I. A. E V2500 COURSE NOTES I.A.E. V2500 COURSE NOTES CONTENTS INTRODUCTION TO V2500 PROPULSION UNIT PART ONE - ENGINE PART TWO - NACELLE SECTION 1 INTRODUCTION SECTION 1 INTRODUCTION SECTION 2 MECHANICAL ARRANGEMENT SECTION 2 MECHANICAL ARRANGEMENT SECTION 3 ELECTRONIC ENGINE CONTROL SECTION 3 THRUST REVERSER SECTION 4 POWER MANAGEMENT SECTION 4 NACELLE VENTILATION AND SECTION 5 FUEL SYSTEM FIRE PROTECTION SECTION 6 OIL SYSTEM SECTION SECTION 7 HEAT MANAGEMENT SYSTEM SECTION 8 COMPRESSOR AIRFLOW CONTROL PART THREE - GENERAL SYSTEM SECTION 1 TROUBLE SHOOTING SECTION 9 SECONDARY AIR SYSTEMS SECTION 2 COMPONENT LOCATION GUIDE SECTION 10 ENGINE ANTI-ICE SYSTEM SECTION 3 ENGINE G.A. DIAGRAM SECTION 11 ENGINE INDICATING SECTION 4 BORESCOPING SECTION 12 STARTING AND IGNITION SYSTEM SECTION 5 TRIM BALANCING 5 ENGINE REMOVAL/INSTALLATION PROPULSION UNIT - INTRODUCTION I.A.E. V2500 PROPULSION UNIT – INTRODUCTION The V2500 is an advanced technology aircraft propulsion unit designed primarily for the 150 seat, short to medium range aircraft. V2500 is a new advanced design and incorporates technology from five major engine manufacturers: Rolls-Royce Pratt & Whitney Japanese Aero Engine Corporation M.T.U. Fiat Aviazion - England - U.S.A. - Japan - Germany - Italy The propulsion unit shown below is the V2500 for the Airbus A320 Aircraft. Propulsion Unit - Data T.O. Thrust (S.L. Static) Flat Rated Temperature Total Airflow By-pass Ratio Overall Pressure Ratio Fan Diameter Propulsion Unit overall length Engine overall length Propulsion Unit Weight Bare Engine Weight : 25000 lbs (111205 KN). : I.S.A. + 15°C : 783 lbs (355kgs)/second : 5.42:1 : 29.4:1 : 63 inches (160cm) : 198.39 inches (503.91cm) : 126 inches (320cm) : 7300 lbs (3311kgs) : 4942 lbs (2242kgs) V2500 PROPULSION UNIT Propulsion Unit Introduction Gas Path A simplified view of the propulsion unit is shown below. All the air entering the engine passes through the inlet cowl to the fan. At the fan exit the air stream divides into two flows:the core engine flow • the by-pass flow the by-pass ratio is 5.42:1. • Core Engine Flow The core engine flow passes through the fixed inlet guide vanes to the three stage booster then to the HP compressor, the combustion section and the HP & LP turbines and finally exhausts into the CNA. By-pass Flow The fan exhaust air (cold stream) entering the by-pass duct passes through the fan outlet guide vanes and flows along the by-pass duct to exhaust into the CNA. Common Nozzle Assembly (C.N.A.) The core engine (hot) exhaust and the (coo1) by-pass flow are mixed in the CNA before passing through the single propelling nozzle to atmosphere. PROPULSION UNIT OUTLINE INTRODUCTION ENGINE MARK NUMBERS For easy identification of the present and all future variants of the V2500, International Aero Engines has introduced a new engine designation system. • All engines will retain V2500 as their generic name. • The first three characters of the full designation are V25, identifying each engine as a V2500. • • • The next two figures indicate the engine's rated sea-level takeoff thrust. The following letter shows the aircraft manufacturer. The last figure represents the mechanical standard of the engine. This system will provide a clear designation of a particular engine as well as a simple way of grouping by name, engines with similar characteristics. • The designation V2500-D collectively describes, irrespective of thrust, all engines for McDonnell • Douglas applications and V2500-A all engines for Airbus Industrie. Similarly, V2500-5 describes all engines built to the 5 mechanical standard, irrespective of airframe application. The only engine exempt from these is the current service engine, which, having already been certified, will retain the original and current designation V2500-A1 V2530-A5 V25 : Generic to all V2500 engines 30 : Take off thrust in thousands of pounds A(D) : Air frame manufacture 5 A : for Airbus Industrie D : for McDonnell Douglas : Mechanical standard of engine MK NO Takeoff Thrust (lb) Aircraft V25OO - A1 25,000 A32O-20O V253O - A5 3O,000 A321-100 V2525 - A5 2 5,000 A3 2O-2OO V2528 - D5 2 8,000 MD-9O-4O V2522 - D5 2 2,000 MD-9O-10 V2500 COURSE NOTES PART ONE ENGINE V2500 COURSE NOTES PART ONE ENGINE CONTENTS SECTION 1 SECTION 2 SECTION 3 SECTION 4 SECTION 5 SECTION 6 SECTION 7 SECTION 8 SECTION 9 SECTION 10 SECTION 11 PART ONE - SECTION 1 ENGINE INTRODUCTION PROPULSION UNIT INTRODUCTION INTRODUCTION - ENGINE The V2500 is a twin spool, axial flow, high by-pass ratio turbofan. The engine incorporates several advanced technology features which include:• • • • • Full Authority Digital Electronic Control - FADEC Wide chord fan blades Single crystal HP. turbine blades Powdered Metal HP. turbine discs . A two piece, annular combustion system which utilises segmental liners. Engine Mechanical Arrangement Low Pressure (LP) spool • The low pressure (LP) spool comprises: o a single stage fan o three stage axial flow (booster) or LP Compressor, o five stage LP turbine. • The booster stage has an annular bleed valve to improve starting and handling. LP spool speed is indicated as N1 (%). The LP Spool is supported in three bearings, one ball and two roller. • • High Pressure (HP) spool • The HP spool comprises: o a ten stage axial flow compressor o two stage turbine. • The HP compressor has: o variable inlet guide vanes (VIGVs), o four stages of variable stator vanes (VSVs) and o four bleed valves which are all used to improve starting and handling. HP spool speed is indicated as N2 (%). • • The HP spool is supported in two bearings, one ball and one roller. Lubrication The lubrication system is: • self-contained, • re-cirulatory, • full flow (unregulated pressure). Primary oil cooling is by a fuel/oil heat exchanger located in the LP fuel system, additional cooling, as required, is provided by an air/oil heat exchanger. V2500 ENGINE CUTAWAY Introduction Active Clearance Control A.C.C. (Turbine) Full Authority Digital Electronic Control (FADEC) Active clearance control (ACC) is used on both the LP and HP turbine casings, this system uses cool air taken from the fan duct. The heart of the FADEC is the Electronic Engine Control (EEC). Engine Air Bleeds Engine air bleed is utilised for: • • • • • • Aircraft systems Compressor stability system HP and LP active clearance control 10th stage make-up cooling air (turbine cooling) Air cooled air cooler (buffer air) Air cooled oil cooler Customer Services Bleed HP compressor stage 7 and stage 10 bleeds are available to the aircraft manufacturer. The EEC receives rotor speed, pressure and temperature signals from the engine, it uses these parameters along with aircraft inputs to command outputs to engine mounted actuators to provide control of:• Engine fuel flow • Automatic engine starting • Compressor airflow control system • Heat Management system • 10th stage make-up air system • thrust reverser The E.E.C. also provides protection for:• Nl overspeed • N2 overspeed • Engine surge SLTO PRESSURE - TEMPERATURE MAP ENGINE DIMENSIONS AND PRESSURE TEMPERATURE MAP V2500 AIR OFFTAKES FAN AIR • • • • • • COOLING FLOW FOR ACAC. ACTIVE CLEARANCE CONTROL SYSTEM (HP TURBINE & LP TURBINE). COOLING FLOW FOR AIR COOLED OIL COOLER (ACOC). PRE-COOLER, CUSTOMER SERVICES BLEED COOLING FLOW. IGNITION EXCITERS & HIGH TENSION LEADS COOLING C-DUCT ACTUATORS & OIL SUPPLY PIPE COOLING STAGE 7 • • • STAGE 8 • BOOSTER - 2.5 HANDLING BLEED VALVE (B.S.B.V.) STAGE 6 • SEALING - FRONT BEARING COMPARTMENT (CARBON SEALS) (1 SERIES ENGINE ONLY) INLET COWL ANTI-ICING HANDLING BLEED VALVES CUSTOMER SERVICES BLEED:o E.C.S. o WING ANTI-ICING o POTABLE WATER TANK o HYD HEADER TANK • COOLING o H.P. COMPRESSOR o L.P. TURBINE CAVITY SEALING o FRONT BEARING COMPARTMENT HYD SEAL (INTERSHAFT) o No. 5 bearing compartment (front seal) V2500 AIR OFFTAKES STAGE 10 • • CUSTOMER SERVICE BLEED MAKE UP AIR SYSTEM; ADDITIONAL COOLING FOR SPACE BETWEEN 1 & 2 HP TURBINE, DISCS & STAGE 2 HP TURBINE BLADES • No 4 BEARING SCAVENGE VALVE SUPPLY (CONTROL PARAMETER & MUSCLE AIR) HANDLING BLEED VALVES H.P.T. STAGE 2 NGV's • • STAGE 12 • • • • • • BUFFER AIR - No 4 BEARING CHAMBER COOLING FLOW MUSCLE AIR FOR HANDLING BLEED VALVES STAGE 1 HPT NGV's COOLING STAGE 1 HPT DISC FRONT FACE COOLING. STAGE 1 HPT BLADES COOLING. INNER & OUTER COMBUSTOR LINERS. PART ONE - SECTION 2 ENGINE MECHANICAL ARRANGEMENT Mechanical Arrangement General The engine is an axial flow, high by-pass ratio, twin spool, turbo fan. The general arrangement is shown below. LP System Four stage L.P. compressor comprising: o 1 Fan stage o 3 Primary stages, driven by a 5 stage axial flow LP Turbine • An annular bleed valve is located at the outlet from the booster stage • HP System • • • • • Ten stage axial flow compressor driven by a 2 stage axial flow HP Turbine Variable angle Inlet guide vanes 4 stages of variable stator vanes Handling bleed valves on stage 7 and stage 10 Customer service bleed at stages 7 and 10 Combustion System • Annular, two piece, with 20 fuel spray nozzles Gearbox • • Radial drive via a tower shaft from HP Compressor shaft to fan case mounted Angle and Main gearboxes . Gearbox provides mountings and drive for all engine driven accessories and the pneumatic starter motor ENGINE - GENERAL ARRANGEMENT Mechanical Arrangement Engine Main Bearings The main bearing arrangement, and the bearing numbering system is shown below. The five bearings are located in three bearing compartments. • The front bearing compartment: o located at the center of the intermediate case, o houses No's 1,2 & 3 bearings • The center bearing compartment located in the diffuser/combustor case houses No 4 bearing . • The rear bearing compartment located in the turbine exhaust case houses No 5 bearing No 1 Bearing • • • LP shaft axial location bearing Takes the thrust loads of the LP shaft Single track ball bearing No 2 Bearing • • • Radial support for the front of the L.P. turbine shaft Single track roller bearing Squeeze film oil damping No 3 Bearing • • • • • HP shaft axial location bearing Radial support for the front of the HP shaft Takes the thrust loads of the H.P. shaft Single track ball bearing Mounted in an Hydraulic damper which is centred by a series of rod springs (Squirrelcage). No 4 Bearing • • Radial support for turbine end of HP shaft Single track roller bearing No 5 Bearing • • • Radial support for the turbine end of the LP shaft Single track roller bearing Squeeze film oil damping No 2 BEARING No 1 BEARING No 3 BEARING No 4 BEARING No 5 BEARING ENGINE - MAIN BEARINGS Mechanical Arrangement Bearing Compartments Front Compartment Gearbox Drive The No's 1, 2 and 3 bearings are located in the front bearing compartment which is at the centre of the intermediate module (32). The HP stubshaft, which is located axially by No 3 bearing, has at its front end a bevel drive gear which, through the tower shaft, provides the drive for the main accessory gearbox. The compartment sealing utilises carbon seals, brush seals and sealing airflows obtained from the 6 th stage compressor manifold. The H.P. stubshaft separates from the HP compressor module at the curvic coupling and remains as part of the intermediate case module. An oil filled (hydraulic) seal is used between the two shafts, this seal is supported by 8 th stage air. Adequate pressure drops across the seals to ensure satisfactory sealing, are achieved by venting the compartment, by an external tube, to the de-oiler. The bearing compartment pressure, and therefore the sealing flows, are controlled by a restrictor in the vent tube. FRONT BEARING COMPARTMENT Bearing Compartments Front Compartment Continued The drawing below shows details of No2 and No3 bearings. A phonic wheel (1) is fitted to the LP stub shaft, this interacts with speed probes to provide LP shaft speed signals (Nl) to the EEC. The hydraulic seal (6) prevents oil leakage from the compartment passing rearwards between the HP and LP shafts. No3 bearing is hydraulically damped. The outer race is supported by a series of spring rods (14) which allow some slight radial movement of the bearing. The bearing is centralised by oil pressure fed to an annulus (12) around the bearing outer race. The gearbox drive gear (8) is splined onto the HP shaft and retained by No3 bearing nut (7). 1 Fonic wheel 2 LP Stub shaft 3 No.2 bearing 4 Squeeze film damper 5 No.2 bearing support 6 Hydraulic seal 7 No.3 bearing nut 8 Internal gearbox driving gear 9 No.3 bearing rotor center 10 No.3 bearing 11 No.3 bearing housing 12 Hydraulic damper 13 No.3 bearing seat support 14 Spring rod 15 No.3 bearing rear oil sea) No 2, No 3 BEARING ARRANGEMENT Mechanical Arrangement Bearing Compartments No 4 (Centre) Bearing The No 4 bearing compartment is situated in a high temperature, high pressure environment at the centre of the combustor section. The bearing compartment is encased in a thermally insulated shielding. A supply of cooled air (buffer air) is admitted to the space between the chamber double skinned walls. The buffer air is exhausted from the cooling spaces close to the upstream side of the carbon seals, creating an area of cooler air from which the seal leakage is obtained. This results in an acceptable temperature of the air leaking into the bearing compartment. Buffer air flow rates are controlled by restrictors at the outlet from the cooling passages. The bearing compartment internal pressure level is determined by the area of the variable scavenge valve, (called No 4 bearing scavenge valve and described in the oil system). This valve acts as a variable restrictor in the compartment vent line. No 4 BEARING COMPARTMENT Mechanical Arrangement bearing Compartments No 5 Bearing (Rear) The rear bearing compartment is located at the centre of the L.P. turbine module (module 50) and houses No 5 bearing which supports the L.P. turbine rotor. The compartment is sealed at the front end by a carbon seal, a simple labyrinth seal provides secondary sealing to protect against oil loss in the event of carbon seal failure. Separate venting is not necessary for this compartment because with only one carbon seal the airflow induced by the scavenge pump gives the required pressure drop across the seal. The compartment is covered by an insulating heat shield. REAR BEARING COMPARTMENT STAGE 8 SEALING AIR LP TURBINE SHAFT CARBON SEAL REAR BEARING COMPARTMENT Mechanical Arrangement: Engine Stations Stage Numbering 1 :Intake/Engine inlet interface Compressor 2 :Fan inlet Stage 1 2.5 :LP Compressor exit 12.5 :Fan exit Stage 1.5 :Booster Stage 3 :HP Compressor exit Stage 2 4 :Combustion section exit Stage 2.5 :Booster Stage 4.5 :HP Turbine exit 4.9 :LP Turbine exit Stage 3 to Stage 12 Pressure and Temperature Signals The following pressures and temperatures are sensed and transmitted to the Electronic Engine Control (EEC): P2 T2 P2.5 T2.5 P3 (Pb) T3 P4.9 P12.5 T4.9 :Fan :Booster Stage :HP Compressor Stages :HP Compressor Stages Turbine Stage 1 } Stage 2 } :HP Turbine Stages Stage 3 } to } Stage 7 } :LP Turbine Stages STATIONS 2 ROTATING STAGES 12,5 2.5 1 1.5 2 2.5 3 4 4 5 6 7 8 9 10 11 12 1 COMPRESSOR STAGES 4.5 4.9 2 3 4 5 6 7 TURBINE STAGES ENGINE STATION AND STAGE NUMBERING ENGINE Mechanical Arrangement Modular Construction Modular construction has the following advantages:• • • • • • • • • • • lower overall maintenance costs. maximum life achieved from each module. reduced turn-around time for engine repair. reduced spare engine holdings. reduced transportation costs. ease of transportation and storage. rapid module change with minimum ground running easy hot section inspection. vertical/horizontal build strip. split engine transportation. compressors/turbines independently balanced. Module Designation Module No 31 32 40 41 45 50 60 Module Fan Intermediate HP System HP Compressor HP Turbine LP Turbine External gearbox (31) FAN (32) INTERMEDIATE (50) LOW PRESSURE (60) EXTERNAL GEARBOX ENGINE MODULES Mechanical Arrangement Module 31 - Description Module 31 (Fan Module) is the complete Fan assembly and comprises: • • • • 2 hollow fan blades 2 annulus fillers the fan disc the front and rear blade retaining rings The blades are retained in the disc radially by the dovetail root. Axial retention is provided by the front and rear blade retaining rings. Blade removal/replacement is easily achieved by removing the front blade retaining ring and sliding the blade along the dovetail slot in the disc. 22 annulus fillers form the fan inner annulus. The nose cone and fairing smooth the airflow into the fan. REAR BLADE RETAINING RING WIDE CHORD FAN BLADES (22) FRONT BLADE RETAINING FAIRING FAN DISK NOSE CONE ANNULUS FILLERS (22) LP COMPRESSOR (FAN) Mechanical Arrangement Fan Nose Cone The Glassfibre cone smoothes the airflow into the fan. It is secured to the front blade retaining ring by 18 bolts. The nose cone is balanced during manufacture by applying weights to its inside surface. The nose cone is unheated. Ice protection is provided by a soft rubber cone tip. The nose cone retaining bolt flange is faired by a titanium fairing which is secured by 6 bolts. NOTE Be careful when removing the Nose Cone retaining bolts. Balance weights may be fitted to some of the bolts. The position of these weights must be marked before removal to ensure they are refitted to the same position. The arrangement is shown below. TRIM BALANCE WEIGHT LP COMPRESSOR {Fan} MODULE TRIM BALANCE WEIGHT FAIRING TRIM BALANCE WtiGHT FRONT RETAINING RING NOSE CONE Mechanical Arrangement Fan - Front Blade Retaining Ring Note: The blade retaining ring is secured to the fan disc by a ring of 36 bolts. A second {outer) ring of bolts passes through the retaining ring and screws into each of the 22 annulus fillers. The fan blades and annulus filler positions are not identified. For this reason it is important to identify the blade and annulus filler position, relative to the numbered slots in the fan disc, before disassembly. Both rings of bolts roust be removed before attempting to remove the front retaining ring. This is done using a temporary marker. After all the securing bolts (22 + 36) have been removed the retaining ring can be removed by screwing pusher bolts into the 6 threaded holes provided for this purpose. Balance weights, if required are located on the retaining ring, as shown below. The front blade retaining ring can only be fitted in one position which is determined by three off-set locating dowels on the fan disc. 1 2 3 4 5 6 7 8 9 10 11 12 13 T mark Front blade retaining ring Stage 1 fan disk Annulus filler Guide pin Headless pin (3 off) Stage 1 fan blade Pusher threaded hole (6 off) Boit (22 off} Bolt (36 off) Balance weight flange BaSance weight fim balance weight (engine pass off) 10 LP COMPRESSOR (FAN) BLADE RETAINING RING Mechanical Arrangement Fan - Fan Blades and Annulus Fillers After removal of the Front Blade retaining ring the annulus fillers can be removed as follows:• • • • • lift the front end of the annulus filler 3 to 4 inches twist the annulus filler through about 60 deg counter clockwise draw the annulus filler forward to clear the blades Remove the annulus fillers on either side of the blade to be removed. The blade to be removed can than be pulled forward to clear the dovetail slot in the fan disc. 1 Stage 1 fan disk 2 Stage 1 fan blade 3 Annulus filler FAN BLADE/ANNULUS FILLER Mechanical Arrangement Fan Blades Inspection/Repair Fan blade inspection / repair procedures are briefly described in these notes. This information is for guidance only. Before any repair is carried out reference must be made to the Maintenance Manual Chapter 72-31-11 Page Block 800. General The fan blade surface area is divided into three zones A,B and C as shown below. The acceptance limits for damage may vary depending on which zone is damaged. Inspection Blades are inspected for signs of nicks, cracks, dents, scores, scratches on the surface, and bends on the leading or trailing edges. The blades should also be inspected for signs of arc burns (lightning strikes). Arc burns or cracks are grounds for rejection and a replacement blade must be fitted. The acceptance limits for nicks, scratches, scores and dents are detailed below. If damage exceeds these limits refer to the appropriate repair scheme (see MM 72-31-11 Page Block 800 - Approved Repairs). LEADING EDGE TRAILING EDGE 0.75m (19,05 mm) 2.00m (50,80 mm) 12.00m (304,80 mm) 1.50 in (38,10 mm) 3.00 in (76,20 mm) MAXIMUM SERVICEABLE LIMITS FOR SURFACE DAMAGE DEPTH ON CONVEX AND CONCAVE SURFACES A 0.008in (0,20mm) B 0.008in (0,20mm) STAGE 1 FAN BLADES REPAIR LIMITS Mechanical Arrangement Fan Blades Inspection/Repair - General Acceptance Limits The leading and trailing edges of the blades should be examined for bends (deformations). X maximum = 0.2 in (5,08mm) if X is more than 0.2 in reject the blade. Note: Z must be not less than 8 times dimension X if Z is less than 8 times X reject the blade. • the maximum number of bent blades in any fan rotor assembly is three. • no blade may have more than one bend. • if any bend has associated cracks, kinks, creases, tears or nicks -reject the blade. • bends must be outboard of the annulus fillers, if any bend extends below the annulus filler platform reject the blade. Y must be not less than 20 times X if Y is less than 20 times X reject the blade. Note: There must be a smooth transition between the undamaged aerofoil surface and the bent area. if there is not a smooth transition reject the blade. STAGE 1 FAN BLADES REPAIR LIMITS Mechanical Arrangement Annulus Fillers - Inspection/Repair The outer surfaces of the annulus fillers should be inspected for cracks, nicks, dents, scores. • if any cracks are found reject the annulus filler. • accept dents, nicks, scores up to 0.010 inches (0.25mm) • annulus fillers with damage in excess of 0.010 may be repaired in accordance with the appropriate repair scheme (see MM 72-31-11 page block 800) 1 0.010 in. (0,25rnm) maximum blend depth 2 Apply heat resistant ES coating to this area 3 0.400in. (10,16mrn) minimum No blending permitted in this area ANNULUS FILLERS - REPAIR LIMITS Mechanical Arrangement Fan Blades - Cropping Before carrying out any blade repairs refer to the Maintenance Manual 72-31-11 page block 800 - Approved Repairs. The following pages illustrate typical cropping and scalloping limits. The following points must be noted:• always do a dye penetrant crack test when the repair is completed. • when t he repair is completed write the repair scheme number e.g. VRS 1002, in the engine log book. • at the next shop visit, after repair, the repair scheme number should be etched on the blade root. • blades that are repaired on wing must be glass bead peened at the next overhaul. FAN BLADE CROPPING LIMITS Mechanical Arrangement Fan Blade Repairs Shown below are examples of blades of blade cropping and scalloping limits. TYPICAL LEADING AND TRAILING EDGES One scallop only is permitted in this zone. POINT AP POINT AR Two scallops only are permitted on leading and/or trailing edges provided that all other conditions are met in this zone. One scallop only is Pitted in this zone on lead.ng or trailing edge. (21,59mm) One scallop only is permitted on leading or trailing edge. FAN BLADE MINIMUM ACCEPTABLE CONDITIONS BETWEEN TIP CROPPING AND SCALLOPING. The minimum dimensions quoted apply to any combination of cropping and scalloping in the blade tip area. If tip cropping has not been carried out, these dimensions apply from the blade tip points AR and AP FAN BLADE SCALLOPING LIMITS (21,59mm) Mechanical Arrangement Fan Blade Repairs Further examples of scalloping proportions and limits are shown below. 