stator Vane Actuator ) ) Engine ) Core

Active Clearance Control Actuator ) Fuel diverter valve Air Cooled Oil Cooler actuator Fuel metering unit LP/HP fuel pumps ) ) Engine fan case ) Engine fan case ) Gearbox

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 Hydraulic Pump Oil tank scupper ) Gearbox ) ) 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 ModulesThe 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 OPERATION1.2.1 GeneralThe 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 AssemblyThe 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 moduleThe 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 ModuleThe 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 GeneralThe 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 AssembliesA. 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 AssemblyThe 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 SectionA. 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 SectionThe 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 GeneralThe 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 GearboxA. 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 GearboxA. 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 CONTROLLING3.1 ENGINE FUEL AND CONTROL GENERAL DESCRIPTION AND OPERATION 3.1.1 GeneralThe 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 DistributionThe 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 ControllingThe 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 IndicatingThe 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 OPERATION3.2.1 GeneralThe 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 TubesThe 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 PumpThe 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 FilterA. 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 NozzleA. 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 ValveA. 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 ValveA. 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 OPERATION1.6.1 GeneralThe 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 DescriptionA. 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. Fan and core over speed protection to prevent engine running over certified red lines Engine turbine outlet gas temperature monitoring.

Engine limits protection -

Power management 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), N1, N2, WF, Exhaust GasTemperature (EGT) monitoring Abort/Recycle capability on ground. Control of thrust reverser actuation Thrust reverser control 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.

-

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. 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

The FADEC provides:

specific fault message transmission. 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.

(b) Alternate start logic with authority of the FADEC limited to:

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 DescriptionA. 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 GeneralThe 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 DescriptionA. 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 InterfraceA. 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 InterfaceA. 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 SensorAll 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 GeneralThe 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 IndicationA. 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 IndicationA. 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 DescriptionThe 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 OperationThe 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 DescriptionThe 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 DescriptionThe 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 GeneralThe ignition system is controlled by: the EEC upon commanded signals from ENG START panel 115VU through the EIU.

2.4.2 DescriptionA. 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 DescriptionThe 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.

3

AIR3.1 AIR GENERAL DESCRIPTION AND OPERATION

3.1.1 GeneralThe 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 DescriptionAll 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 N_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 indicationThis system enables the nacelle core zone ambient temperature indication to be displayed.

3.2 COOLING DESCRIPTION AND OPERATION 3.2.1 GeneralThe 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 LocationA. 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 GeneralThe 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 DescriptionA. 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. 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.

(6) The maximum permitted air temperatures are:

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 GeneralA. The engine main bearings are contained in three bearings compartments: The front bearing compartment. The N4 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 N4 bearing compartment cool.

3.4.2 DescriptionA. 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 sealingA. 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 CoolerA. 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 GeneralA. 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.