10XAB SCALLOP PROPORTIONS MIN R = 7xAB MINR = 7xAB Permitted only if extents of blend rads do not overlap — AB ,1AX = AB + 0.064«p, 11,63mm) MIN = AB AB — TYPICAL BLENDING AREA FAN BLADE SCALLOP PROPORTIONS 0.858in (21,79mm) AT TIP R 0.Q43in(1.09 mm) 0.026in (0,66mm) LEADING EDGE VIEW AB 8.955in (227,46mm) . BLEND LEADING EDGE BACK TO THIS LINE 4.220in (107,18mm) R 2.532in (64,31mm) LEADING EDGE TRAILING EDGE FAN BLADE FLY BACK REPAIRS FAN BLADE FLY BACK REPAIRS Mechanical Arrangement Module 32 - Intermediate Case The intermediate module comprises: • the fan case. • the fan duct. • the fan outlet guide vanes. • the three stage L.P.'Booster' compressor. • the booster stage bleed valve (B.S.B.V.). • the front engine mount structure. • the front bearing compartment which houses: o No 1, 2 and 3 bearings. o the drive gear for the power off-take shaft (gearbox drive). • the L.P. stub shaft. • 10 inner support struts. • 10 outer support struts • Vee gioove locations for the inner and outer barrels of the C-ducts Instrumentation The following pressures and temperatures are sensed and transmitted to the E.E.C.:• P12.5 • P2.5 • T2.5 1 Fan case 2 Fan case front panel fairing 3 LP compressor outlet guide vane 4 Intermediate structure front fairing 5 LP compressor inlet guide vane 6 Inner ring 7 LP stub shaft curvic teeth FORWARD INTERMEDIATE MODULE - FRONT VIEW Mechanical Arrangement Module 32 - Intermediate Case The rear view of the intermediate case is shown below. 1 2 3 4 5 6 7 Fan case Fan case rear panel Fan frame outer strut Intermediatestructure Forward mount Fan frame inner strut No.3 bearing curvic teeth INTERMEDIATE CASE - REAR VIEW Mechanical Arrangement Nodule 40 - H.P. Compressor The general arrangement of the H.P. compressor is shown below. The H.P. compressor has 10 stages. It utilises variable inlet guide vanes at the inlet to stage 3 and variable stator vanes on stages 3, 4, 5 and 6. The front casing, which houses stages 3 to 6, is made in two halves which bolt together along horizontal flanges, it is bolted to the intermediate casing (module 32) at the front and to the outer casing at the rear. The rear compressor casing has inner and outer casings as shown. Flanges on the inner case form annular manifolds which provide 6, 7 and 10 stage air offtakes. HP COMPRESSOR ROTOR ASSEMBLY REAR INNER CASE REAR OUTER CASE FRONT COMPRESSOR CASES VARIABLE STATOR VANE VSV OPERATING MECHANISM HP COMPRESSOR Mechanical Arrangement H.P. Compressor Compressor Drums - (Rotor) The rotor assembly is in two parts: • the stage 3 to 8 drum • the stage 9 to 12 drum The two rotor drums are bolted together with a vortex reducer installed between the 8 and 9 stages. The vortex reducer straightens the stage 8 air flow which passes to the centre of the engine for internal cooling and sealing. Mechanical Arrangement Combustion Section The combustion section includes the diffuser section, the combustion inner and outer liners, and the No 4 bearing assembly. Diffuser Casing The diffuser section is the primary structural part of the combustion section. The diffuser section has 20 mounting pads for the installation of the fuel spray nozzles. It also has two mounting pads for the two igniter plugs. Combustion Liner The combustion liner is formed by the inner and outer liners.The outer liner is located by five locating pins which pass through the diffuser casing. The inner combustion liner is attached to the turbine nozzle guide vane assembly. The inner and outer liners are manufactured from sheet metal with 100 separate liner segments attached to the inner surface. The segments can be replaced independently. COMBUSTION SECTION (1) Mechanical Arrangement Combustion Section The drawing below shows the arrangement of the diffuser casing and the outer section of the combustion liner. Also shown is the front section of the No 4 bearing compartment. COMBUSTION SECTION (2) DIFFUSER CASE ASSEMBLY HP COMPRESSOR EXIT STATOR FUEL NOZZLE (Cool air flow) OUTER COMBUSTION CHAMBER LINER No. 4 BEARING FRONT HEATSHIELD THRUST BALANCE STATIC SEAL No, 4 BEARING SUPPORT ASSEMBLY No. 4 BEARING FRONT COOLiNG AIR DUCT ' No. 4 BEARING LOCK AND NUT No. 4 BEARING FRONT COMPARTMENT No. 4 BEARING FRONT SEAL ASSEMBLY COMBUSTION SECTION (2) Mechanical Arrangement Combustion Section The drawing below shows the arrangement of the inner combustion liner and the H.P. stage 1 nozzle guide vanes. Also shown is the cooling air inlet arrangement which provides the cooling air supplies for the H.P. turbine 1st stage disc and turbine blades. The cooling air duct is known as the Tangential On Board Injection (T.O.B.I.) duct. STAGE 1 HPT SUPPORT ASSEMBLY (Coo! air flow) COMBUSTION CHAMBER INNER LINER STAGE 1 HPT VANE CLUSTER ASSEMBLY STAGE 1 HPT DUCT SEGMENT (Cool air flow} STAGE 1 HPT COOLING DUCT ASSEMBLY (Cool air flow) COMBUSTION SYSTEM (3) Mechanical Arrangement H.P. Turbine Shown below is the arrangement of the H.P. Turbine. Cooling airflows are also shown. 10th STAGE COMPRESSOR AIR FOR 2nd VANE AND 1-2 SEAL TOB! FEED TO 1st BLADE 10th STAGE AND HPC DISCHARGE"AIR TO 2nd BLADE HP TURBINE ASSEMBLY Mechanical Arrangement Module 50 - L.P. Turbine The five stage LP turbine extracts energy from the gas stream to provide the rotational drive for the L.P. compressor and fan. The four principal elements of the LP Turbine Module are: • • • • LP Turbine case, vanes and static seals Five stage LP Turbine rotor LP Turbine shaft Turbine exhaust case Seal clearance and L.P turbine case thermal expansion are controlled by an external Active Clearance Control.(A.C.C.) system. The A.C.C. system uses fan discharge air which is directed externally to the L.P. turbine case via the eight A.C.C. tubes. Two boroscope ports are provided on the case, one on each side. These ports enable inspection of the L.P. turbine (stage 3) rotor blades and also stage 2 H.P. turbine rotor blades (rear side). Each port is sealed by a plug which incorporates features to prevent incorrect installation. Axial positioning of the L.P. turbine rotor assembly is achieved by selection of an appropriate adjusting washer fitted at the front end between the L.P. turbine shaft and the L.P. compressor stubshaft. The L.P. turbine shaft is supported at the front by No 2 bearing and at the rear by No 5 bearing. The turbine exhaust case serves to straighten the gas flow, provides structural support for the No 5 bearing and incorporates the rear engine mount lugs. The struts incorporate provision to sense exhaust gas temperature T4.9 and pressure P4.9. REAR ENGINE MOUNT LUGS STAGE 3 NGV LP TURBINE SHAFT LP TURBINE NUT ACC COOUNG AIR INLET TURBINE EXHAUST CASING ADJUSTING WASHER STUBSHAFT SPLINES ACC TUBES LOW PRESSURE TURBINE MODULE Mechanical Arrangement L.P. Turbine Exhaust Case The L.P. Turbine exhaust case provides the support for No 5 bearing. The hollow support struts provide the location for the 4 thermocouples which measure the E.G.T. (T4.9). Three of the struts also house the P4.9 pressure measuring rakes. The casing also provides the rear engine mounting location. OUTER RING ENGINE MOUNTS INNER RING OIL SCAVENGE TUBE FORWARD PRESSURE CONTROL MANIFOLD TURBINE EXHAUST CASE ASSEMBLY TURBINE EXHAUST CASE ASSEMBLY Mechanical Arrangement Nodule 60 - External Gearbox The gearbox assembly transmits power from the engine to provide drives for the accessories mounted on the gearbox front and rear faces. During engine starting the gearbox also transmits power from the pneumatic starter motor to the engine. The gearbox also provides a means of hand cranking the H.P. rotor for maintenance operations. Location The gearbox is mounted by 4 flexible links to the bottom of the fan case. • main gearbox 3 links • angle gearbox 1 link Type Cast aluminium housing. Features • individually replaceable drive units • magnetic chip detectors o main gearbox 2 magnetic chip detectors o angle gearbox 1 magnetic chip detector Front Face Mount Pads • • • • • De-oiler Pneumatic starter Dedicated generator Hydraulic pump Oil Pressure pump Rear Face Mount Pads • • • Fuel pumps and Fuel Metering Unit (FMU) Oil scavenge pumps unit Integrated Drive Generator (I.D.G.) OIL FILTER FUEL PUMP DRIVE PAD O!L SCAVENGE PUMP MAIN GEARBOX DEOILER INTEGRATED DRIVE GENERATOR SYSTEM (iDGS) DRIVE PAD STARTER DRIVE PAD DEDICATED ALTERNATOR (PMA) FRONT VIEW OIL TANK HYDRAULIC PUMP DRIVE PAD OIL PRESSURE PUMP ANGLE AND MAIN GEARBOX Mechanical Arrangement Engine Drain System Leakage and drainage from the engine accessories and fuel operated actuators is conducted by tubes to the engine drains mast. The drains mast discharges to atmosphere through the bottom of the fan cowls. 10 1 9 2 8 3 4 1 : OIL TANK SCUPPER DRAIN 2 : HYDRAULIC PUMP SEAL DRAIN 3 : AIR STARTER MOTOR SEAL DRAIN 4 : IDG SEAL DRAIN 5 : AIR COOLED OIL COOLER ACTUATOR DRAIN 6 : DRAIN MAST 7 : CORE ENGINE DRAINS 8 : FUEL DIVERTER VALVE DRAIN 9 : FUEL PUMP SEAL DRAIN 7 5 GASKET 6 BOLT (TYPICAL 4 PLACES) 10 : FUEL MODULATING DRAIN UNIT ENGINE DRAIN MAST-INSTALLED PART ONE - SECTION 3 ELECTRONIC ENGINE CONTROL (E.E.C.) Electronic Engine Control Introduction The V2500 uses a Full Authority Digital Electronic Engine Control (FADEC). The FADEC comprises the sensors and data input, the electronic engine control unit (E.E.C.) and the output devices which include solenoids, fuel servo operated actuators and pneumatic servo operated devices. The FADEC also includes electrical harnesses. Engine Electronic Control The heart of the FADEC is the Engine Electronic Control (EEC) unit. The EEC is a fan case mounted unit which is shielded and grounded as protection against E.M.I. mainly lightning strikes. Features • • • • • • • Vibration isolation mountings Shielded and grounded (lightning strike protection) Size :15.9 X 20.1 X 4.4 inches Weight :41 lbs Two independent electronic channels Two independent power supplies from dedicated generator Built in handle facilitates removal and handling. HARNESS ANTI-ICE DUCT P2/T2 PROBE SENSOR LINE COOLING AIR OUTLET (2) EEC COOLING EJECTOR EEC-CHANNEL (A) ELECTRICAL CONNECTORS EEC COOLING OUTLET ANTI-VIB MOUNTINGS (4) EEC - CHANNEL (B) ELECTRICAL CONNECTORS COOLING AIR INLETS (2) ELECTRONIC ENGINE CONTROL The EEC Description The E.E.C. has two identical electronic circuits which are identified as Channel A and Channel B. Each channel is supplied with identical data from the aircraft and the engine- This data includes throttle position, aircraft digital data, air pressures, air temperatures, exhaust gas temperatures and rotor speeds. This data is used by the E.E.C, to set the correct engine rating for the flight conditions. The E.E.C. also transmits engine performance data to the aircraft. This data is used in cockpit display, thrust management and condition monitoring systems. Each of the EEC channels can exercise full control of all engine functions. Control alternates between Channel A and Channel B for consecutive flights, the selection of the controlling channel being made automatically by the E.E.C. itself. The dual channels which are contained in the two piece housing are separated from each other through a unique circuit mounting board system. The two channels exchange data through the data crosslink. ENGINE ELECTRONIC CONTROL UNIT EEC DESCRIPTION Electronic Engine Control Harness (electrical) and Pressure Connections Electrical Connections Two identical, but separate electrical harnesses provide the input/output circuits between the EEC and the relevant sensor/control actuator, and the aircraft interface. The harness connectors are 'keyed' to prevent misconnection. Front Face Note: Single pressure signals are directed to pressure transducers - located within the EEC - the pressure transducers then supply digital electronic signals to channels A and B. The following pressures are sensed: • Pamb :ambient air pressure - fan case sensor • Pb :burner pressure P3/T3 probe • P2 :fan inlet pressure - P2/T2 probe • P2.5 :booster stage outlet pressure • P5 (P4.9) :LP Turbine exhaust pressure - P5 (P4.9) rake • P12.5 - fan outlet pressure - fan rake Harness Connector Plug Identification • • • • • J1 J2 J3 J4 J11 E.B.U 4000KSA D202P Engine D202P Engine D203P Engine D204P Engine D211P Rear Face • • • • • • J5 J6 J7 J8 J9 J10 Engine D205P Data Entry Plug E.B.U. 4000 KSB Engine D208P Engine D209P Engine D210P COOLING AIR INLET J5 EEC HARNESS CONNECTORS THIS SIDE TOWARD FNGINE EEC HARNESS CONNECTORS Pamb VIEW A Pb P2.5 VIBRATION - ISOLATED MOUNTS (4) P2 P5 EEC - HARNESS / PRESSURE CONNECTIONS COOLING AIR OUTLET VIEW B Electronic Engine Control Cooling System Internal cooling of the E.E.C. is provided by an induced airflow. The cooling air is drawn in through an inlet in the R-H. fan cowl by an ejector which is located in the cooling air exhaust duct. The ejector uses 7th stage air tapped from the nose cowl anti-icing ducting. The induced airflow is supplemented by ram airflow during forward movement of the aircraft. HARNESS EEC COOLING AIR EXHAUST ANTI-ICE DUCT P2/T2 PROBE SENSOR LINE EEC COOLING EJECTOR SEAL EEC UNiT STARTER DUCT EEC COOLING AIR INLET BRACKET EEC EXHAUST COOLING AIR DUCT EEC COOLING INLET (ON FAN COWL DOOR) ELECTRONIC ENGINE CONTROL-INSTALLATION Engine Electronic Control (E.E.C.) Overview Fault Monitoring The E.E.C. provides the following engine control functions:- The E.E.C- has extensive self test fault isolation logic built in. This logic operates continuously to detect isolate defects in the E-E.C. • • • • • • • • • • • • • • Power Setting (E.P.R.) Acceleration and deceleration times Idle speed governing Overspeed limits (N1 and N2) Fuel flow Variable stator vane system (V.S.V.) Compressor handling bleed valves Booster stage bleed valve (B.S.B.V.) Turbine cooling (10 stage make-up air system) Active clearance control (A.C.C.) Thrust revsrser Automatic engine starting Oil and fuel temperature management Note: The fuel cut off (engine shut down) command comes from the flight crew and is not controlled by the EEC Electronic Engine Control Failures and Redundancy Improved reliability is achieved by utilising dual sensors, dual control channels, dual selectors and dual feedback. • Dual sensors are used to supply all E.E.C. inputs except pressures, (single pressure transducers within the E.E.C. provide signals to each channel A and B). • The E.E.C. uses identical software in each of the two channels. Each channel has its own power supply, processor, programme memory and input/output functions. The mode of operation and the selection of the channel in control is decided by the availability of input signal and output controls. Each channel normally uses its own input signals but each channel can also use input signals from the other channel required i.e. if it recognises faulty, or suspect, inputs. • An output fault in one channel will cause switchover to control from the other channel. • In the event of faults in both channels a predetermined hierarchy decides which channel is more capable of control • In the event of loss of both channels, or loss of electrical power, the systems are designed to go to the fail safe positions. EEC-REDUNDANCY Electronic Engine Control Failures and Redundancy In the event of loss of both input signals, loss of both channels, or loss of electrical power, the system is designed to go to the fail safe positions shown in the table below. • Metering Valve Torque Motor Minimum Fuel Position • Fuel Shut-off Valve Last Commanded Position • Overspeed Valve Solenoid Normal Fuel Metering Seventh Stage Bleed Valves Valve Open Tenth Stage Bleed Valve Valve Open Combined Active Clearance Control unit • High A.C.C. Valve Closed • Low A.C.C. Valve Partially (45%) Open Low Compressor (2.5) Bleed Actuator Valve Open Stator Vane Actuator Vanes Open Fuel Diverter and Back to Tank Valve • Fuel Diverter Valve Solenoid De-energised (Mode 4 or 5) • Fuel Back to Tank Valve Valve closed - No Return to Tank (Mode 3 or 5) Air/Oil Cooler Control Valve Actuator Valve Open Tenth Stage "Make-up" Cooling Air Valve Valve Open * Thrust Reverser Control Unit Reverser Stowed PT2/TT2 Heater Relay Box • Ignition Relays Ignition ON • Probe Heater Relays Heater OFF • Starter Air Valve Valve closed Note: If there is a failure of the thrust reverser control unit arming valve while the reverser is deployed, will remain deployed Operation and Control E.E.C. Power Supplies The electrical supplies for the E.E.C. are normally provided by a dedicated alternator which is mounted to and driven by the external gearbox - shown below. Dedicated Alternator The unit is a permanent magnet alternator which has two independent sets of stator windings and supplies two independent, 3 phase, frequency wild A.C. outputs to the EEC. These unregulated A.C. supplies are rectified to 28 volts DC within the EEC. The dedicated alternator also supplies the N2 signal for the EEC. The EEC also utilises aircraft power to operate some engine systems:• 115 volts AC 400 Hz power is required for the ignition system and inlet probe anti-icing heater • 28V DC is required for some specific functions which include the thrust reverser, fuel on/off and ground test power for EEC maintenance. In the event of a dedicated alternator total failure the EEC is supplied from the aircraft 28V DC bus bars, 28V DC from the same source is also used by the EEC during engine starts until the dedicated alternator comes on line at approximately 10% N2. The dedicated alternator comes on line and supplies the EEC power requirement when the N2 roaches approx 10%. Switching between the aircraft 28V supply and dedicated alternator power supplies is dour automatically by the EEC. SHAFT SUPPORT STATOR DEDICATED ALTERNATOR EEC Power Supplies Electrical Harness The E.E.C. power supplies and speed signal harness connections are as shown below. EEC ELECTRICAL POWER SUPPLIES AND SPEED SIGNALS HARNESS Electronic Engine Control E.E.C. Electrical Connections The harness connections, pin identification and data identification for the E.E.C. junctions 1-11 are listed in the next five pages, These tables may be used in conjunction with the electrical harness diagrams which appear in the relevant engine system description. A. 1 Sine t II »'. t â– IfANHEl 8TOO3W.-O BT004H.M) 1' u i Cosine Return 8T006B20 8T00SW20 T J High Return 8TOO1W2O 8T002B20 h Line A Line B 8T021B20 8T929W29 g t Line A 8TJ12 Line B 8TJ13 Line A Line B 8TO18W2O 8T019B20 8T027W20 8T028B2O p N e M Line A Line B 8T017B20 8T016W20 2 d Line A Line B 8T014H20 8T015B20 Discrete Return L/Eng Discrete R/Eng Discrete Eng D'crete Rt T TRA Exci t .it I- n i i s h «T»f><)W.'(i 8IO/0H.M( !<â– '( t i l n 8T072B20 8T071W?0 Cos i ne Neturn 8T067W20 8T0e8B20 iiigh Return a 8T048B20 8T047W20 Line A ARINC Input ARINC Input From EIU p N e 8T027W20 8T028B20 Line A Line B E n Line A Line B ARINC II Out d 8T059B20 BTO58W2O L K ARINC 112 Out L K 8T056W20 8TO57B20 Line A Line B 8TU6B20 8T115W20 A X External Reset A X 8TU7W20 8T118B20 Discrete Return 8T024B2O 8T022W20 8T023O20 r 3 r 8T066B20 8T064W20 L/Eng D;-5c-...ia R/En§ Discs-ira Eng D 'tiocs Kr at t 8T065O20 z T/R Arming Valve Pressure Switch s. c z T/R Stow and Lock ncnsoi s 8T060B20 8T061W20 8T062W20 8T0fi3B20 E F a 8T051B20 8TC49W2O 8T050O20 R 8T053W20 8T052B20 8T02SW20 8T026B20 Discrete Discrete Rtn Discrete Rtn 6TJ11 Discrete 8T013W20 8T012B20 8TO11B2O 8TO1OW2O On Off Discrete Rtn 8T009B20 8T007W20 8T008O20 E F a 8T02SW20 8T026B20 R n Line A Line B u C Instinctive Disconnect (Autotthrust) ARINC Input From ADI R.I 7 n H.E.C. HARNESS CONECTIONS PIN NO S Line B Di sere â– 'â– .â– â– : 'â– â– •-8TJO3 .* K DL*t.i.^:~. Discra-- on Off Discrete R 8TJ02 Line A Lirse B Line A 8TJ01 Line B PART ONE - SECTION 4 POWER MANAGEMENT Power Management Throttle Control Lever Mechanism The throttle control lever (Thrust Lever) is based on the "fixed throttle" concept, there is no motorised movement of the throttle levers. Each throttle control lever drives dual throttle resolvers, each resolver output is dedicated to one E.E.C. channel. The throttle lever angle (TLA) is the input to the resolver. The resolver output, which is fed to the E.E.C, is known as the Throttle Resolver Angle (TRA) The relationship between the throttle lever angle and the throttle resolver angle is linear and:1 deg TLA = 1.9 deg TRA. CONTROL LEVER REVERSER LATCHING LEVER AUTOTHRUST INSTINCTIVE DISCONNECT PUSHBUTTON ENGINE POWER SETTING Power Management Throttle Control Lever Mechanism The throttle control mechanism for one engine is shown below. The control system consists of:• • • the throttle control lever the mechanical box the throttle control unit The throttle control lever movement is transmitted through a rod to the mechanical box. The mechanical box incorporates 'soft' detents which provides selected engine ratings, it also provides "artificial feel" for the throttle control system. The output from the mechanical box is transmitted by a second rod to the throttle control unit. The throttle control unit incorporates two resolvers and six potentiometers. Each resolver is dedicated to one E.E.C. channel, the output from the potentiometers provides T.L.A. signals to the aircraft flight management computers. A rig pin position is provided on the throttle control unit for rigging the resolvers and potentiometers. Note The E.E.C. incorporates resolver fault accommodation logic which allows engine operation after a failure/loss of T.R.A. signal. Bump Rating Push Button In some cases (optional) the throttle control levers are provided with "Bump" rating push buttons, one per engine. This enables the E.E.C. to be re-rated to provide additional thrust capability for use during specific aircraft operations. REVERSE LATCHING AUTOTHRUST INSTINCTIVE LEVER DISCONNECT PUSHBUTTON CONTROL LEVER REVERSER LATCHING LEVER AUTOTHRUST INSTINCTIVE DISCONNECT PUSHBUTTON THROTTLE CONTROL LEVER MECHANICAL BOX BUMP PUSHBUTTONS ROD THROTTLE CONTROL UNIT ENGINE CONTROL-THROTTLE MECHANISM Power Management Throttle Control Lever Mechanism The throttle control lever moves over a range of 65 degrees, from minus 20 degrees to plus 45 degrees. An intermediate retractable mechanical stop is provided at 0 degrees. Forward Thrust Range The forward thrust range is from (0 to plus 45) degrees. 0 deg. = forward idle power 45 deg. = rated take off power Two detents are provided in this range:at 25 degrees: max climb (MCLB) at 35 degrees: max continuous (MCT)/flexible (deated) take off power (FLTO) Reverse Thrust Range Lifting the reverse latching lever allows the throttle to operate in the range 0 degrees to minus 20 degrees. Two detents are provided in this range:at minus 6 degrees: reverse idle power, corresponds to thrust reverse deploy commanded at minus 20 degrees: max reverse power. Auto Thrust System (A.T.S.) The Auto Thrust System can only be engaged between 0 degrees and plus 35 degrees. ENGINE CONTROL - THROTTLE SETTING Power Management EEC-Fuel System Interface Movement of the pilots throttle control lever is sensed by the dual resolvers which signal the TRA to the EEC. The EEC computes the fuel flow which will produce the required thrust. The computed fuel flow request is converted to an electrical current (I) which drives the torque motor in the Fuel Metering Unit (FMU) which modulates fuel servo pressure to move the Fuel Metering Valve (FMV) and sets the fuel flow. Movement of the FMV is sensed by a dual resolver which is located in the fuel metering unit next to the FMV. The dual resolver translates the fuel metering valve movement into an electrical feedback signal which is fed back to the EEC. The basic control loop is shown below. POWER SETTING - BASIC CONTROL LOOP Power Management Basic Control Loop NOTE The EEC uses closed loop control based on Engine Pressure Ratio (EPR) or N1 if EPR is unobtainable. EEC controls the rate of change of fuel flow, and thus acceleration/deceleration times, as a function of the rate of change of N2. EPR Closed Loop Control The E.E.C. computes a Target EPR as a function of:T.R.A. Tamb T2 Alt Mn : (Throttle Resolver Angle) : (Ambient temperature) : (Engine air inlet temperature) : (Altitude) : (Mach Number) The EPR target is compared to the actual EPR to determine the EPR error. The EPR error is converted to a rate controlled fuel flow command (WF) which is summed with the measured fuel flow (WF actual) to produce the WF error. The W.F. error is converted to a current I which is sent to the FMU to drive the torque motor, this moves the FMV to change the fuel flow. The change in fuel flow causes the engine to accelerate/decelerate and bring about a change in actual EPR. This process continues until there is no EPR error. Nl Reversion Loss of the P2 or P4.9 signal will cause an automatic reversion to Rated N1 closed loop control. If the T2 signal also fails this causes reversion to Unrated N1 closed loop control. NOTE: • • N1 Rated control is dispatchable N1 Unrated control is non dispatchable POWER SETTING -EPR/N1 CLOSED LOOP CONTROL Power Management Thrust Modes The engine operates in one of three thrust modes, AUTO, MEMO and MANUAL. Entering, exiting these three modes is controlled by inputs to the Engine Interface Unit (EIU) In the Memo mode the thrust is frozen, to the last actual EPR value, and will remain frozen until the throttle lever is moved manually, or, auto thrust is reset. Auto Thrust Mode Manual Thrust Mode The Auto Thrust mode is only available between idle and MCT when the aircraft is in flight. This mode is entered anytime the conditions for AUTO or MEMO are not present. In this mode thrust is a function of throttle lever position. After take off the throttle is pulled back to the max climb position, the auto-thrust system will be active and the Automatic Flight system will provide an E.P.R. target to provide either:• • • • max clistib thrust an optimum thrust a minimum thrust an aircraft speed (mach number) in association with the auto pilot. Memo Mode The memo mode is entered automatically, from Auto mode if:• the EPR target is invalid • one of the instinctive disconnect buttons on the throttle is activated • Auto thrust is disconnected by the E.I.U Alpha Floor Protection If an aircraft stall is imminent the Auto Thrust System sets the engine power to T/O regardless of throttle position. POWER MANAGEMENT -THRUST MODE SETTING PART ONE - SECTION 5 FUEL SYSTEM Fuel System Introduction The primary purpose of the fuel system is to provide a completely controlled continuous fuel supply in a form suitable for combustion, to the combustion system. Control of the fuel supply is by the EEC via the FMU. High pressure fuel is also used to provide servo pressure (actuator muscle) for some actuators. The major components of the fuel system include: • • • • • • • high and low pressure fuel pumps (dual unit) fuel-oil heat exchanger LP fuel filter fuel metering unit (FMU) fuel distribution valve fuel injectors (20) fuel diverter and back to tank valve (FDRV) The fuel system is shown schematically below. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. LP fuel pump. LP fuel filter by-pass valve. Differential pressure switch. LP fuel temperature probe. HP fuel pressure relief valve. HP fuel pump. Fuel metering valve (FMV). Overspeed valve (OS) Pressure raising and shut-off valve (PRSOV). Pressure drop governor and spill valve ACC actuator. ACOC air control valve actuator. VSV system actuator. BSBV - Master actuator. BSBV - Slave actuator. V2500 FUEL SYSTEM - SIMPLIFIED OVERVIEW Fuel System Components Description Fuel Pumps H.P.STAGE The fuel pump unit, shown below, consists of low pressure and high pressure stages which are driven from a common gearbox output shaft. Purpose. LP Stage Purpose To provide the necessary pressure increase to: • • • Type account for pressure losses through the FCOC and the LP fuel filter. suppress cavitation. maintain adequate pressure at the inlet to the HP stage Shrouded, radial flow, centrifugal impeller, with an axial inducer. To inciease the fuel pressure to that which will ensure adequate fuel flow and good atomisation at all engine operating conditions. Type Two gear (spur gear) pump. Features • provides mounting for fuel metering unit (F.M.U.) • integral relief valve UNDERSIDE VIEW FORWARD 1 LP pump discharge port 2 LP pump inlet port 3 LP return from FMU 4 HP pump discharge port 5 HP pump inlet port 6 Adapterhousing 7 Accessory drive clamp LP/HP FUEL PUMPS UNIT Fuel System Component Fuel Cooled Oil Cooler Purpose To transfer heat from the oil system to the fuel system to:• • reduce the temperature of the engine lubricating oil prevent fuel ing Type Single fuel pass/multi pass oil flow. Features • • a single casing houses the FCOC and the LP fuel filter. provides location for the fuel diverter and back to tank valve (unit not shown). • fuel differential pressure switch. • fuel temperature thermocouple. 1 2 3 4 5 6 7 8 9 10 11 :Fuel inlet connection :Oil outlet connection :Oil let connection in :Fuel/oil inlet heat exchanger assembly :Fuel inlet connection from FD and RV :Fuel fitter inlet connection from FD and RV :Fuel filter by-pass valve :Fuel filter pressure plate :Fuel drain plug :Fuel outlet connection :Fuel filter element FUEL OIL HEAT EXCHANGER Fuel System Component Low Pressure Fuel Filter Purpose To remove solid contaminants from the fuel. TYPE Woven, glass fibre, disposable, 40 micron (nominal) Features • a differential pressure switch, which generates a flight deck message "Fuel Filter Clog" if the differential pressure across the filter reaches 5 psid. • a by-pass valve which opens and allows fuel to bypass the filter if the differential pressure reaches 15 psid. • a fuel drain plug, used to drain filter case or to obtain fuel samples. FUEL COOLED OIL COOLER END SEAL 1 2 3 4 5 6 7 FUEL PLUG LP FUEL FILTER Fuel Drain plug O-ring Bolt(5 off) Washer (5 off) O-ring Fuel filter element cap assembly Fuel System Component Fuel Metering Unit (F.M.U.) The FMU is the interface between the EEC and the fuel system. It is located on the dual fuel pumps unit, on the rear of the main gearbox, and is retained by four bolts as shown below. All the fuel delivered by the HP fuel pumps which is much more than the engine requires passes to the FMU. The FMU, under the control of the EEC, meters the fuel supply to the fuel spray nozzles, it also supplies HP fuel for the operation (muscle) of a number of actuators. Any fuel supplied by the HP pumps which is not needed for these two uses is returned, from the FMU to the LP side of the fuel system. In addition to the fuel metering function the FMU also houses the Overspeed Valve and the Pressure Raising and Shut Off Valve. The Overspeed valve under the control of the EEC, provides overspeed protection for the LP (N1) and HP (N2) rotors. The Pressure Raising and Shut Off Valve provides a means of isolating the fuel supplies to start and stop the engine. Note: There are no mechanical inputs to, or outputs from, the FMU. 7 1 1 2 3 4 5 6 7 8 9 IP/HP fuef pumps unit Fuel metering unit-FMU LP pump in let LP pump outlet HP pump inlet HP pump outlet-to FMU inlet FMU spill-to LP FMU outlet to ffowmeter EEC-Channel (A) Electrical connectors 10 EEC-Channel (B) Electrical connectors 11 PRSOV Electrical connectors FUEL METERING UNIT DISCONNECT POINTS Fuel System Component Fuel Metering Unit A simplified schematic representation of the Fuel Metering Unit is shown below. The three main functions of the FMU are: 1. metering the fuel supplies to the fuel spray nozzles. 2. overspeed protection for both the LP (N1) and HP (N2) rotors. 3. isolation of fuel supplies for starting / stopping the engine. These three functions are carried out by three valves, arranged in series, as shown: • • • the Fuel Metering Valve (FMV). the Overspeed Valve (OS). the Pressure Raising and Shut Off Valve (PRSOV). The position of each valve is monitored and positional information is transmitted back to the EEC. This ensures that the EEC always knows that the valves are in the commanded position. TO FLIGHT DECK EEC TM TM FEED TO EXT SERVICES HP FUEL PUMP SERVO PRESSURE REGULATOR POSITION RESOLVER FUEL METERING VALVE SPEED VALVE TM M.SW M.SW PRESSURE RAISING AND SHUT-OFF VALVE PRESSURE DROP GOVERNOR AND SPILL VALVE TM = TORQUE MOTOR M.SW = MICROSWiTCH FUEL METERING UNIT SCHEMATIC TO FLOWMETER Fuel System Component Fuel Metering Unit - operation The three major functions: • fuel metering • overspeed protection • pressure raising and shut off are explained in more detail in the following pages. CHANNEL B CHANNEL A TO FUEL DISTRIBUTION VALVE HP FUEL FROM PUMP HP FEED TO EXT SERVOES FMU SPILL AND SERVO RETURN TO LP PUMP INLET FUEL METERING UNIT 1 :SPILL PRESSURE RAISING VALVE 2 :PRESSURE DROPREGULATOR AND SPILL VALVE 3 :FLOW WASH FILTER 4 :SERVO PRESSURE REGULATOR 5 :CONNECTORS 6 :MV TORQUE MOTOR DENSITY ADJUST 7 :POSITION RESOLVER 8 :MICRO SWITCH 9 :DRY DRAINS 10 :LATCHING SHUT-OFF MOTOR 11 :HEAT SHIELD AND EMI COVER 12 :MICRO SWITCH 13 :DRY DRAINS 14 :PRESSURE RAISING AND SHUT OFF VALVE 15 :OVERSPEED VALVE 16 :METERING VALVE 17 :MV BY-PASS ADJUSTER 18 :MIN FLOW ADJUSTER 19 :SERVO SWITCHING VALVE Fuel Metering Operation Fuel metering is achieved by the Fuel Metering Valve and the Pressure Drop Regulator and Spill Valve, which act together in the following sequence. Signals from the EEC, cause the torque motor to change position which directs fuel servo pressure to re-position the Fuel Metering Valve. This changes the size of the metering orifice through which the fuel passes which in turn changes the pressure drop across the metering valve. The change in the pressure drop is sensed by the Pressure Drop Regulator which will re-position the spill valve and so increase/decrease the fuel flow through the fuel metering valve until the pressure drop is restored to its datum value. The increase/decrease in fuel flow causes the engine to accelerate/decelerate until the actual EPR is that demanded by the EEC signal. Movement of the Fuel Metering Valve is transmitted through a rack and pinion mechanism to drive a dual output position resolver, the resolver output is fed back to the EEC. The fuel metering valve incorporates a manual adjuster to allow setting for changes in fuel density. Automatic compensation for changes in fuel temperature is provided by bi-metallic washers located in the pressure drop governor and spill valve assembly. CHANNEL CHANNEL B FUEL METERING VALVE-POSITION FEED-BACK SIGNAL TO EEC TORQUE MOTOR FUEL DENSITY ADJUSTER SERVO PRESSURE REGULATOR POSITION RESOLVER HP FUEL FROM PUMP FLOW WASH FILTER LP SPILL RETURN MIN FLOW STOP PRESSURE DROP REGULATOR AND SPILL VALVE FMU FUEL METERING VALVE PRESSURE DROP REGULATOR AND SPILL VALVE Fuel Metering Unit Overspeed Protection The Overspeed Valve is positioned down stream, in series, with the Fuel Metering Valve. Note: It should be understood that this device is not incorporated to provide the usual N1/N2 red line limiting of max T.o speed of 100% The EEC will act through the Fuel metering Valve to trim the fuel flow if Nl or N2 reach 100%. Operation The overspeed valve is spring loaded to the closed position, it is opened by increasing fuel pressure during engine start and during normal engine operation is always fully open. In the event of an overspeed the EEC sends a signal to the overspeed valve torque motor which changes position and directs H.P. fuel to the top of the overspeed valve this fully closing the valve. A small by-pass flow is arranged around the overspeed valve to prevent engine flame out The overspeed valve is hydraulically latched in the closed position, thus preventing the engine from being reaccelerated - the recommended procedure is for the flight crew to close the engine down, and not re-start. Closing down the engine is the only way to release the hydraulic latching. Note: Because the overspeed valve is spring loaded to the closed position, and opened by fuel pressure, the overspeed valve will close on every engine shut down. CHANNELA CHANNEL B DE-ENERGIZED TORQUE MOTOR ENERGIZED TO PRESSURE RAISING VALVE OVERSPEED VALVE METERING VALVE MIN FLOW ORIFICE PRESSURE DROP REGULATOR AND SPILL VALVE FMU OVERSPEED PROTECTION Fuel Metering Unit Pressure Raising and Shut off Valve The third valve in the F..M.U. is the Pressure Raising and Shut off Valve (PRSOV). Its primary function is to Isolate the fuel supplies to the fuel spray nozzles for starting and stopping the engine, it also acts as a pressure raising valve to ensure that, during engine starts, fuel is not passed to the fuel spray nozzles until fuel pressures in the F.H.U. are high enough to ensure the control devices will function correctly. The two position torque motor which controls H.P. fuel pressure to operate the PRSOV also controls a spill valve servo line. When the torque motor is selected to close the PRSOV, to shut down the engine, the spill valve servo line is opened, this will fully open the spill valve and direct all the H.P. fuel pump delivery back to the L.P. fuel system. The PRSOV torque motor is commanded open by the EEC during AUTO starts. It is commanded open by a cockpit mounted switch during MANUAL starts. The PRSOV can be commanded closed by the EEC during AUTO start sequences if the sequence has to be stopped for any reason. The EEC's ability to close the shut off valve is inhibited above 50% N2. Above 50% N2, and in flight, the PRSOV can only be closed by the crew operated switch in the cockpit. LATCHING SHUT-OFF MOTOR {Shown in open position) CH A CH B MICROSWITCH COCKPIT SHUT-OFF SHUT PX FIXED ORIFICE SPILL VALVE SERVO LP PRV AND SHUT-OFF PISTON (Shown in shut-off position) FMU PRESSURE RAISING AND SHUT-OFF VALVE Component Description Fuel Distributor Valve Purpose To accept the fuel supplied from the FMU and apportion it to the ten fuel manifold tubes. The ten fuel manifolds each supply two fuel spray nozzles. Location At the 4 o'clock position on the front flange of the diffuser casing. Features • • • integral fuel filter with by-pass valve single fuel metering (check) valve o spring closed o fuel pressure opened ten fuel outlet ports REAR OUTER CASING FRONT FLANGE BRACKET REAR MOUNT BRACKET FUEL SUPPLY MANIFOLD SHUT OFF SEAL FUEL INLET FUEL DISTRIBUTION VALVE DIFFUSER CASE FRONT FLANGE FILTER (With bypass) TRANSFER TUBE FUEL SUPPLY TUBES FUEL DISTRIBUTION VALVE DRAIN LINE (8 places) FUEL OUT TO MANIFOLD (10 places) Component Description Fuel Distributor Valve (Cont) The diagram below shows an 'exploded' view of the fuel distributor valve. It also shows how access is gained to the filter for the periodic inspection which is called up in the Aircraft Maintenance Manual. 1 1 2 3 4 5 2 3 : FUEL DISTRIBUTION VALVE : PACKING : FILTER : PACKING : FILTER COVER 4 5 6 7 8 6 : PACKING 7 : WASHER 8 : BOLT 9 : BOLT 10 : FUEL INLET LINE BOLT FUEL FILTER - FUEL DISTRIBUTION VALVE 9 10 Component Description Fuel Manifold To conduct fuel from the fuel distribution valve to the fuel spray nozzles. Location Mounted around the diffuser casing. Type Single pipe with twin branch. Features • • 10 manifolds each manifold supplies two fuel spray nozzles 10 NOZZLE CLUSTERS (2 nozzles each) 20 FUEL NOZZLES IGNITER PLUGS FUEL IN FROM FMU FUEL DISTRIBUTION VALVE VIEWED FROM REAR FUEL DISTRIBUTION TUBES Component Description Fuel Spray Nozzles Purpose To inject the fuel into the combustion chamber in a form suitable for combustion by:• atomizing the fuel • mixing it with air • controlling the spray pattern Location Mounted around the diffuser casing. Type Single orifice - air blast. Features • 20 fuel spray nozzles • inlet fitting houses fuel filter • internal and external heat shields to reduce coking. FUEL FLOW FUEL INLET CONNECTOR CONNECTOR MOUNTING FLANGE FUEL STRAINER ELEMENT INSERT METERING PLUG FAIRING (Both sides) SUPPORT FAIRING FUEL- FUEL- AIR FLOW FAIRING (Both sides) NOZZLE HEAT SHIELD FAIRING BAFFLE FUEL SPRAY NOZZLES (20) NOZZLE Fuel Spray Nozzles Removal/Installation The installation details for one of the fuel spray nozzles is shown below, the other 19 are similar but there are slight differences. Reference must be made to the Engine Maintenance Manual CH 73-13-41. The following points must be observed:• all the gaskets and packing used must be discarded on removal and replaced with new items on re• lubricate the bolt threads with engine oil • observe the torque loading limitations quoted in the Engine Maintenance Manual. • apply anti-galling compound (V10-032) to the shoulders of the end fittings of the fuel supply tubes on re-assembly. • apply white petrolatum (V10-041) to the rubber packings on re-assembly. Note Reference numbers - e.g. V10-032 refer to consumable materials - a full list of these can be found in the Engine Maintenance Manual CH 70-30-00. DIFFUSER CASE BOSS FLANGE K FORWARD GASKE FUEL NOZZLE PACKINGS BRACKET FUEL TRANSFER TUBE BOLTS GASKET NUT FUEL SUPPLY TUBE BOLT CLAMP FUEL SPRAY NOZZLE REMOVAL/INSTALLATION Electrical Harness Fuel Control (F.M.U.) The fuel control electrical harness connections are shown schematically below. 401VC 409VC IDG POWER CABLE 454VC 450VC 452VC 407VC 403VC 457VC 406VC 405VC 408VC 448VC 402VC 404VC 447VC FUEL METERING UNIT FUEL METERING UNIT 1 Fuel metering valve - T/M 2 Fuel metering valve - resolver 3 Overspeed valve - T/M 4 Overspeed valve - M/Sw 5 PRSOV-T/M 6 PRSOV- M/Sw FUEL SYSTEM CONTROL HARNESS Fuel System Electrical Harness, The fuel system electrical harness connections are as shown below. 407VC 403VC 401VC 409VC IDG POWER CABLE 454VC 450VC 452VC D600 1 1 2 3 4 457VC 406VC 405VC 408VC 3 4 2 Fuel Flowmeter Fuel Filter ∆P S/W Fuel Temperature/ TC LP Fuel Filter FUEL SYSTEM HARNESS 448VC 402VC 404VC 447VC PART ONE - SECTION 6 OIL SYSTEM OIL SYSTEM-INTRODUCTION The oil system is a self contained, full flow recirculating type design to ensure reliable lubrication and cooling under all circumstances. Oil cooling is controlled by a dedicated Heat Management System which ensure that engine oil, I.D.G. oil and fuel temperatures are maintained at acceptable levels while ensuring the optimum cooling configuration for the best engine performance. A simplified system diagram is shown below. SYSTEM MONITORING The operation of the engine oil system may be monitored by the following flight deck indications: • engine oil pressure • engine oil temperature • oil tank contents In addition warnings may be given for the following nonnormal conditions: • low oil pressure • scavenge filter clogged or partly clogged (high differential pressure) • No. 4 compartment scavenge valve inoperative V2500 SIMPLIFIED OIL SYSTEM Oil System No 1, 2 and 3 Bearing Lubrication The oil feeds, and scavenge return, for No's 1, 2 and 3 bearings, including oil damping supplies for No 3 bearing are shown in detail below. RESTRICTOR OIL DlSTRIBUTORS STRAINERS STRAINER No 1 BEARING (LP) No 3 BEARING - DAMPER JET TO No 2 BEARING SEAL JET TO No 1 BEARING JET TO HYDRAULIC SEAL JET TO No 3 BEARING CARBON SEAL JET TO BEVEL GEAR MESH SEAL DRAINS ET TO BEVEL GEAR BEARINGS AND STEADY BEARING STEADY BEARING No 1 2 3 BEARING LUBRICATION Oil system No 4 Bearing Lubrication The oil feed, and scavenge, arrangements for No 4 bearing are shown in detail below. STRAINER THERMALLY INSULATED SHIELDING HP COMPRESSOR DELIVERY AIR VIA AIR COOLED AIR COOLER (ACAC) TO No 4 BEARING FAN AIR USED AS COOLANT JET TO No 4 BEARING AND REAR CARBON SEAL JET TO FORWARD CARBON SEAL No 4 BEARING BUFFER AIR FEED (Cooling air) SCAVENGE OIL SCOOP AND STRAINER No 4 BEARING LUBRICATION Oil System No 5 Bearing Lubrication The oil feed and scavenge arrangements for No 5 bearing are shown in detail below. ' LAST CHANCE STRAINER 5 BEARING FEED RESTRICTOR SQUEEZE FILM JET JFT TO No 5 BEARING CARBON SEAL STRAINER OIL SCALLOPS INSULATION BLANKET EXHAUST CASING No 5 BEARING LUBRICATION Oil System Gearbox Lubrication The gearboxes, main and angle, are lubricated and scavenged as shown below. ANGLE GEARBOX MAGNETIC CHIP DETECTOR STRAINER SCAVENGE PUMPS STARTER RIGHT HAND SCAVENGE PICK-UP LEFT HAND GEARBOX SCAVENGE PICK-UP DEDICATED GENERATOR HYDRAULIC PUMP PRESSURE PUMP AND FILTER FUEL PUMPS GEARBOX LUBRICATION Component Description Oil Tank Purpose To store the oil supply. Location Located to the top L.H. side of the external gearbox. Type Pressurised, hot tank. Features • • • • • • oil system servicing: o gravity fill port o prismalite oil level indicator o tank capacity, 28.3 US quarts usable oil 24 US quarts. Internal cyclone type de-aerator Tank pressurization valve (6 psi) ensures adequate pressure at inlet to oil pressure pump. Strainer in tank outlet to pressure pump. Provides mounting for scavenge filter on rear face. Provides mounting for scavenge filter and master magnetic chip detector (MMCD). OIL TANK PRESSURIZATION VALVE OIL TANK BREATHER TUBE PRISMALITE OIL LEVEL INDICATOR OIL TANK CONTENTSTRANSMITTER BREATHER TUBE RETAINING FLANGE MASTER MAGNETIC CHIP DETECTOR MANUAL OIL FILLER SCUPPER DRAIN OIL TANK OIL SCAVENGE FILTER ASSY OIL TANK Component Description Pressure Pump Pressure Filter Purpose Purpose To supply oil under pressure to the engine bearings, gearbox drive, and accessory drives. To trap solid contaminants. Type Mesh type filter - cleanable - nominal 125 micron rating. Standard gear type (speed is 21.4% N2). Type Location Location On the pressure pump housing. Attached to front face of external gearbox on L.H. side just below oil tank. Oil from the tank is taken to the pressure pump by a passage cast within the gearbox. Features Features • provides the housing for the pressure • filter • cold start pressure limiting valve • flow trimming valve • • pressure priming connection anti-drain valve FLOW TRIMMING VALVE (for adjusting oil fow to a standard level during pass-off testing) ANTI-DRAIN VALVE (prevents oil loss when fitter is removed) COLD START PRESSURE LIMITING VALVE (opens at 450 psi) PRESSURE PUMP ' (Driven from gearbox) FILTER ELEMENT (125 microns (ju) filtration) OIL PRESSURE PUMP STRAINER OIL PRESSURE PUMP FILTER UNIT OIL TANK Air/Oil Heat Exchanger (ACOC) Purpose To reject excess heat from the oil system to the fully modulated cooling (fan) air flow, as required by the E.E.C. Type Corrugated fin and tube - double pass. Location Attached to the fan casing on the lower R.H. side. Features • • oil by-pass valve modulated air flow as commanded by E.E.C. (heat management system) Air flow regulated by air control valve. AIR CONTROL VALE OIL IN OIL OUT AIR INLET DUCT Air/Oil Heat Exchanger Oil System Air/Oil Beat Exchanger The cooling air and the oil flows through the air/oil heat exchanger are shown below. MODULATING VALVE (fuel powered and controlled by signal from engine EEC electronic control unit) NACELLE OUTLET PANEL AIR FLOW OIL FLOW Full oil flow at all times, heat rejection controlled by varying air flow BYPASS VALVE (used in event of cooler blockage - opens at 50 psi) AIR AND OIL FLOWS REAR OF FAN CASE Component Description Air/Oil Heat Exchanger Air Modulating Valve Purpose To govern the flow of cooling (fan) air through the air/oil heat exchanger, as commanded by the Heat Management Control System. (EEC) Type Plate type supported at either end by stubshafts. Operated by an Electro-Hydraulic Servo Valve mechanism. Location Bolted to the outlet face of the air/oil heat exchanger. Features • • • fails safe - valve fully open - maximum cooling position fire seal forms an air tight seal between the unit outlet and the cowling orifices. control by either channel A or B of E.E.C. • • • valve position feed back signal via L.V.D.T. to each channel of E.E.C. valve positioned by fuel servo pressure acting on a control piston fuel servo pressure directed by the ElectroHydraulic Servo Valve assembly which incorporates a Torque motor. TORSION SPRING FIRE SEAL VALVE HOUSING FUEL (SERVO) RETURN AIR VALVE STOP FUEL (SERVO) SUPPLY STOP PLATE VIEW ON A LINEAR VARIABLE DIFFERENTIAL TRANSFORMER ASSEMBLY (LVDT) ELECTRO-HYDRAULIC SERVO VALVE ASSEMBLY (EHSV) CHANNEL A ELECTRICAL RECEPTACLES CHANNEL 8 PUSH ROD ACTUATING LEVER AIR MODULATING VALVE Component Description Air Control Valve Electro Hydraulic Servo Valve (E-H.S.V.) Purpose It provides the muscle to move the air control valve to the EEC commanded position. Type two stage directional flow valve • stage 1: electrically activated torque motor and Jet pipe • stage 2: spool valve Location Bolted to the air control valve casing. Features • • • • • two independent torque motor windings -one connected to each channel of EEC. operation, from either channel of EEC. jet pipe protected by 90 micron filter. E.H.S.V. biased to ensure air control valve fully open at engine start condition. single fuel servo supply from FMU. CHANNEL A FUEL SUPPLY FUEL RETURN DRAIN 1 Torque motor 2 Torque motor jet pipe 3 Spool vaive 4 Actuating piston 5 Air valve 6 Spring 7 Push rod and lever 8 LVDT ACOC ELECTRO-HYDRAULIC SERVO VALVE Component Description Fuel/Oil Heat Exchanger Purpose • • cool the engine oil heat the fuel Type Single pass fuel flow multi pass oil flow. Forms an integral unit with the LP fuel filter. Location Bolted to the fan casing - LH side - on the engine centre line. Features Differential pressure relief valve permits oil by-pass if oil is congealed or cooler blocked Anti-syphon drilling Mttrix assy tubes & paffle p|ate Fuel filter Fuel in from LP pump Fuel out to FMU & HP pump Pressure relief valve opens at 50 psi Bypass setction of cooler when mattrix is blocked or when oil is congeealed. FUEL-OIL HEAT EXCHANGER Component Description Scavenge Pumps Unit Purpose Returns scavenge oil to tank. Type Standard gear type pumps (6). All pumps rotate at the same speed (22% N2), pump capacity is determined by the width of the gears. Location All 6 scavenge pumps are housed together as a single unit on rear L.H, side of the gearbox. Features combines the flows from all the pump outlets to return to oil tank DEVELOPED COMPOSITE SECTION THROUGH PUMPS SHOWING FUNCTION OIL FILTER FUEL PUMP DRIVE PAD GEAR PUMPS OIL SCAVENGE PUMP ANGLE GEARBOX 1 STRAINER 6 4 2 3 5 LEFT HAND GEARBOX PICKUP MAGNETIC CHIP DETECTOR STRAINER SCAVENGE PUMPS UNIT RH HAND GEARBOX PICK UP Component Description De-oiler Purpose • • To separate the breather air/oil mixture, it return the oil to the oil scavenge system via its own scavenge pump, and vent the air overboard through the R.H. fan cowl. Type Centrifugal separator. Location Bolted to the front face (R.H. side) of the external gearbox. Features • provides mounting for the No 4 bearing chamber scavenge valve. • provides location for No 4 bearing Magnetic Chip Detector housing. DEOILER OIL TANK OIL PRESSURE PUMP DEOILER Component Description Scavenge Filter Purpose To trap solid contaminants. Type Disposable. Location Mounted to the rear of the oil tank -shown below. Features • • • • disposable. by-pass valve. ∆P switch connections. provides housing for master magnetic chip detector. OIL FILTER PRESURE RELIEF VALVE (Opens at 20 psi pressure drop) FUEL PUMP DRIVE PAD OIL SCAVENGE PUMP ANGLE GEARBOX MASTER MAGNETIC CHIP DETECTOR (Vatric type) SCAVENGE FILTER PRESSURE DROP WARNING SWITCH (∆p 12 psi pressure setting) FILTER ELEMENT (30 microns filtration fine) OIL TEMPERATURE SENSOR (Weston instruments) VIEW ON A OIL SCAVENGE FILTER Component Description No 4 Bearing Scavenge Valve Purpose Maintains No 4 bearing compartment seal deferential air/oil mixture to the de-oiler. Type Pneumatically operated, two position valve; • fully open at low speeds. • closed to minimum area at high engine speeds. Location Mounts on the front face of the de-oiler casing. Features • • • position feed back signal to EEC. uses stage 10 air as servo air. uses value of pressure of stage 10 air as operating parameter: o stage 10 air < 150 psi: valve max open position o stage 10 air > 200 psi: valve min open position 10TH STAGE MAGNETS SCAVENGE FLOW REED SWITCH 4 5 Bolt (3 off) Sealing ring DE-OILER CASE No. 4 BEARING SCAVENGE VALVE ELECTRICAL CONNECTOR Component Description Magnetic Chip Detectors (M.C.D.) A total of 7 MCD's are used in the oil scavenge system. Each bearing compartment and gearbox has its own dedicated M.C.D. (two in the case of the main gearbox) although that for the No 4 bearing is located in the De-oiler scavenge outlet. Access to the dedicated MCD's is by opening the L and R hand fan cowls. The Master M.C.D It is located in the combined scavenge return line, on the scavenge filter housing. The Master MCD is accessible through an access panel in the LH fan cowl. If the master MCD indicates a problem then each of the other MCD's is inspected to indicate the source of the problem. FUEL METERING UNIT HP/LP FUEL PUMPS OIL SCAVENGE PUMPS Nos 1, 2 AND 3 BEARING SCAVENGE CHIP ANGLED GEARBOX SCAVENGE CHIP DETECTOR No 5 BEARING SCAVENGE CHIP DETECTOR OIL SCAVENGE FILTER MASTER CHIP DETECTOR INTEGRATED DRIVE GENERATOR DE-OILER OIL TANK DE-OILER CHIP DETECTOR FORWARD OIL PUMP AND FILTER GEARBOX CHIP DETECTORS STARTER DETECTORS HYDRAULIC CHIP DETECTORS - LOCATION 2 POSITION OIL SCAVENGE VALVE Component Description Master Magnetic Chip Detectors The master magnetic chip detector is located in the scavenge filter case. The master MCD samples the combined scavenge return oil flow. Access to the master MCD is through a dedicated access panel in the LH fan cowl. 1 Magnetic probe 2 Seal ring (2 off) 3 Detector housing MASTER CHIP DETECTOR HOUSING PROBE (With 'O' ring missing) SPRING SAFETY PIN View showing pin against groove when 'O' ring missing) MASTER MAGNETIC CHIP DETECTOR Component Description Magnetic Chip Detectors Location The MCDs for: • • • No's 1, 2 and 3 bearings main gearbox - L.H. scavenge pick up angle gearbox are located to the rear of the main gearbox on the LH side as shown below. Nos 1, 2 & 3 BEARltfG SCAVENGE CHIP DETECTORS ANGLED GEARBOX SCAVENGE CHIP DETECTOR GEARBOX LH CHIP DETECTOR LOCATION MAGNETIC CHIP DETECTORS LEFT-HAND Magnetic Chip Detector Location The MCDs for: • • • No 5 bearing De-oiler (No 4 bearing) Main gearbox - R.H. scavenge pick up are located as shown below. DE OILER CHIP DETECTOR (NO 4 BEARING) No 5 BEARING SCAVENGE CHIP DETECTOR GEARBOX RH CHIP DETECTOR VIEW A VIEW B LOCATION MAGNETIC CHIP DETECTORS RIGHT-HAND OIL SYSTEM INDICATIONS The oil system parameters are displayed on the Engine page on the Lower ECAM screen. 1. OIL TEMPERATURE (deg C) • Normal - green indication • 170 deg C or above: o flashing green indication • 190 deg C or above: o steady amber indication o master caution light o single chime o message (upper ECAM) OIL HI TEMP • Oil low temperature warning: - Throttle above idle and - Engine running o message OIL LO TEMP o single chime 2. Oil Quantity • • Normal - green Less than 4 quarts - flashes green 3. Oil Pressure • • o o o o Normal: green indication 60 psid or below: flashing red indication master warning light f audio warning message (upper ECAM) ENG 1(2) OIL LOW PRESS THROTTLE 1(2) IDLE 4. Scavenge Filter Clog • If filter ∆p > 12 psi: o OIL FILTER CLOG message appears on Engine page. OIL SYSTEM INDICATIONS Component Description Differential Oil Pressure Transmitter & Low Oil Pressure Warning Switch Both units are located on the upper LH side of the fan case. OIL PRESSURE TRANSMITTER LOW OIL PRESSURE WARNING SWITCH OIL PRESSURE TRANSMITTER & LOW OIL PRESSURE WARNING SWITCH Component Description Oil System Electrical Harness The oil system sensors, controls, actuators are connected by the main Electrical harness to the EEC and ECAM computer, as shown below. 1 Low oif pressure switch 2 Lube oil - pressure tx 3 Lube oil temp sensor 4 Lube oif qty tx 5 Scav fitter AP sw 6 IDG oil temp (HMS) 7 ACOC air modulating valve 8 No 4 bearing scav valve 9 No 4 bearing compartment pressure tx 10 Lube oil temp sensor - HMS OIL SYSTEM HARNESS Engine Oil System Pressure Filter Removal/Installation Refer to the AMM Ch 79 Task 79-21-44-000-010 The procedure is summarised below. Note: the pressure filter can be cleaned ultrasonically. OIL PRESSURE PUMP ASSEMBLY 1 Filter casing 2 Lock wire 3 Seal ring 4 Drain plug 5 Bolt 6 Washer 7 Filter cover 8 Seal ring 9 Filter element OIL PRESSURE FILTER REMOVAL/INSTALLATION Oil system Scavenge Filter - R & I Refer to the AMM Ch 79 Task 79-22-44-000-010 Note The scavenge filter is not re-usable. OIL OUTLET (Return to oil tank) MASTER MAGNETIC CHIP DETECTOR 1 OIL INLET (From scavenge pumps) 1 Filter casing 2 Guide pin 3 Lock wire 4 Seal ring 5 Drain plug 6 Bolt 7 Washer 8 Filter cover 9 Seat ring 10 Filter element OIL SCAVENGE FILTER - REMOVAL/INSTALLATION PART ONE - SECTION 7 HEAT MANAGEMENT SYSTEM Beat Management System Purpose Mode 1 The system is designed to provide adeguate cooling, to maintain the critical oil and fuel temperatures within specified limits, whilst minimising the requirement for fan air offtake. This is the normal mode and is shown below. In this mode all the heat from the engine oil system and the I.D.G. oil system is absorbed by the L.P. fuel flows. Some of the fuel is returned to the aircraft tanks where the heat is absorbed or dissipated within the tank. Three sources of cooling are available:• • • the LP fuel passing to the engine fuel system the LP fuel which is returned to the aircraft fuel tanks fan air There are four basic configurations between which the flow paths of fuel in the engine L.P. fuel system are varied. Within each configuration the cooling capacity may be varied by control valves which form the Fuel Diverter and Back to Tank Valve. The transfer between modes of operation is determined by software logic contained in the E.E.C. The logic is generated around the limiting temperatures of the fuel and oil within the system together with the signal from the aircraft which permits/inhibits fuel spill to aircraft tanks. This mode is maintained if the following conditions are satisfied:• • • engine not at high power setting (T/O and early part of climb). cooling spill fuel temperature less than 100 deg C. fuel temperature at pump inlet less than 54 deg C. Heat Management System Mode 3 The second mode shown below is the mode adopted when the requirements for fuel spill back to tank can no longer be satisfied i.e. engine at high power setting. spill fuel temperature above limits (100 deg C tank fuel temperature above limits (54 deg C) In this condition all the heat from the engine and I.D.G. oil systems is absorbed by the burned fuel. If however, the fuel flow is too low to provide adequate cooling the engine oil will be pre-cooled in the air/oil heat exchanger, by a modulated air flow, before passing to the fuel/oil heat exchanger. Heat Management System Mode 4 Mode 4 as shown below is the mode adopted when the burned fuel flow is low i.e. at low engine speeds with a high H.P. fuel pump inlet temperature. In this mode the fuel/oil heat exchanger is operating as a fuel 'cooler' and the heat passed to the engine oil is extracted by the air/oil heat exchanger. Heat Management System Mode 5 Mode 5, shown below, is the mode which is used when system conditions demand operation as in Mode 3 but this is not permitted because:• • the IDG oil system temperature is excessive, or fuel spill to aircraft tank is not permissible because of high spill fuel temperatures. Heat Management System Fuel Diverter and Back to Tank Valve The fuel diverter valve and the back to tank valve together form a single unit. The unit is bolted to the rear of the fuel/oil heat exchanger as shown below. The valves are positioned on commands from the E.E.C. Fuel Diverter Valve This valve is a two position valve and is operated by a dual coil solenoid. The control signals to energise / deenergise the solenoid come from the E.B.C. • • solenoid energised :Mode 4 or 5. Solenoid de-energised : Mode 1 or 3 (fail safe position) Back to Tank Valve This valve is a modulating valve and will divert a proportion of the L.P. fuel back to the aircraft tanks as directed by the E.E.C. The interface between the E.E.C. and the valve is a modulating torque motor, the torque motor (or will direct H.P. servo fuel to position the valve. The fail safe position is with the valve fully closed - no fuel return to tank 10 1 Fuel cooled oil cooler (FCOC) (79-21-43) 2 Fuel filter element (73-12-42) 3 HP servo supply connection 4 LP servo return connection 5 To/from IDG FCOC port connection 6 To/from FCOC connection 7 Fuel filter in ret connection 8 Fuel metering unit spill flow connection 9 FCOC inlet connection 10 Aircraft fuef tank inlet connection 11 Drainconnection 12 ECU LVDT connection (19 pins) 10 ECUT/M • solenoid valve - microswitch connection (26 pinsf FUEL DIVERTER AND RETURN TO TANK VALVE PART ONE - SECTION 8 COMPRESSOR AIRFLOW CONTROL SYSTEM Compressor Airflow Control System Introduction The engine incorporates two air bleed systems and a variable stator vane (VSV) system which together are used to:• ensure stable airflow through the compressor at "off design" conditions • ensure smooth, surge free, accelerations and decelerations (transient conditions) • improve engine starting characteristics • assist in re-stabilising the engine if surge occurs (surge recovery) The complete system comprises three subsystems, which are: • an LP compressor air bleed located at engine station 2.5 and known as the Booster Stage Bleed Valve (BSBV). • HP compressor air bleeds on stages 7 and 10 • the VSV system which comprises variable inlet guide vanes, at the inlet to the H.P. compressor, and 4 stages of variable stator vanes. The three systems are controlled by the E.E.C. AIR-FLOW CONTROL SYSTEM - SCHEMATIC Booster Stage Bleed Valve (B.S.B.V.) Purpose B.S.B.V. Mechanical Arrangement The B.S.B.V. bleeds air from the rear of the L.P. compressor at engine station 2.5, the bleed air is vented into the fan air duct. The annular bleed valve comprises 27 flaps which are attached by 25 link arms and 2 power arms to a synchronise ring. Two actuating rods connect the two power arms to two actuators. The two actuators utilise H.P. fuel as an hydraulic medium, and are hydraulically "linked" to ensure simultaneous movement. The bleed valve provides improved surge margin during starting, low power and transient operations. The bleed valve is controlled by the E.E.C. and is fully modulating, between the fully open and fully closed positions, as a function of:• • • Nl corrected speed Altitude Aircraft forward speed (Mn) For starting the bleed valve is fully open and will progressively close during engine acceleration, during cruise and take off the valve is fully closed. For decelerations and operation in reverse thrust the valve is opened. In the event of an engine surge the valve is opened to enhance recovery. The mechanical arrangement is shown below. 1 2 3 4 5 6 Support ring Power arm (2off) Valve flap Link arm Synchronize ring Bleed duct BOOSTER STAGE BLEED VALVE Compressor Airflow Control BSBV Actuators Description The two BSBV actuators utilise HP fuel as a hydraulic operating medium. The actuators are located on the rear of the intermediate casing on either side of the HP compressor. Only one of the actuators, the one on the left hand side, interfaces with the EEC. This actuator is called the Master actuator, the RH actuator is called the Slave actuator. The two actuators are hydraulically linked by two tubes which pass across the top of the H.P. compressor case. The master actuator incorporates an LVDT which transmits actuator positional information back to the EEC. The slave actuator incorporates two overload relief valves which prevent overpressurisation of the actuators in the case of faults, such as a mechanically seized actuator. 1 2 3 4 5 Bleed valve actuating rod Piston jack fork end LPC bleed-master actuator LPC Weed-slave actuator Intermediate structure BSBV ACTUATORS BSBV Master Actuator Removal / Installation Removal/installation of the master actuator is quite straightforward. The following points should be noted:all sealing rings must be discarded on removal and new sealing rings fitted on installation all threads should be lubricated with clean engine oil on installation observe the torque maintenance manual loadings quoted in the the bolt which secures the actuator fork end to the actuating rod is locked by a double key washer a new washer must be used on installation on completion of the actuator change carry out Test No. 1 or 3 (leak test), followed by Test No. 11 (High Power Assurance test). The full procedure to remove/install the B.S.B.V. master actuator can be found in the AMM CH 7531-42 EXTEND LINE TO SLAVE ACTUATOR RETRACT LINE TO SLAVE ACTUATOR GUIDE PIN (2 off) HP FUEL BLEED VALVE ACTUATING ROD INTERMEDIATE STRUCTURE DRAIN TUBE HARNESS (Feed back signals) HARNESS BSBV MASTER ACTUATOR REMOVAL / INSTALLATION Compressor Airflow Control BSBV Slave Actuator Removal/Installation Removal/installation straightforward. of the slave actuator is quite The disconnect points are shown below. Points to note are the same as the notes for the Master actuator shown on the previous page. For full removal/installation procedures refer to the aircraft maintenance manual CH 75-31-43. INTERMEDIATE STRUCTURE RETRACT LINE EXTEND LINE PISTON JACK FORK END BLEED VALVE ACTUATING ROD FUEL DRAIN GUIDE PIN (2 off) MOUNT BRACKET BSBV SLAVE ACTUATOR REMOVAL/INSTALLATION Compressor Airflow Control BSBV Electrical Harness The BSBV electrical harness connections are as shown below. BSBV (2.5) BLEED ACTUATOR HARNESS Compressor Airflow Control Variable Stator Vane System (V.S.V.) Introduction The entry of air into the H.P. compressor is controlled by Variable Incidence Stator Vanes. The variable vanes control the angle at which the air enters the first five stages of the H.P. compressor. The five unison rings are connected by short rods to a crankshaft. The crankshaft is connected by a short rod to an actuator which utilizes HP fuel as a hydraulic operating medium. The angle varies with the HP compressor speed (N2), this reduces the risk of blade stall and compressor surge. Signals from the EEC direct HP fuel to extend/retract the actuator. Actuator movement causes the crankshaft to rotate, and, through the unison rings, reposition the variable stator vanes. The five stages of variable incidence stators comprise inlet guide vanes to stage 3 and stages 3, 4, 5 and 6 stator vanes. Mechanical Arrangement Each vane has pivots at its inner and outer ends which allow the vane to rotate about its longitudinal axis. The outer end of each vane is formed into a shaft which passes through the compressor case and is attached by a short lever to a Unison ring, (one unison ring for each stage). The actuator incorporates an LVDT which signals actuator positional information back to the E.E.C. UNISON RINGS FUEL POWERED RAM VARIABLE STATOR VANE ACTUATION CRANKSHAFT AND MOUNTINGS FRONT BRG HOUSING IGV CRANK SHAFT LEVER ASSEMBLY STAGE 3 LEVER STAGE 5 LEVER BRIDGE PIECE REAR BRG ENGINE SPLIT CASING FUEL POWERED RAM STAGE 4 LEVER INPUT AND STAGE 4 UNDERSIDEVIEW OF ENGINE HP COMPRESSOR VSV ACTUATION SYSTEM Compressor Airflow Control Variable Stator Vane System (S.V.S.) Actuator Removal/Installation The following notes summarize the removal / installation procedures. For a full description of the procedure refer to the aircraft maintenance manual CH 75-32-41. Access to the actuator, which is mounted on the HP compressor case, L.H. lower side, is by opening the L.H. 'C duct, see Part Two, Section Two of these notes and refer to the aircraft maintenance manual CH 78-32-00. Note • The VSV actuator fuel pressure and return lines must be drained before work begins. • The fuel lines are drained at the union locations shown below. • The fuel is drained at this point because it is the lowest point in the system, and also, because fuel drained from here is least likely to contaminate the engine electrical harness. FUEL METERING UNIT (FMU) 1 UNION 2 UNION VSV ACTUATOR-FUEL DRAIN UNIONS Compressor Airflow Control Variable Stator Vane System (VSV) Actuator Removal/Installation (Cont) Before the actuator is removed it is important that the VSV crankshaft assembly is locked in order to prevent damage to the stator vanes. Rig pins are provided to lock the crankshaft and the actuator, as shown below. After the fuel supply and return tubes have been disconnected the crankshaft should be rotated to align the rig pin holes in the input lever and the front bearing housing. Spanner (Wrench) flats are provided on the crankshaft for this purpose. Installing the rig pin locks the crankshaft assembly with the actuator and vanes in the high speed position (actuator fully retracted). VSV ACTUATOR REMOVAL/INSTALLATION (3) Compressor Airflow Control Variable Stator Vane System (V.S.V.) Actuator Removal/Installation (Cont) The actuator disconnect points are shown below. Note: Discard the sealing rings, from the fuel lines, on removal • • fuel tubes are lock-wired a 15/16 inch bi-hex crowsfoot spanner is required to disconnect the fuel supply tube ELECTRICAL CONNECTOR VARIABLE STATOR VANE ACTUATOR ELECTRICAL CONNECTOR LP RETURN TUBE SEALING RING HP SUPPLY TUBE ELECTRICAL CONNECTOR FUEL DRAIN TUBE ELECTRICAL CONNECTOR VSV ACTUATOR - REMOVAL/INSTALLATION (1) Variable Stator Vane System (V.S.V.) Actuator Removal/Installation (Cont) The following points should be noted:• During removal do not allow the upper support bracket to take the full weight of the actuator since this can damage the bracket. • The surfaces marked * should be cleaned with V01003 (cleaning fluid) and coated with V04-0O4 (jointing compound) on assembly. • Ensure the torque loading instructions contained in the aircraft maintenance manual are carried out. • Fit the rigging pin to the replacement actuator, with the actuator in the fully retracted (high speed) position before installation. • If it is necessary to adjust the length of the control rod end ensure the control rod ends are in "safety" on completion. • On completion carry out Test No 1 or 3 (leak checks) followed by Test No 11 (high powor assurance test). VSV ACTUATOR - REMOVAL/INSTALLATION (2) Compressor Airflow Control V.S.V. System - Electrical Harness The VSV Electrical harness connect ions are as shown below. VSV ACTUATOR HARNESS Compressor Airflow Control Handling Bleed Valves Introduction Handling bleed valves are fitted to the H.P. compressor to improve engine starting, and prevent engine surge when the compressor is operating at off-design conditions. A total of four bleed valves are used, three on stage 7 and one on stage 10. The handling bleed valves are two position only - fully open or fully closed, and are operated pneumatically by their respective solenoid control valve. The solenoid control valves are scheduled by the E.E.C. as a function of N2 and T2.6 (N2 corrected). When the bleed valves are open air bleeds into the fan duct through ports in the inner barrel of the 'C ducts. The servo air used to operate the bleed valves is H.P. compressor delivery air known as PB or Pb. The bleed valves are arranged radially around the HP compressor case as shown below. Silencers ate used on some bleed valves. All the bleed valves are spring loaded to the open position and so will always be in the correct position (open) for starting. COMPRESSOR HANDLING BLEED VALVES Compressor Airflow Control Handling Bleed Valves - Location R.H. The diagram below shows the location of the three bleed valves mounted on the R.H. side of the engine H.P. compressor case. STAGE 7 UPPER RH BLEED VALVE STAGE 10 RH BLEED VALVE STAGE 7 LOWER RH BV BLEED VALVES - RH Compressor Airflow Control Handling Bleed Valves - Location L.H. The diagram below shows the location of the bleed valve mounted on the L.H. side of the engine. BLEED VALVES - LH Compressor Airflow Control Handling Bleed Valves Operating Schedule The handling bleed valves have tluec operating regimes:• • • steady state transient surge/reverse Operation of the bleed valve is scheduled against N2 corrected for changes of H.P. compressor (T2.6) inlet temperature -known as N2C26. Steady State The valves are commanded open whenever N2C26 is below the steady state closing speed. Transient The valves are commanded open at the beginning of accelerations/decelerations and will close when either the speed limits are exceeded or timers expire. Surge/Reverse The valves will be commanded open in the event of a surge within the speed range shown. In reverse thrust laws similar to the transient laws apply but the reverse speeds, shown below, are used. BLEED VALVE REGIME OPEN CLOSE Corrected N2 9000 9400 below 15000 ft 10000 10400 above 15000 ft Transient 9000 9400 Surge/Reverse 12562 12772 7B Steady state 7650 8000 7C Steady state 6800 7000 Transient 11800 12250 below 15000 ft 12000 12450 above 15000 ft Surge/Reverse 12352 12562 Steady state 7650 8000 Surge/Reverse 10667 10667 7A 10 Steady state BLEED VALVE OPERATING SCHEDULE Compressor Airflow Control Handling Bleed Valves operating Schedule The schedule for one bleed valve (7C) is shown, in detail, below. Note the transient regime is slightly modified for operation above 15000 ft but operates in the same way. Steady State Surge/Reverse It can be seen that the valve will be commanded closed at stabilized min. idle, 8600 N2C26, and will not be opened again in Steady state. If the engine is operating in reverse thrust operation is the same as Transient but different speeds apply. Transient The valve will be commanded open during engine acceleration whenever N2C26 is below the transient closing speed. Thus during an acceleration from min idle to max speed the valve will be opened and will remain open until the speed passes the transient closing speed. If the acceleration is to a speed below the transient closing speed the valve will remain open until the acceleration timer expires (30 seconds). During decelerations the valve will be commanded open whenever N2C26 is below the transient opening speed. The valve remains open until the deceleration ceases and a deceleration time, 2 seconds, expires. In the event of an engine surge the valve will be commanded open, if the speed is below the open speed, and will remain open until the engine restabilises. BLEED VALVE OPERATING SCHEDULE (7C) Compressor Airflow Control Handling Bleed Valves - Operation The bleed valves and the solenoid control valves all operate in the same manner. The operation of one bleed valve only is described. Bleed Valves The bleed valve is a two position valve and is either fully open or fully closed. The bleed valve is spring loaded to the open position and so all the bleed valves will be in the correct position - open -for engine start. When the engine is started the bleed air will try to close the valve. The valve is kept in the open position by servo air (P3) supplied from the solenoid control valve, (solenoid de-energised) as shown below. P3 SUPPLY VENT VENT SOLENOID DE-ENERGISED OPENING CHAMBER BLEED VALVE OPEN SOLENOID CONTROL VALVE BLEED VALVE - OPERATION Compressor Airflow Control Handling Bleed Valves - Operations The bleed valves will be closed at the correct time during an engine acceleration by the EEC. The bleed valve is closed by the EEC which energises the solenoid control valve, as shown below. Energising the solenoid control valve vents the P3 servo air from the opening chamber of the bleed valve, and the bleed valve will move to the closed position. During an engine deceleration the reverse operation occurs and the bleed valve opens. P3 SUPPLY VENT VENT SOLENOID ENERGISED BLEED VALVE CLOSED SOLENOID CONTROL VALVE BLEED VALVE - OPERATION PART ONE - SECTION 9 SECONDARY AIR SYSTEMS ACTIVE CLEARANCE CONTROL SYSTEM 10TH STAGE HAKE-UP AIR SYSTEM AIRCRAFT SERVICES BLEED SYSTEM Secondary Air Systems Active Clearance Control (A.C.C.) System Introduction The system improves engine performance by ensuring that the HP and LP turbines operate with the optimum turbine blade tip clearances. This is achieved by directing a controlled flow of cooling air to reduce the thermal growth of the turbine casings. This minimises the increase in turbine blade tip clearances which otherwise occurs during the climb and cruise phases. Operation An air scoop directs fan air to a dual control valve which modulates the flow to two cooling manifolds, on the H.P. and L.P. turbine casings. The modulating air control valves are positioned by a fuel pressure operated actuator. The actuator input to the air control valves is through a cam mechanism which provides different cooling flow rates to the two separate manifolds. The actuator is positioned by signals from the E.E.C. which thus controls the cooling flows as a function of:• • corrected N2 aircraft altitude An actuator mounted L.V.D.T. transmits cooling valve position feedback signals to the E.E.C. Loss of control (EEC) or loss of fuel pressure drives the actuator to the fail safe position to provide maximum turbine blade tip clearances. ACC SYSTEM - SCHEMATIC Secondary Air Systems Active Clearance Control Component Location The components in the system comprise:• • • • the HP Turbine cooling manifold the LP Turbine cooling manifold the dual air control valve the actuator ACTIVE CLEARANCE CONTROL (ACC) ACTUATOR ACTIVE CLEARANCE CONTROL (ACC) VALVE AIR FROM THRUST REVERSER DUCT TO HP/LP CASE COOLING HP TURBINE CASE COOLING LP TURBINE CASE COOLING FROM RH T/R DOOR ACTIVE CLEARANCE CONTROL-INSTALLATION Secondary Air Systems Active Clearance Control Operation The A.C.C. system is shown diagrammatically below. Operation of the system is summarized as follows:• • • • • signals from the E.E.C. channel A or B position the jet pipe servo valve the jet pipe servo valve moves to direct HP fuel to one end of the spool valve the spool valve is positioned to port HP fuel to one side of the actuator piston, the other side of the piston is ported to LP fuel return the actuator extends/retracts to position the L.P. and HP cooling air control valves to the commanded position actuator movement is sensed by the LVDT which signals actuator position to the E.E.C, channels A and B. JET PIPE SERVO VALVE 90 µ FILTER EEC PRESSURE PORT DUAL LVT FEEDBACK TRANSDUCER RETURN PORT DRAIN HP Turbine LP Turbine ACC SYSTEM Secondary Air Systems Active Clearance Control H.P. Turbine Manifold The assembly consists of LH and RH tube assemblies which are a simple push fit into the manifold. The tube assemblies are sealed off at their upper ends. Air from the air control valve enters the manifold and is directed to the left and right tubes. Air outlet holes on the inner face of the tubes direct the air onto the H.P. turbine casings. HP TURBINE - ACC MANIFOLD Secondary Air Systems Active Clearance Control L.P. Turbine Manifold The assembly consists of upper and lower tube assemblies with integral manifolds, both ends of the cooling tubes are sealed. Air from the air control valve enters a supply tube which then splits to feed air into two tubes which supply the upper and lower manifolds. The manifolds direct the air into the cooling air tubes. Air outlet holes on the inner surfaces direct the air onto the L.P. turbine cases. LP TURBINE - ACC MANIFOLD Secondary Clearance Control Operating Schedule The graph shown below shows control valve position, and actuator position related to operation points A to E. Engine Stopped With the engine stopped, the position of the actuator piston is point A. At this point:• The control valve for the H.P. turbine ACC is closed. • The control valve for the LP turbine ACC is not less than 44 per cent opened. Engine Operation During engine operation, the E.E.C. controls the position of the actuator piston between point B and point E. Take-off During take-off, the position of the actuator piston is at point C. At this point:• The control valve for the HP turbine ACC is closed. • The control valve for the LP turbine ACC is not less than 70 per cent opened. Note The actuator position between point C and point E depends on Altitude. Fail Safe When there is no torque motor current or no fuel servo pressure, the actuator piston moves to point A. The actuator piston remains at this point at all defective conditions. PISTON TRAVEL (%) ACC OPERATING SCHEDULE Secondary Air Systems Active Clearance Control The ACC Electrical harness connections are as shown below. HPT/LPT ACC HARNESS Secondary Air Systems 10th Stage Make-up Air System Introduction The purpose of this system is to provide additional cooling airflows to the H.P. turbine 2nd stage disc and blades. The cooling air used is taken from the 10th stage manifold, and is controlled by a two position pneumatically operated valve. The valve position is controlled by the EEC as a function of corrected N2 and altitude. Operation Signals from the EEC will energise/de-energise the solenoid control valve. This directs pneumatic servo supplies to position the 10th stage air valve to the open/close position. In the open position the valve allows 10th stage air to flow through two outlet tubes down the left and right hand side of the diffuser case and then pass into the engine across the diffuser area. The air then discharges into the area around No4 bearing housing. The make up air supplements the normal airflows in this area and increases the cooling flow passing to the H.P. turbine, stage 2. The EEC will keep the air valve open at all engine operating phases except cruise. The valve incorporates a micro switch when transmitting valve position feedback signals to the EEC. The failsafe position is valve open. MAKE-UP AIR VALVE - OPERATION MAKE-UP AIR SYSTEM -SCHEMATIC Secondary Air Systems 10th Stage Make-up Air System Component Location The components in this system comprise:• • • the two-position stage 10 on-off valve bolted to the 10th stage manifold at the top of the engine compressor case. the solenoid control valve located on the lower RH fan case. two air supply tubes. TURBINE COOLING CONTROLLED AIR TUBES ON/OF VALVE SOLENOID CONTROL VALVE VIEW ON A MAKE-UP AIR SYSTEM - COMPONENT LOCATION Secondary Air Systems 10th Stage 'Make-up' Air System 10th Stage Air Valve Removal/Installation These notes are for guidance only, reference must be made to the Maintenance Manual -75-23-51. Access to the valve is by opening the 'C ducts. The disconnection and location points are shown below. The removal and installation procedure is straightforward but the following points should be noted:• • a new 'C seal (3) must be fitted. the threads of the retaining bolts (4) must be cleaned and coated with anti-seize compound, 1 2 3 4 5 6 7 Electrica (connector) Outlet air tube C-seal Bolt Air Offtake to No. 4 Bearing Scavenge Valve Lock wire Servo air tube 10 STAGE AIR VALVE REMOVAL/INSTALLATION Secondary Air Systems Engine Air Bleeds - Aircraft Services The system provides pressure/flows for:• • • • • cabin pressurisation and conditioning wing anti-icing engine crossfeed starting hydraulic system pressurisation water system pressurisation The required Sir is bled from the HP compressor of each engine. APU BLEED VALVE TO WING ANTI-ICING CROSS8LEED VALVE TO AIR CONDITIONING PACKS TO WING ANTI-ICING HP GROUND CONNECTOR PRECOOLER FAN AIR VALVE OVER PRESSURE VALVE IP CHECK VALVE PRESSURE REGULATING VALVE HP VALVE AIRCRAFT PNEUMATIC SYSTEM MANIFOLD Secondary Air System Aircraft Pneumatic System The aircraft pneumatic system is shown, schematically, below. No 1 Engine installation only is shown, No 2 engine is identical. AIRCRAFT PNEUMATIC SYSTEM Secondary Air Systems Engine Air Bleeds - Aircraft Services The system is identical on each engine. One engine system is shown, schematically below. Two air off-takes are provided, one each from:. stage 7 . stage 10 Air is taken from stage 10 or stage 7. Stage 10 supplies the requirements at low power setting. Stage 7 supplies the requirements at higher power settings. Automatic change over from stage 10 to stage 7 occurs during engine acceleration. TO AIRCRAFT PNEUMATIC SYSTEM TO NOSE COWL TAI OVERBOARD STARTER AIR VALVE STARTER STAGE 7 STAGE 10 HPC AIR OFF-TAKES - SCHEMATIC 1 2 3 4 5 HP Regulating valve PRV 0/PV FAV Check valve Secondary Air Systems Engine Air Bleeds - Aircraft Services Component Location The engine mounted components of the aircraft pneumatic system comprise:• • • • high stage control valve pressure regulating valve check valve associated ducting up to the engine / aircraft interface All these components are located on the L.H. side of the engine core as shown below. The remainder of the components of the system are supply, fit and responsibility of the aircraft manufacturer. STARTER DUCT OVER/PRESSURE VALVE PRECOOLER INLET COWL ANTI-ICE DUCT FAN AIR 7TH STAGE CHECK VALVE ANTI-ICE VALVE STARTER DUCT REGULATING VALVE HIGH STAGE CONTROL VALVE CORE COMPARTMENT TEMPERATURE SENSOR NACELLE - PNEUMATIC SYSTEM - COMPONENTS Secondary Air Systems Engine Air Bleeds - Aircraft Services The system operates under the control of the Bleed Air Monitoring Computer (BMC) and will automatically:• • • select the compressor stage from which air is bled. regulate bleed air pressure. control bleed air temperature. TCT CTS A/C SYSTEMS TLT PRECOOLER PNEUMATIC SYSTEM SCHEMATIC Secondary Air Systems Engine Air Bleeds - Aircraft Services Sensing Lines The arrangement of the pneumatic sensing lines is shown below. PRESSURE REGULATING VALVE CHECK VALVE HIGH PRESSURE CONTROL VALVE PNEUMATIC SYSTEM COMPONENTS-SENSE LINES Secondary Air Systems Engine Air Bleeds - Aircraft Services The air bleed electrical harness connections are as shown below. 402VC 404VC 447VC DUAL OUTPUT 454VC 450VC PRESSURE REGULATING VALVE V HIGH STAGE VALVE D510P(A) D550P(B) HYDRAULIC LP WNG SWITCH MISC SYSTEMS HARNESS ANTI-ICING CONTROL VALVE Secondary Air Systems Engine Air Bleeds - Aircraft Services High Stage Control Valve Removal / Installation Refer to M.M. Removal Refer to AMM, Task 36-11-51-000-010 Installation Refer to AMM, Task 36-11-51-400-010 The procedure is illustrated below. Note Discard the seals on removal and always fit new seals on installation ELECTRICAL CONNECTOR MANUAL LOCKOUT PIN PRESSURE FROM PRV CLOSED POSITION OPEN POSITION POSITION INDICATOR HIGH PRESSURE VALVE - REMOVAL/INSTALLATION Secondary Air Systems Engine Air Bleeds - Aircraft Services Pressure Regulating Valve R & I Removal Refer to AMM, Task 36-11-52-000-010 Installation Refer to AMM, Task 36-11-52-040-010 Note Discard the seals on removal and always fit new seals on installation PRESSURE FROM TEMPERATURE LIMITATION THERMOSTAT (TLT) ELECTRICAL CONNECTOR PRESSURE TO HIGH PRESSURE VALVE MANUAL LOCKOUT PIN (STOWED) MANUAL LOCKOUT PIN OPEN ROTATED VIEW 180° POSITION INDICATOR PRESSURE REGULATOR VALVE - REMOVAL/INSTALLATION PART ONE - SECTION 10 ENGINE ANTI-ICE SYSTEMS Engine Anti-Icing System General P2/T2 Probe Heating Ice may form in the inlet cowl when the engine is operating in conditions of low temperature and high humidity. The P2/T2 probe is continuously heated, during engine operation, by an integral 115V heating coil. Ice build up in, and on, the inlet cowl could affect engine performance and could cause compressor damage from ice ingestion. To prevent ice formation anti-icing protection is provided in the following areas:- Rotating Fairing • • • the inlet cowl (thermal). the P2/T2 probe (thermal) the rotating fairing - spinner (dynamic) Inlet Cowl Anti-Icing The inlet cowl anti-icing system utilises air taken from the 7th compressor stage which is ducted down the R.H. side of the engine to an anti-icing control valve located on the rear diaphragm of the nose cowl. From the anti-icing control valve the hot air passes to a distribution manifold located inside the inlet cowl lip. The used air is vented overboard through an exhaust grille on the lower R.H. side of the inlet cowl. The rotating fairing (spinner) is protected against ice build up by a solid rubber nose tip which vibrates naturally to break up and dislodge the ice immediately it starts to form. ANTIICE ENG 1 7TH STAGE DUCT INLET COWL ANTI-ICE VALVE ANTI ICE DISCHARGE GRILLE ANTI-ICE SYSTEM - INSTALLATION ENG2 Engine Anti-Icing System Inlet Cowl Ducting Hot air from the anti-icing control valve is ducted between the inner and outer skins of the inlet cowl to a spray ring located inside the inlet cowl lip skin. The spray ring has multiple outlets which direct the hot air to heat the inner surfaces of inlet cowl lip skin. The air is then exhausted overboard through an outlet grille on the lower R.H. side of the inlet cowl. The arrangement is shown below. FORWARD BULKHEAD AFT BULKHEAD FORWARD BULKHEAD AIR SUPPLY EXHAUST DUCT DISCHARGE VIEW ON A INLET COWL ANTI-ICE DUCT Engine Anti-Icing System Anti-Icing Control Valve The anti-icing control valve is located in the air supply ducting, as shown below, which is attached to the rear of the inlet cowl. Manual Override and Lock The valve can be manually overridden and locked in either the open or closed position. Without inlet pressure applied to the unit, the valve can be manually moved to the desired position by applying a standard wrench to the hexagonal nut attached to the butterfly shaft. The valve is locked in the selected position by removing the locking pin from its stow position and inserting it through the locking hole on the valve exterior and the mating hole in the valve piston. The pin is retained by a ball detent mechanism in the end of the locking pin. LOCKING PIN OP (OPEN) RING MID CL CLOSE LOCK RING INLET COWL ANTI-ICE VALVE - REMOVAL/INSTALLATION Engine Anti-Icing System Anti-Icing Control Valve Description and operation The anti-icing control valve is a butterfly type valve located in the anti-icing supply duct. The valve is positioned through a mechanical linkage by a piston which is operated by pneumatic pressure. The piston is spring loaded to the valve closed position. The two valve positions are shown below. ANTI-ICING OFF Solenoid Energised Inlet supply pressurised When the solenoid is energised, the solenoid plunger is retracted against the spring force. Inlet pressure shuttles the pilot ball against the pilot vent seat, thus admitting inlet pressure to chamber "A". Inlet pressure acting on the chamber "A" side of the piston, overcomes the opposing force of inlet pressure acting on the chamber "B" side of the piston which has less effective area. This pneumatic closing force, aided by the power spring force, moves the valve to the closed position. ANTI-ICING ON Solenoid De-energised Inlet supply pressurised When the solenoid is de-energised, the solenoid plunger is extended by the spring. The pilot ball is held against the inlet pressure seat on the pilot valve, thus venting chamber "A" to ambient. Inlet pressure in chamber "B", acting on the reduced piston area, overcomes the power spring force and moves the valve to the open position. DUCT PRESSURE ANTI-ICE OFF SOLENOID ENERGISED VALVE CLOSED ANTI ICE ON SOLENOID DE ENERGISED VALVE OPEN (Fail safe position) INLET COWL ANTI-ICE VALVE Engine Anti-Icing System Inlet Cowl Ducting The ducting which connects the anti-icing control valve to the inlet cowl ducting is shown below INLET COWL ANTI-ICE DUCTING - RH NLET COWL PART ONE - SECTION 11 ENGINE INDICATING Engine Indicating Flight Deck Indications The primary engine parameters listed below are displayed on the upper ECAM CRT. Primary Engine Display • • • • Engine Pressure Patio (EPR) Exhaust Gas Temperature (EGT) N1 N2 Secondary Engine Display The secondary engine parameters listed below are available for display on command or, during engine start, automatically. • • • • • • • • • Total fuel used Oil quantity Oil pressure Oil temperature Nacelle temperature Vibration – N1 and N2 Oil filter clog Fuel filter clog No 4 bearing Scavenge Valve position FLIGHT DECK INDICATIONS Engine Indicating Speed Indicating System Purpose To provide signals of N1 & N2 speeds to be used for:• • flight deck indications the EEC control circuits Note In addition to the speed signals a dedicated signal from the LP rotor (N1) is provided for trim balancing operations. Type N1: speed probes used in conjunction with a phonic wheel N2: uses the frequency of the dedicated alternator output Trim Balance: speed probe (used with the phonic wheel as N1). Operation A schematic arrangement of the speed indicating system is shown below. FRONT BEARING COMPARTMENT BREAK CONNECTION (Bifurcation panel) SPEED PROBE 'A' SPEED PROBE 'B' ARINC 429 RMS EEC SPEED PROBE (Spare) N1 SIGNALS N1 and N2 signals) TRIM BALANCE PROBE TO ELECTRONIC CENTRALIZED AIRCRAFT MONITORING SYSTEM (ECAM) AIRCRAFT/ ENGINE INTERFACE JUNCTION BOX MAIN ACCESSORY GEARBOX CHANNEL A' POWER SUPPLY CHANNEL B' POWER SUPPLY DEDICATED GENERATOR N2 SIGNALS ENGINE SPEED MEASUREMENT -SCHEMATIC TO ENGINE VIBRATION MONITORING UNIT (EVMU) Engine Indicating Speed Indicating System Speed Probes (N1) The probes comprise two pole pieces, a permanent magnet, and a coil wound on to one of the pole pieces. The pole pieces span two teeth of the phonic wheel. The phonic wheel is an integral part of the fan stubshaft and has 60 teeth. As the shaft rotates and the teeth of the phonic wheel pass the pole pieces, a voltage 'pulse' is produced in the winding. The number of pulses produced is directly proportional to the speed of the shaft. This signal is passed to the E.E.C., processed and is used to display N1 speed on the flight deck and also for the engine control circuits as required. Trim Balance Probe The signal from this probe is only used during trim balance operations and provides the phase relationship between any out of balance forces present and a datum position. The trim balance probe senses the passage of one specially modified tooth on the phonic wheel and produces one pulse per revolution. POLE PIECES LP STUB SHAFT SPEED PROBE PERMANENT MAGNET FAN SPEED PROBE PHONIC WHEEL PERMANENT MAGNET POLE PIECES TRIM BALANCE PROBE 1 PULSE/REVOLUTION SLOT ENGINE SPEED MEASUREMENT SPEED PROBES Engine Indicating Speed Indicating System Speed Probes - Location The N speed probes and the trim balance probe are located, as shown below, in the front bearing compartment. No 3 STRUT No 2 BEARING SUPPORT INNER STRUT FAN SPEED PROBE RWARD LP STUB SHAFT TRIM BALANCE PROBE LP SHAFT PHONIC WHEEL ENGINE SPEED MEASUREMENT SPEED PROBES LOCATION 1 Terminal block 2 NF tube 3 Electrical lead 4 Fan speed probe 5 Trim balance probe Engine Indicating Exhaust Gas Temperature (E.G.T.) Indicating System The EGT is measured by 4 thermocouples which are located in the turbine exhaust case support struts (engine station 4.9). The 4 thermocouples are connected to the junction box by a thermocouple harness. The materials used for the thermocouples and harness are Chrome (CR) and Alumel (AL). An extension harness connects the EGT junction box to channels A and B of the EEC. Indication The EGT indication appears on the upper ECAM display unit. The ECAM provides the EGT indication : • • in 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 EGT is greater than (TBD) deg C: • • • • the indication becomes amber the MASTER CAUT light comes on accompanied by the single chime The following message appears on the ECAM (TBD) When EGT is greater than (TBD) deg C : • • • • the indication becomes red the MASTER WARN light comes on accompanied by the repetitive chime the following message appears on the ECAM (TBD) If the EGT has exceeded TBD deg C : • the maximum value reached is memorised • a small red line remains positioned on the analog scale, at that value (max point). T/C 3 A CIRCUIT B CJRCUJT T/C 4 JUNCTION BOX T/C 2 CHROMEL B CIRCUIT CHROMEL A CIRCUIT ALUMEJ 8 CIRCUIT ALUMEI A CIRCUIT EGT MANAGEMENT - SCHEMATIC T/C 1 Engine Indicating EGT Indicating System Thermocouple locations The 4 thermocouples are located in four of the hollow LP turbine exhaust struts. PISTON RING SEALS ALUMEL STUDS VANE No IT (Similar - 3, 6 & 8) CHROMEL VIEW ON B TYPICAL AT 4 POSITIONS EGT MEASUREMENT - THERMOCOUPLE Engine Indicating EGT Indicating System EGT Junction Box The location of the EGT junction box is shown below. ENGINE FLANGE ACAC EXHAUST DUCT FIRE DETECTION SUPPORT BRACKET EGT JUNCTION BOX Indicating The P3/T3 Sensor The P3/T3 sensor is a dual purpose aerodynamically shaped probe. It measures the pressure and temperature of the air stream at the inlet of the diffuser case. The data which results is transmitted to the EEC for control purposes. BOLT P3/T3 RIGHT LOWER FRONT-COWLS OPEN BOSS DIFFUSER CASE ASSEMBLY P3/T3 PROBE Engine Indicating E.G.T. and P3/T3 Measurement Electrical Harness The EGT and T3 measurement harness electrical connections are shown below. TEMPERATURE MEASUREMENT HARNESS ENGINE Indicating Engine Pressure Ratio (EPR) Indicating System The Engine Pressure Ratio (EPR) is used to set and control the engine thrust EPR = P4.9 : P2 P2 is measured by the P2/T2 Probe located in the inlet cowl. The associated indications are:• EPR max : thick amber line • EPR limit : max EPR value corresponding to thrust limit mode, which can be: o o o o P4.9 is measured by a pressure rake located at the LP Turbine exhaust. The pressures from these sensors are routed to the EEC. The EEC processes the pressure signals to totm actual EPR and transmits the EPR value to the ECAM for display on the upper screen. Each of the two EEC channels perform this operation independently. EPR: normal. EPR: if exceeds limit. EPR: if exceeds Red line limit (TBD) with master caution and single chime. G.A Go Around mode FLX Flexible Take Off mode MCT Maximum Continuous Thrust mode CL Climb mode • Flex T/O Temperature Assumed temperature entered by crew through the FMS - MCDU • EPR Reference Predicted EPR value according to TRA. EPR INDICATIONS Engine Indicating P2/T2 Sensor The P2/T2 sensor is a dual purpose probe which measures the total air temperature and pressure in the inlet air stream. The temperature and pressure signals are fed to the E.E.C The sensor is installed at the 12 o'clock position in the air inlet cowl. The temperature is measured by two platinum resistance elements, each channel of the E.E.C. monitors one of the elements. The pressure signal is fed to a pressure transducer in the E.E.C. The sensor is electrically heated to provide anti-ice protection. The E.E.C. software corrects any temperature signal errors caused by heating. PROBE ELECTRICAL CONNECTIONS REAR BULKHEAD P2/T2 MOUNTiNG BOLTS MOUNTING BRACKET SEALING BLOCK PROBE PRESSURE CONNECTION P2/T2 PROBE ACOUSTIC LINER P2/TZ PROBE - INSTALLATION Engine Indicating P2/T2 Sensor - Pneumatic Line The single pressure signal from the P2/T2 probe is routed to a pressure transducer located within the E.E.C. The pressure transducer has two outputs, one to channel A, one to channel B. Routing of the pressure signal - probe to E.E.C. - is shown below. PY LON INLET COWL REAR BULKHEAD ELECTRONIC ENGINE CONTROL UNiT RELAY BOX P2/T2 PROBE PNEUMATIC LINE - ROUTING Engine Indicating P2/T2 Probe - Electrical Harness The electrical harness connections for the P2/T2 probe are as shown below. Note the probe anti-icing heater utilises 115V AC from the aircraft: electrical system. P2/T2 PROBE HARNESS Engine Indicating Engine Vibration Indicating System The system monitors engine vibration for engine 1 and engine 2. Monitoring is done by a vibration transducer on each engine fan case. This produces an electrical signal in proportion to the vibration detected and sends it to the Engine Vibration Monitoring Unit (EVMU). Two channels come from each engine. The EVMU provides signals of Vibration, N1, N2 for display on the Engine page of the ECAM. The vibration transducer is installed on the fan case at the top left side of the engine. It is attached with bolts and is installed on a mounting plate. Indications The engine vibration indications are displayed in green on the lower ECAM display unit on the engine and cruise pages. The ECAM display unit receives the information through the ARINC 429 data bus via the SDAC 1 and SDAC 2. If the advisory level is reached, the indication flashes (0.6 sec bright, 0.3 sec normal). If the indication is not available, the corresponding indication is replaced by 2 amber crosses. VIBRATION ACCELEROMETER - REMOVAL/INSTALLATION PART ONE - SECTION 12 STARTING AND IGNITION SYSEM Engine Starting and Ignition System General The system comprises:• • • • • a pneumatic starter motor a starter air control valve a dual ignition system pneumatic ducting start control panels on the flight deck Starting and cranking cycles are initiated through switches/controls on the engine start panel and provide, on selection:• • • • • automatic starting manual starting dry cranking wet cranking continuous ignition Pneumatic supplies for the starter motor can be provided by:• • • the A.P.U. ground supply the other engine (if already started) During auto start attempts the critical parameters are monitored by the FADEC and in the event of a faulty start i.e. • • • • hot start hung start no light-up HP fuel valve failure the FADEC performs an automatic shut down (start abort) and provides a motoring cycle to clear fuel vapours and cool the engine. SIMPLIFIED STARTING SYSTEM Engine Starting and Ignition System Starter Air Duct Air supplies for the pneumatic starter motor may be supplied from:• • • the aircraft APU. the other engine - if already running a ground starter trolley Note: minimum duct pressure for starting should be approx 25 psi. All ducting in the system is designed for high pressure and high temperature operation and gimbal joints are incorporated to permit working movement. Air leakage is prevented by E-type seals interposed between all mating flanges, mating flanges are secured by Vee-band coupling clamps. STARTER STARTER SHUT OFF VALVE STARTRT DUCT- INSTALLATION Engine Starting and Ignition System Starter Motor The pneumatic starter motor is mounted on the forward face of the external gearbox and provides the drive to rotate the HP compressor to a speed at which light up can occur. The starter motor is connected by ducting to the aircraft pneumatic system. The starter motor ge.as and bearings arc lubricated by an integral lubrication system. Servicing features include:• an oil filler-level plug • drain plug with a built in magnetic chip detector The starter motor is attached to the external gearbox by a quick attach/disconnect adaptor (QAP). A quick detach Vee clamp connects the tarter motor to the adaptor. (FGEARBOX) STARTER INSTALLATION COUP I ING STARTER ADAPTER ENGINE STARTER ENGINE STARTER-INSTALLATION Engine Starting and Ignition System Starter Motor - Operation The starter is a pneumatically driven turbine unit that accelerates the H.P. rotor to the required speed for engine starting. The unit is mounted on the front face of the external gearbox. The starter comprises a single stage turbine, a reduction gear train, a clutch and an output drive shaft - all housed within a case incorporating an air inlet and exhaust. Compressed air enters the starter, impinges on the turbine blades to rotate the turbine, and leaves through the air exhaust. The reduction gear train converts the high speed, low torque rotation of the turbine to low speed, high torque rotation of the gear train hub. The ratchet teeth of the gear hub engage the pawls of the output drive shaft to transmit drive to the external gearbox, which in turn accelerates the engine H.P. compressor rotor assembly. When the air supply to the starter is cut off, the pawls overrun the gear train hub ratchet teeth allowing the turbine to coast to a stop while the engine HP turbine compressor assembly and, therefore, the external gearbox and starter output drive shaft continue to rotate. When the starter output drive shaft rotational speed increases above a predetermined rpm centrifugal force overcomes the tension of the clutch leaf springs, allowing the pawls to be pulled clear of the gear hub ratchet teeth to disengage the output drive shaft from the turbine. TURBINE ROTOR REDUCTION GEARS DRIVE SHAFT AIR INLET DISENGAGING MECHANISM TURBINE ROTOR GEAR TRAIN TURBINE NOZZLES CLUTCH AIR INLET TWO PIECE OUTPUT SHAF AIR EXHAUST (Elongated for clarity) STARTER - SCHEMATIC ROTATING ANNULUS Engine Starting and Ignition System Starter Air Control Valve The starter air control valve is a pneumatically operated, electrically controlled shut-off valve positioned on the lower right hand side of the L.P. compressor (fan) case. The start valve controls the air flow from the starter air duct to the starter motor. The start valve basically comprises a butterfly type vaive housed in a cylindrical valve body with in-line flanged end connectors, an actuator, a solenoid valve and a pressure controller. Manual Operation The starter air valve can be opened/closed manually using a 0.375 in square drive. Access is through a panel in the R.H. fan cowl. A valve position indicator is provided on the valve body. A micro switch provides valve position feed back information to the FADEC. STARTER SHUT-OFF VALVE - INSTALLATION Starting and Ignition System Starter Air Valve - Operation Valve Opening Air from immediately upstream ot the butterfly valve is filtered and routed through an orifice in the solenoid valve. Air upstream of the orifice is also admitted to the smaller piston of the double acting actuator. When the solenoid is energised the ball valve opens to admit air to the larger piston whilst simultaneously closing the vent port. The air acting on the larger piston oveicomes the combined force of upstream air pressure acting on the smaller piston and the actuator spring. Movement of the actuator is translated through the linkage to rotate the butterfly valve towards the open position. Valve Closing When the solenoid is de-energised, at approximately 6000 rpm (43%) N2, the ball valve closes and air acting on the larger piston is vented to atmosphere through the vent- Air pressure and actuator spring pressure acting on the smaller piston then closes the butterfly valve. Any loss of air pressure will cause the butterfly valve to close under the action of the actuator spring. STARTER AIR VALVE -SCHEMATIC Starting and Ignition System Starter Air Valve - Electrical Harness The starter air control valve electrical harness connections are as shown below. Starting and Ignition System Ignition Two independent ignition systems are provided. Continuous ignition may also be selected manually. The system comprises:- The ignition exciters provide approx 22.26 KV and the igniter discharge rate is 1.5/2.5 per second. • • • two ignition exciter units' - located as shown below two igniter plugs - located in the combustion system adjacent to No's 7 & 8 fuel spray nozzles two air cooled High Tension ignition connector leads Dual ignition is automatically selected for: • • • all in flight starts manual start attempts continuous ignition Single alternate ignition is automatically selected for ground auto starts. Continuous ignition is automatically selected when the engine anti-ice system is 'ON' or when the aircraft flaps are extended for:• • • take-off approach landing Test Operation of the ignition system can be checked on the ground, wj th the engine shut down, through the maintenance menu mode of the CFDS. CLAMP INPUT LEAD IGNITER PLUG COOLING AIR INLET COOLING AIR EXHAUST HIGH TENSION LEAD (inside cooling jacket) IGNITION EXCITER IGNITION SYSTEM Starting and Ignit ion System Ignition - Relay Box The ignition system utilises 115VAC supplied from the AC 115V Normal and standby bus bars to the relay box. The 115V relays - which are used to connect / isolate the supplies - are located in the relay box and are controlled by signals from the E.E.C. Note the same relay box also houses the relays which control the 115V AC supplies for P2/T2 probe heating. RELAY BOX Starting and Ignition System Electrical Harness The ignition system electrical harness connections are shown below. Starting and Ignition System Engine Starting Controls and Indication Two Master switches and a 3-position rotary control switch are mounted on the pedestal for auto mode engine start and pushbutton switches are mounted on the overhead panel for alternate mode (manual) start. The rotary switch is used in conjunction with either Master switch during auto mode starts or with MAN START push button switches in the alternate mode. IGN position • • used to perform starting in the auto and alternate modes. to call ENG START page on the lower ECAM display unit. NORM position • • rest position after engine starting. clears ENG START page CRANK position • used to pet form dry or wet motoring. Indication is displayed by lower and upper ECAM display units and by warning lights. The ENG MAN START button incorporates a blue ON legend and is normally in the released position with the ON legend off. Pressing the switch opens the pneumatic starter valve and illuminates the ON legend. An amber "fault" warning light on the pedestal mounted control panel illuminates when a disagreement occurs between the pneumatic starter valve position and that commanded by the starting sequence in the "auto" mode. ECAM lowet and upper screens display various warning and caution messages as well as the ENGINE start page. STARTING - CONTROLS AND INDICATIONS Starting and Ignition System Electrical Control Engine Interface Unit (EIU) Receives discrete electrical signals from the cockpit. Digitizes these signals and transmits them to the EEC. Also sends discrete signals to close air conditioning pack flow valves and to speed up the Auxiliary Power Unit (APU) if required. Electronic Engine Control (E.E.C.) The EEC generates pneumatic starter valve opening / closing signal in respect of control switch selection (rotary selector, master lever, MAN START push button switch) and N2 speed signal. Generates warning and caution messages for display on the ECAM. Starting and Ignition System Start Procedures (A) Pre-start 2. Manual Start Procedure • thrust lever : IDLE • master switch : OFF • mode selector : AUTO IGN • manual start button : OFF • aircraft booster pumps : ON (B) Start The engine may be started using the: (1) Auto Start or (2) Manual Start procedure 1. Auto Start Procedure • mode selector • master switch • on completion of start, : IGN/START : ON : return mode selector to NORM • mode selector • manual start button • when N2 reaches max motoring speed (min 15%) • on completion of start : IGN : ON : master switch ON : mode selector to AUTO IGN manual start push button OFF Starting and Ignition System Start Procedures Auto Start - Automatic actions • • • at approx 15% N2: ignition is activated at approx 19% N2: fuel PRSOV opens at approx 42% N2: starter valve closes ignition off Manual Start • • • • Starter air valve opens when manual start push button is pressed ON Fuel PRSOV opens and ignition is switched on when master switch is moved to ON position Starter air valve closes and ignition is cancelled automatically at approx 42% N2 Manual start push button must be selected OFF on completion of start General Auto starts or manual starts can be interrupted at any time by moving the Master switch to the OFF position. Starting and Ignition System Wet and Dry Cracking (Motoring) A dry motoring cycle will be required to: • • ventilate (blow through) the engine carry out leak checks. Method • • During this operation the starter motor is engaged but the fuel PRSOV remains closed and both ignition systems are OFF. Method • • Place ignition mode selector to CKANK on overhead panel, push MAN START push button for appropriate engine o o o ON legend illuminates starter air valve opens engine accelerates to max motoring speed Wet Motoring During wet motoring cycles the starter motor is engaged, fuel is ON, both ignition systems are OFF. Place ignition mode selector to CRANK On overhead panel, push MAN START push button for appropriate engine ON legend illuminates starter air valve opens o engine accelerates to max motoring speed Place master control switch to ON o o • o fuel PRSOV opens V2500 COURSE NOTES PART TWO - NACELLE V2500 COURSE NOTES PART TWO - NACELLE CONTENTS PART TWO - SECTION 1 INTRODUCTION PART TWO - SECTION 2 NACELLE – MECHANICAL ARRANGEMENT PART TWO - SECTION 3 THRUST REVERSEH PART TWO - SECTION 4 NACELLE VENTILATION AND FIRE PROTECTION PART TWO - SECTION 5 ENGINE REMOVAL / INSTALLATION PART TWO - SECTION 1 INTRODUCTION Introduction - Nacelle The propulsion unit comprises the Engine and the Nacelle. The major components which comprise the Nacelle are:• the air inlet cowl • the fan cowls (left and right hand) • the C - ducts (left and right hand) which incorporate the hydraulically operated thrust reverser unit • the combined nozzle assembly Components Weights Nose Cowl : 238 lbs (107.98 kg) Fan Cowl R.H. : 86 lbs (39.01 kg) Fan Cowl L.H. : 79 lbs (35.84 kg) C Ducts : 578 lhs (each) (262.25 kg) FAN AiR VALVE DISCHARGE P2/T2 PROBE C DUCT TRANSLATING COWL CARBON FIBRE ANT-ICING D CHAMBER ALUMINIUM FAN COWL DOORS CARBON-FIBRE INLET COWL CARBON FIBRE V2500 NACELLE COMMON NOZZLE ASSEMBLY (CNA) TITANIUM Introduction - Nacelle Access to Engine Mounted Units Access to units mounted on the LP compressor (fan) case and external gearbox is gained by opening the hinged fan cowls. Access to the core engine, and the units mounted on it, is gained by opening the hinged 'C ducts. RIGHT HAND C-DUCT THRUST REVERSER PYLON RIGHT HAND FAN COWL DOOR COMMON NOZZLE ASSEMBLY (CNA) HOLD OPEN ROD LEFT HAND C-DUCT THRUST REVERSER" INLET COWL HOLD OPEN ROD NACELLE SYSTEM-DETAILS LEFT HAND FAN COWL DOOR Introduction - Nacelle Access Panels and Doors Access panels and servicing doors are provided as shown below. ACTUATOR ACCESS (TYPICAL 4 PLACES) P2/T2 PHOBE ACCESS EEC GOOLING AIR EXHAUST EEC GOOLING AIR INLET INIfcHPHONI. JACK ANI-ICE DISCHARGE GRILLE AIR COOLED OIL COOLER GEARBOX OVERBOARO DISCHARGE MASTER CHIP DETECTOR STOW LOCKOUT (TYPICAL 2 PLACES) STARTER SHUTOFF VALVE/PRESSURE RELIEF DOOR OIL TANK SERVICE DOOR VENTILATION EXIT GRILLE DRAIN MAST NACELLE ACCESS PANELS AND DOORS STOW LOCKOUT PIN STOWAGE (TYPICAL PART TWO - SECTION 2 NACELLB MECHANICAL ARRANGEMENT NACELLE Mechanical Arrangement The propulsion unit comprises the Engine and the Nacelle. The major assemblies which together form the propulsion unit are shown below. 1. Inlet Cowl Bolted to the LP Compressor (Fan) case. 2,3. Fan Cowl Doors Secured to the strut at 4 hinge points on each side. 4,5. C-Ducts/Thrust Reverser Secured to the strut at 4 hinged points on each side. The C - ducts incorporate the hydraulically operated, cold stream, thrust reverser. 6. Common Nozzle Assembly (C.N-A.) Bolted to the rear flange of the turbine exhaust case 7,8. Engine Mounts Front mount • • attaches to the engine intermediate case with two brackets and a monoball mount. It transfers vertical, lateral and thrust loads. Rear mount • • attaches to the LP turbine exhaust case. It transfers vertical, lateral and torque loads. Propulsion Unit General Arrangement NACELLE Mechanical Arrangement Inlet Cowl The inlet cowl is bolted to the front of the L.P. Compressor (Fan) case. Purpose To supply all the air required by the engine, with minimum pressure losses and with an even pressure face to the fan. To minimise nacelle drag. Construction Hollow, inner and outer skins supported by front (titanium) and rear (Graphite / Epoxy composite) bulkheads Inner and outer skins manufactured from composites. Leading edge - Aluminium. Ice Protection Integral thermal anti-icing utilizing stage 7 air off take. P2 / T2 Probe • Probe located at TDC, attached to inner skin • Senses total inlet pressure (P2), and total inlet temperature (T2). • Access panel provided for maintenance. Handling • Two hoisting points provided. • Inlet cowl weighs 238 lbs. Ventilation Intake Ram air inlet to provide ventilation of Zone 1. INLET COWL NACELLE Mechanical Arrangement Inlet Cowl Details • Door locators: Automatically align the fan cowl doors to ensure good sealing. • Strut brackets: Provide location for the L and R hand fan cowl door support struts (front struts only). • Alignment pins: Ensure correct location of the inlet cowl to the fan case when the inlet cowl is installed. OUTER BARREL INLET VENTILATION P2/T2 PROBE ACCESS PANEL DOOR LOCATOR DOOR LOCATORS ALIGNMENT PIN (4 PLACES) HOiST POINTS (4 PLACES) RING STRUTBRACKET (2 PLACES) INNER BARREL INTERPHONE JACK DOOR LOCATOR TAI DISCHARGE GRILLE AFT BULKHEAD INLET COWL ASSEMBLY - DETAILS NACELLE Mechanical Arrangement P2/T2 Probe Access An access panel is provided maintenance of the P2/T2 probe ACCESS PAEL PYLON CAP VENTILATION INTAKE DOOR ALIGNMENT FEATURE ZONE VENTILATION OUTLET P2/T2 PROBE ACCESS PROBE ELECTRICAL CONNECTIONS DOOR ALIGNMENT FEATURE PROBE PRESSURE CONNECTION P2/T2 PROBE ACCESS PANEL NACELLE Mechanical Arrangement Inlet Cowl Removal/Replacement Note: Refer to the AMM Task 71-11-11-000-010. The procedure to remove/replace the Inlet Cowl is summarised below. • Open the L and R fan cowl doors • Attach the sling to the inlet cowl and the hoist. • Remove the coupling at the anti-ice duct joint and discard the seal. Fit now seal on installation. Disconnect the four electrical connectors at the top RH side of the cowl aft bulkhead. Disconnect the P2 signal pipe. Take the weight of the cowl on the sling with the hoist. Remove the cowl securing bolts. Move cowl forward carefully and lower onto dolly. • • • • • Replacement This is a reversal of the removal procedure. When offering up the inlet cowl use the 4 location spigots to ensure correct alignment. Tests Required • • • Test 30-21-00-710-010 - TAI system functional Test 71-00-00-790-010 - Leak test TAI system Test 73-22-00-710-010 - FADEC Operational test 1 3 2 1 SLING 2 PLASTIC SHIM 3 INLET COWL INLET COWL - SLIN6 NACELLE Mechanical Arrangement Fan Cowl Doors The doors extend rearwards from the inlet cowl to the leading edge of the 'C ducts. Purpose To provide access to the fan case and gearbox mounted accessories. Construction Graphite skins enclosing, aluminium honeycomb. Aluminium reinforcement at each corner minimises handling/impact damage and wear. Door Fitting 4 hinges on each door locate to the bottom of the pylon. Doors abut along bottom centre line and are secured to each other by 4 quick release - adjustable - latches. Hold-open Struts Doors can be propped open, for maintenance operations, using 2 swivel, telescopic struts which are stowed inside the doors when not in use. Warnings The fan cowl hold open struts must be in the extended position and both struts must always be used to hold the doors open. • Be careful when opening the doors in winds of more than 26 knots (30 mph) . • The fan cowl doors must not be opened in winds of more than 52 knots (60 mph). • HINGES DOOR ALIGNMENT FEATURES ECC COOLING INTAKE HOISTING POINTS ACOC COOLING OUTLET QUICK ACCESS MASTER CHIP DETECTOR QUICK ACCESS AIR STARTER VALVE AND BLOW OUT DOOR WHEELCASE BREATHER OUTLET QUICK ACCESS OIL FILL AND SIGHT GLASS BLOW OUT DOOR STRUTS LATCHES FAN COWL DOORS ADJUSTABLE KEEPERS NACELLE Mechanical Arrangement Fan Cowl Doors - Removal/Replacement Refer to the AMM • • Task 71-ll3-ll-000- 010 (Removal) Task 71-13-11-400-010 (Replacement) The procedure is summarised below. 1. Remove the blanking caps from the cowl slinging points. 2. Attach sling to cowl door and heist. 3. Open cowl door to gain access to hinges. 4. Remove split pins from hinge bolts. 5. Remove nuts and shouldered bolts. 6. Remove cowl door and lower onto dolly. Replacement This is the reversal of the removal sequence. On completion, check the cowl door alignment and latch tension. PYLON FAN COWL DOOR TYPICAL 4 PLACES 1AE-1N20404 BLANK CAP FAN COWL DOOR-REMOVAL SLING Mechanical Arrangement Fan Cowl Doors Latch Adjustment and Alignment The mismatch between the two cowl doors can be adjusted by fitting/removing shims as shown below. Latch tension is adjusted by use of the adjusting nut at the back of the latch keeper as shown below. The latch closing load should be between 45-60 lbs f / in. SHIM LEFT HAND FAISI COWL DOOR SHIM RIGHT HAND FAN COWL DOOR ALLOWABLE MISMATCH VIEW LOOKING FWD SHIM HEXAGONAL WRENCH SHIM KEEPER ASSEMBLY FAN COWL DOORS - LATCH ADJUSTMENT ADJUSTING NUT NACELLE Mechanical Arrangement Inlet Cowl - Handling/Transportation The arrangement for storage/transportation of the inlet cowl is shown below. IAE-1N20401 FAN COWL -TRANSPORT AND WORKSTAND NACELLE Mechanical Arrangement Fan Cowls - Hold Open Struts When in the open position the fan cowls are supported by two telescopic hold-open struts, using anchorage points provided on the fan case (rear) and inlet cowl (front). Stowage brackets are provided to securely locate the struts when they are not in use. The arrangement for stowing and anchoring the hold open struts are shown below. ANCHORAGE BRACKET (TYPICAL 2 PLACES) STOW BRACKET STOW BRACKET ATTACH POINT BRACKET ATTACH POINT BRACKET LATCH HANDLE {TYPICAL 4 PLACES) RH FAN COWL DO IR-HOLD-OPEN ROD BRACKETS NACELLE Mechanical Arrangement Fan Cowl Doors Detail The L.H. fan cowl door detail is shown below. HINGES (4 PLACES) DOOR ALIGNMENT FEATURES HOISTING FEATURE (TYPICAL 2 PLACES) DOOR STRUT ATTACHMENT DOOR STRUT ATTACHMENT OIL FILLER ACCESS DOOR MASTER CHIP DETECTOR ACCESS DOOR STRUT STOWAGE STRUT STOWAGE VENTILATION GRILLE ADJUSTABLE KEEPERS LH FAN COWL DOOR-HOLD-DETAILS DOOR ALIGNMENT FEATURE HINGES (4 PLACES) DOOR ALIGNMENT FEATURES EEC COOLING AIR DISCHARGE EEC COOLING INTAKE HOISTING FEATURE DOOR STRUT ATTACHMENT STARTER VALVE ACCESS/PRESSURE RELIEF DOOR ACOC AIR GEARBOX OVERBOARD BREATHER OUTLET DOOR ALIGNMENT FEATURE DOOR STRUT ATTACHMENT STRUT STOWAGE VENTILATION GRILLE RH FAN COWL DOOR-HOLD-DETAILS LATCHES (4 PLACES) NACELLE Nechanical Arrangement Fan Cowl Doors - Storage/Transportation The arrangement for storing / transporting the cowl doors is shown below. IAE-1N20402 FAN COWL DOOR-HANDLING DOLLY NACELLE Mechanical Arrangement C – Ducts / Thrust Revecser The C - ducts extend rearwards from the fan cowls to the Combined Nozzle Assembly (CAN). An overview of the C – ducts / reverser assembly is shown below. Purpose The C - ducts: • form the cowling around the core engine (inner barrel) • form the fan air duct between the fan case exit and the entrance to the C.N.A. • house the thrust reverser operating mechanism and cascades • form the outer cowling between the fan cowls and C.N.A. Construction Mostly composites but some sections are metallic - mainly aluminium - e.g. inner barrel, Mocker doors and links. C - Duct fitting Each C-duct is located to bottom of the strut by 4 hinge brackets. The 'C ducts abut along the bottom cent-re line and are secured to each other by a series of latches which are located under a hinged access panel. Opening/Closing C-Ducts Access to the core engine, for maintenance operation is gained by opening the hinged C ducts. The 'C ducts are supported/locked in the open position by integral support struts. Opening is effected through an integral, self contained, hand pump operated, hydraulic system. UPPER TRACK AND HINGE FITTING PRECOOLER DUCT DOOR OPENING ACTUATOR COOLING DUCT TRANSLATING SLEEVE UPPER ACTUATOR ACTUATOR ACCESS RIGHT CDUCT INNER BARREL STOW LOCKOUT (TYPICAL 2 PLACES) HOLD OPEN ROD STOW LOCKOUT/PIN STOWAGE (TYPICAL 2 PLACES) MANUAL C-DUCT HYDRAULIC LINE C-DUCT - OVERALL VIEW NACELLE Mechanical Arrangement LH Reverser ( C - Duct) An overview of the ' C' duct is shown below. Latch Assemblies Hold - open struts Four latch hooks engage with four latch keepers, on the RH C - duct to secure the two C - ducts together. A further double latch hook, at the rear, is used to lock the two translating sleeves together. Two struts on each C - duct support the ducts in the open position to facilitate maintenance operations on the core engine. Bumpers 5 lower bumpers and 4 upper bumpers absorb the fan air compressive loads. The bumpers, incorporate adjusters shim packs - to provide rigging adjustments. Heatshield The whole of the inner barrel is lined with heat reflective / insulating material. Pre-Coolor Ducts Supply fan air to the pre-cooler of the ECS located in the aircraft strut. Note For added safety during maintenance operations and to support the C - ducts during an engine change a GSE safety rod is inserted. Take Up Device This device is used to pull the two 'C ducts together to facilitate engagement of the main latch assemblies. UPPER BUMPER (4 PLACES) HEAT SHIELD UPPER BIFURCATION DOOR OPENING ACTUATOR COOLING DUCT PRECOOLER DUCT TORQUE RING LOAD SHARE FITTING 7TH STAGE BLEED C-DUCT OPENING ACTUATOR OUTER V BLADE ONE PIECE INNER BARREL/ BIFURCATION AFT HOLD OPEN ROD 7TH STAGE BLEED HEAT SHIELD LOWER BUMPER (5 PLACES) LOWER BIFURCATION ACAC INLET FORWARD HOLD OPEN ROD LH THRUST REVERSER ASSEMBLY FORWARD BUMPER AND LATCH NACELLE Mechanical Arrangement RH reverser (C – Duct) An overview of the RH C - duct is shown below. Description is same as LH C - duct . 7th & 10th Stage Bleed Ports Ports in both 'C ducts discharge air from the 7 & 10 stage compressor handling bleed valves, when open, into the fan duct. DOOR OPENING ACTUATOR COOLING DUCT TORQUE RING. UPPER BIFURCATION HEATSHiELD PRECOOLER DUCT UPPER BUMPER (4 PLACES) 7TH STAGE BLEED LOAD SHARE FITTING 10TH STAGE BLEED C-DUCT OPENING ACTUATOR 7THSTAGE BLEED AFT LOWER BUMPER AFT HOLD OPEN ROD AOAC BLEED ONE PIECE INNER BARREL/ BIFURCATION OUTER 'V BLADE HEATSHIELD LOWER BUMPER (5 PLACES) LOWER BIFURCATION FORWARD HOLD OPEN ROD RH THRUST REVERSER ASSEMBLY NACELLE Mechanical Arrangement C - Duct Latches A total of six latches arc used to secure the two C - ducts to each other. Access to latch A is through the L. and R. hand fan cowls. 3 latches B are located under a hinged access panel. Latch C is a double latch assembly but the two latches must be released / latched individually. LATCH DETAIL A LATCH DETAIL B LATCH DETAIL C C DUCTS - LATCHES Mechanical Arrangement Latch Access Panel and Take Up Device An access panel is provided to gain access to three of the C-duct latches and the C-duct take-up device. The take-up device is a turnbuckle arrangement which is used to draw the two C-ducts together. This is necessary to compress the C-duct seals far enough to enable the latch hooks to engage with the latch keepers. The take up device is used when closing and opening the C-ducts. The take up device must be disengaged and returned to its stowage bracket, inside the LH C-duct, when not in use. CNA TRANSLATING SLEEVE DOUBLE LATCH THRUST REVERSER C DUCT LATCHES C DUCT TAKE UP DEVICE DOOR LATCH LATCH ACCESS PANEL BRACKET ACCESS DOOR INSTALLATION HOOK NACELLE Mechanical Arrangement C - Duct Hold Open Struts Two hold open struts are provided on each C - duct to support the C-ducts in the open position. The struts engage with anchorage points located on the engine as shown below. When not in use the struts are located in stowage brackets provided inside the C - duct. The front strut is a fixed length strut. The rear strut is a telescopic strut and must be extended before use. The arrangement for the LH C - duct is shown below, the RH C-duct is similar. Warning Both struts must always be used to support the 'C ducts in the open position. The C-ducts weigh approx. 578 lbs each. Serious injury to personnel working under the C - ducts can occur if the C-duct is suddenly released. ROD END FITTING C DUCT ROD ANCHORAGE BRACKET ROD ANCHORAGE BRACKET LH T/R DOOR HOLD-OPEN RODS-INSTALLED Mechanical Arrangement C-Ducts - Maintenance Each C-duct is attached to the aircraft pylon by four hinges. The three front attachment points ate provided by beams located on the bottom of the pylon. The beams are not rigidly attached to the pylon and this provides a degree of self alignment when closing the 'C ducts. The rear hinge point is a solid location on the side of the pylon. SLEEVE BOLT WASHER WASHER BOLT NUT WASHER NUT NUT WASHER (TYPICAL LHAND RH INSTALLATION) PYLON GSE PIN (AIRBUS FURNISMFH) BOLT BOLT THRUST REVERSER C-DUCT Mechanical Arrangement C-Duct Opening/Closing System Purpose To provide a mechanical method of to raising /l owering the C-ducts facilitate one man operation. Features On each 'C duct • • • • single acting hydraulic actuator self sealing/quick release hydraulic connection rigid and flexible hydraulic hoses pylon and C duct hydraulic actuator attachment brackets Aircraft carried • hand operated hydraulic pump Note The hydraulic fluid used in the system is Engine Lubricating Oil. CAUTION • WING SLATS MUST BE RETRACTED AND DEACTIVATED • ALL 6 LATCHES AND TAKE UP DEVICES MUST BE RELEASED • IF REVERSER IS OE PLOYED. PYLON FAIRING MUST BE REMOVED CAUTION SEE DECAL ABOVE BEFORE OPENING C DUCT 'C DUCT OPEN POSITION C DUCT CLOSED POSITION DOOR OPENING ACTUATORS MANIFOLD/PRESS RELIEF VALVE IAE-1N20009 HAND PUMP FLEX HOSE QUICK DISCONNECT FITTING THRUST REVERSER D00R OPENING SYSTEM - OVERVIEW NACELLE Mechanical Arrangement C-Duct Opening Actuator The installation of the RH C-duct actuator is shown below. LH actuator is similar. The actuator hydraulic supply tube is continuously cooled by air taken from the fan duct. The arrangement of the cooling jacket is shown below. The actuator has an integral, one way, check valve. This restricts the fluid return when the C-duct is closing and thus controls the speed at which the C-duct closes. Note The C-duct closes owing to its own weight, thus it is not necessary to use the hand pump to pressurise the actuator during the closing operations. HINGE BEAM PYLON CDUCT OPENING ACT UAT OR A C T UA T O R CHECK VALVE CLAMP HYDRAULIC TUBE HOSE COVER (EXPANDED) TUBING AND HOSES RELIEF VALVE WITH I N T E G R A L FILTER FILTER VENT MANIFOLD QUICK DISCONNECT FITTING RH T/R - DOOR OPENING ACTUATOR COOLIN6 JACKETS - INSTALLATION NACELLE Mechanical Arrangement Combined Nozzle Assembly (CAN) The CNA is bolted to the rear flange of the turbine exhaust casing. Purpose • forms the exhaust unit • mixes the hot and cold gas streams and ejects the combined flow to atmosphere through a single propelling nozzle. • completes the engine nacelle EXIT NOZZLE INCONEL 625 BRAZED SANDWICH FAIRING/UPPER STRUT TITANIUM SKIN AND FRAME INNER ANNULUS LUMINUM SKIN TITANIUM FRAME TYPICAL LOWER STRUT TITANIUM SKIN COMMON NOZZLE ASSEMBLY - INSTALLATION NACELLE Mechanical Arrangement C-Duct – Maintenance Slinging & Hoisting The slinging and hoisting arrangements are shown below. HOIST IAE:1N20002 HINGE ACCESS PANEL SCREW (11 PLACES SLING - TRUST REVERSER HALF-REMOVAL/INSTLLATION NACELLE Mechanical Arrangement CNA Handling The arrangements for slinging, hoisting, stowage and transportation are shown below. PYLON IUPPER STRUT EXHAUST PLUG (.75 IMCH)(19.5mm) CLEARANCES TYPICAL COMMON NOZZLE EXHAUST COLLECTOR NUT (56 PLACES) IAE-1N20001 COMMON NOZZLE FIXTURE BOLT (56 PLACES) IAE-1N20004 CNA DOLLY COMMON NOZZLE ASSEMBLY-REMOVAL/INSTLLATION Mechanical Arrangement C-Ducts - Maintenance The arrangement for storage/transportation of the C-duct(s) is shown below. STRAP C DUCT DOLLY AND WORKSTAND IAE 1N20005 (LH) IAE 1N20006 (RH) FIGURE 5 DOLLY AND WORKSTAND Exhaust Plug The exhaust, plug is located inside the CNA as shown below. Removal of the exhaust plug provides access to the cover plate of the rear bearing compartment (No5 bearing). TYPICAL 13 PLACES EXHAUST PLUG - INSTALLATION NACELLE Mechanical Arrangement Engine Mounts The engine is mounted to the pylon at two places. Front Mount Locates to the engine intermediate casing at 3 points, 2 brackets and a Monoball mount. Located to pylon by 5 bolts aligned by 2 shear pins. Loads Transfers vertical, lateral and thrust loads. CROSS BEAM MOUNT BEAM ASSEMBY SHEAR PIN BEAM SHEAR PIN (2 places) MONOBALL STOP PLATE THRUST LINK SHEAR PIN BEAM JOINING BOLTS THRUST LINK ANTI-ROTATION AND RETAINING PLATES (10 places) FORWARD ENGINE MOUNT Mechanical Arrangement Engine Mounts Rear Mount Locates to the LP turbine exhaust casing. Transfers vertical, lateral and torque loads. Located to the pylon by 4 bolts aligned by 2 shear pins. BEAM ASSY JOINING BOLTS MOUNT BEAM ASSY SHEAR PIN (2 places) LINK ASSY ANTI-ROTATION AND RETAINING PLATES TURBINE EXIT CASE RETAINING PLATE SLEEVF LINK SHEAR PIN SOLID PIN AND SLEEVE AFT ENGINE MOUNT V2500 NACELLE PART TWO - SECTION 3 THRUST REVERSER NACELLE Thrust Reverser Unit (T.R.U.) Introduction Purpose The thrust reverser provides deceleration forces to slow the aircraft on landing or during an abandoned take off. It is incorporated in the 'C ducts and forms an integral part of the fan stream exhaust duct. The thrust reverser comprises a fixed inner and a movable outer (translating} assembly. Controls Selection of reverse thrust, and control of engine power in reverse thrust is effected by the gated throttle lever {thrust lever). All signals to and from the T.R.U. are through the E.E.C. and the Engine Interface Unit (E.I.U.). Features • electrical control circuit • hydraulic actuation system • positional information feedback system • actuator lock position sensors and feed back • automatic restow system • manual deployment / stow for maintenance • manual lock out allows aircraft dispatch with inoperative thrust reverser An overview of the thrust reverser is shown below. GRAPHITE COMPOSITE CASCADES UPPER TRACK AND HINGE FITTING PRECOOLER DUCT UPPER ACTUATOR ACTUATOR ACCESS TRANSLATING COWL ONE PIECE INNER BARREL STOW LOCKOUT BLOCKER DOORS PIN STOWAGE LOWER ACTUATOR REVERSER OVERVIEW Introduction - Operation A sectional view through the Nacelle showing the translating cowl in the stowed (forward thrust) position and the deployed (reverse thrust) position is shown below. GRAPHITE PANEL DYNAROHR ACOUSTIC PANEL BLOCKER DOOR STOWED - FORWARD THRUST ALUMINIUM OUTER PANEL CASCADE ACTUATOR TORQUE RING CASCADE AFT RING ALUMINIUM BLOCKER DOOR LINK ACCESS DOOR 21 in STROKE TRANSLATING SLEEVE INNER BARREL ENGINE VEE GROOVE DEPLOYED - REVERSE THRUST THRUST REVERSER OPERATION Thrust Reverser Introduction - Controls and Indications Selection of reverse thrust and control of engine power in reverse thrust is effected by the normal thrust lever (throttle lever). Movement of the thrust lever rearwards is limited to forward idle by the latches carried on the thrust levers. When the latches are lifted the thrust lever is allowed to move further rearwards (maximum movement approx. 12 degrees). This initiates the following sequence of events which are all controlled by the E.E.C. • the thrust reverser begins to deploy, engine power commanded to idle. • when the thrust reverser has deployed approx. 78%, EEC commands engine to accelerate to reverse power selected by thrust lever position. • thrust reverser continues to deploy to fully deployed position. Indications NO INDICATION • the thrust reverser is fully stowed • both locks are fully engaged. REV • both locks are disengaged and • the reverser is between fully stowed and fully deployed i.e. in transit. REV • thrust reverser is fully deployed. FORWARD REVERSE UPPER ECAM DISPLAY THRUST REVERSE LATCHING LEVERS THROTTLE CONTROL LEVERS REVERSE POWER RANGE ON QUADRANT ENGINE PANEL THRUST REVERSER CONTROL INDICATING NACELLE Thrust Reverser - Description Hydraulic Actuation System Purpose To provide the force required to move the translating cowl during thrust reverser operation. Features • Two linear hydraulic actuators per translating cowl. • One non-locking (upper) actuator which incorporates a Linear Voltage Differential Transformer (LVDT) to provide actuator positional feedback signals. • One locking (lower) actuator which includes a locking mechanism to hold the reverser in the stowed position, the locks incorporate sensors which signal lock position to the EEC. • The Hydraulic Control Unit (HCU) which incorporates the isolation valve, the directional control valve (DCV) and the pressure switch. • Flexible hydraulic hose assemblies which link the two upper actuators and incorporate the hydraulic feed tubes (hydraulic T unions). • Rigid hydraulic tubes which link the upper and lower actuators. • Actuator synchronisation system which utilizes flexible synchronising cables running inside the hydraulic deploy tubes. • Manual drive system (which utilises the synchronising cables) used to stow or deploy the translating cowl for maintenance operations. TRANSLATING SLEEVE UPPER ACTUATOR HYDRAULIC "T" AIRBUS FURNISHED HYDRAULIC CONTROL UNIT RETRACT HYDRAULIC HOSE ASSY - EXTEND HYDRAULIC HOSE/FLEXSHAFT ASSEMBLY - EXTEND TUBE LOWER LOCKING ACTUATOR RETRAC TUBE T/R HYDRAULIC SYSTEM ACTUATION NACELLE Thrust Reverser - Description Hydraulic Control Unit The Hydraulic Control Unit (HCU) is a self contained line replaceable unit (LRU), providing safe control of the thrust reverser actuators in response to electrical position demands from either channel of the EEC. The HCU comprises the following:• One Isolation valve, which can be mechanically latched in the closed position during maintenance. • One direction control valve to port fluid to the actuators in response to stow or deploy demands. • switch to detect system One pressure pressurisation downstream of the isolation valve. • Two dual-coil solenoid valves to control operation of the isolation and direction control valves in response to electrical signal from either channel of the EEC. • One filter with clogging indicator. • One bleed valve. Location The hydraulic control unit is bolted to the bottom of the engine pylon in the fan case area. Access is gained by opening the L.H. fan cowl. RETURN SUPPLY FILTER HYDRAULIC ISOLATION VALVE DEACTIVATING LEVER SOLENOIDS PRESSURE SWITCH HCU PYLON CLOGGING INDICATOR BYPASS VALVE DEPLOY LOCK PIN STORAGE BRACKET HYDRAULIC CONTROL UNIT DEACTIVATION (Lock out) PIN LOCKOUT LEVER POSITION HYDRAULIC FILTER ELECTRICAL CONNECTORS UNLOCK LEVER POSITION T/R HYDRAULIC CONTROL UNIT DIRECTION CONTROL VALVE NACELLE Thrust Reverser - Description Hydraulic Actuators Four actuators are used for each thrust reverser, two actuators are used for each translating cowl. • the lower actuators incorporate an integral lock mechanism which holds the piston in the fully stowed position. • the upper actuators incorporate an integral Linear Variable Directional Transformer (LVDT) to indicate piston position, and thus translating cowl position, to the EEC. All actuators use hydraulic snubbing at the end of the deploy stroke to slow down the actuators over the final part of the deploy stroke. All actuators also incorporate the necessary deploy stroke mechanical stops. Upper Actuators The two upper actuators are identical Line Replaceable Units (LRU's), and, in conjunction with the two lower locking actuators, control movement of the fan reverser translating elements in response to hydraulic inputs from the hydraulic control unit. EYE END MOUNTING GIMBAL UNION (Stow) ADAPTER (Deploy) UPPER NONLOCKING ACTUATOR BALL ADJUSTABLE EYE END FLOW CONTROL VALVE SEALS SPLIT GUIDE CLUTCH SHAFT PIVOT END CAP BEARING RINGS SEAL PIN SCRAPER ROD GUIDE THRUST NEEDLE BEARING NUT PISTON JACKHEAD ASSEMBLY SCREWSHAFT PISTON WORMWHEEL UNION (Stow) MOUNTING BUSH CLUTCH SHAFT GIMBAL JACKHEAD ASSEMBLY SCREWSHAFT WORMSHAFT LOCKNUT LVDT ADAPTER (Deploy) END CAP UPPER NONLOCKING ACTUATOR - DETAIL Thrust Reverser Lower Locking Actuators The two lower locking actuators are identical Line Replaceable Units (L.R.U's) and, in conjunction with the two upper actuators, control movement of the fan reverser translating elements in response to hydraulic inputs from the hydraulic control unit. The actuators incorporate an integral lock mechanism to hold the piston rod when the actuator is in the fully stowed position. The lock releases on rising hydraulic pressure when deploy is commanded via the hydraulic control unit. The lock mechanism incorporates a manual release facility and proximity switch for electrical lock position feedback to the EEC. Thrust Reverser - Description Lower Locking Actuators - Locking Mechanism In the stowed position (forward thrust) the stow and deploy ports are both connected to return pressure. The piston assembly is positively, mechanically, held in the stow position by the claws of the tine lock. On a deploy command from the E.E.C. the H.C.U. ports hydraulic fluid to both the stow and deploy ports. As the flow into the actuator rod side is not restricted, the pressure rise in the rod side of the piston is greater than that in the head side. The net effect is to initially move the piston in the stow direction, which relieves the locking load on the tine lock. As the pressure in the head side subsequently builds up, the differential area unlock sleeve moves forward against the lock spring preload to release the radial restriction to tine lock movement. Thus, the lock is released, the resultant pressure acting on both sides of the differential area piston drives the piston and synchronising nut through the tine lock in the deploy sense. The release sleeve spring holds the sleeve in its disengaged position. Thrust Reverser - Description Synchronisation System The system is used to provide inter-actuator synchronisation using flexible synchronisation cables inside the deploy hydraulic tubing. The system comprises:• One T-piece connector, to split the deploy hydraulic pipe connection from the hydraulic control unit to each of the upper feedback actuators. • Two flexible tubes to carry the deploy hydraulics from the T-piece to the upper actuators and guide the flexible synchronising shafting running between the two upper actuators. • Two rigid tubes to carry the deploy hydraulics between upper and lower actuators on each side and guide the flexible synchronising shafting running between the upper and lower actuators. • Three flexible shafts with square male end fittings to interconnect the synchronising mechanism of each actuator The two deploy tubes incorporate a telescopic coupling at one end to permit simple removal and replacement without disturbing actuator installation. T-PIECE HOUSING SLEEVE FLEXIBLE GUIDE TUBE O RING SEAL FLEXIBLE DRIVE SHAFT (TOP) FLEXIBLE DRIVE SHAFT (TOP) SHAFTING GUIDE T- PIECE HOUSING ASSEMBLY FLEXIBLE DRIVE SHAFT (SIDE) GUIDE TUBE FLEXIBLE DRIVE SHAFT (SIDE) GUIDE TUBE LOCKNUT R.NGSEAL END FITTING FIGURE 2A END FITTING SLEEVE ACTUATOR - FLEXSHAFT INSTALLATION Thrust Reverser Cascades (Deflector Boxes) In the reverse thrust mode all the fan air is directed through the cascades which eject the air in a forward direction. A total of 16 cascades are fitted as shown below. THURST REVERSER CASCADE Thrust Reverser - Operation The thrust reverser is shown schematically, below, in the forward thrust (flight) position. In this position:• hydraulic pressure from the aircraft system is available as far as the control unit only • the isolation valve is in the closed position - control solenoids de-energised • both sides of the actuator pistons are ported to hydraulic return • the thrust reverser is maintained in the forward thrust position by mechanical locks which are an integral part of the locking (lower) actuators • the directional control valve is in the stow position, control solenoids de-energised THRUST REVERSER TRU - SCHEMATIC Thrust Reverser - Operation Deploy/Stow Reverse Thrust Selection Selection of reverse thrust will - via the EEC - provides signals to: • open the isolation valve to allow hydraulic pressure to the thrust reverser • move the directional control valve (DCV) to the deploy position Hydraulic pressure is then felt on both sides of the actuator pistons but, because of the differential piston areas, the actuators will extend to move the translating cowls to the reverse thrust position. Note the signal from the EEC to the DCV is routed via a relay which is closed by the Engine Interface Unit (EIU) when reverse thrust is selected within the permitted operating envelope. Re-selection of Forward Thrust • moves DCV to stow position • the open signal to the isolation valve is maintained providing hydraulic pressure to the stow side of the actuator pistons • the extend side of the actuator pistons is ported, via the DCV to hydraulic return • the EEC will cancel the open signal to the isolation valve 5 seconds after the translating cowls reach the fully stowed position, to ensure full lock engagement. THRUST REVERSER OPERATION NACELLE Thrust Reverser - Maintenance Manual Deploy/Stow The thrust reverser may be deployed/stowed manually for maintenance - trouble shooting operations. The procedure is summarised below, the full procedure, warnings and cautions may be found in the M.M. CH 78. • • open and tag the following circuit breakers for the appropriate engine T.B.D. by Airbus open the L and R hand fan cowls • move the thrust reverser hydraulic control unit deactivation lever to the de-activated position and insert lockout pin • disengage the locks on the two locking (lower) actuators - insert pins to ensure locks remain disengaged - see below • position t lie non return valve in the hydraulic return line to the by-pass position, see next page (deploy only -not necessary for stow operation) • insert 3/8 inch square drive speed brace into external socket - see below - push to engage drive and rotate speed brace to extend /retract translating cowl as required. Note do not exceed max. indicated torque loading. THRUST REVERSER MANUAL DEPLOY NACELLE Thrust Reverser - Maintenance Manual Deploy During manual deploy operations it is necessary to draw some hydraulic fluid from the aircraft system. This is done by moving the non return by-pass lever, located in the hydraulic return line, to the by-pass position as shown below. Access to the non-return valve is gained by removing the pylon access panel. On completion of the manual deploy operation the by-pass valve must be re-positioned to the normal position and the access panel replaced. A baulking feature on the access panel prevents the panel being closed if the by-pass lever is in the by-pass position. TO HYDRAULIC RESERVOIR FROM HCU VIEW ON A T/R MANUAL DEPLOY NON RETURN VALVE (By-pass) Thrust Reverser - Maintenance De-activation An inoperative thrust reverser may be locked in the forward thrust position for flight, as permitted by M.E.L. requirements). Method The procedure is summarised below, the full procedure is described in the AMM CH 78 Task 78-32-00-040-011. • if the thrust reverser is deployed -stowed manually as previously described • install the lock out pin in the de-activation lever of the hydraulic control unit. Task 78-32-00-041-010. • remove the translating cowl de-activation pins (2) from their stowage and insert them in the deactivation position Note When fully inserted in the de-activation position the pins will protrude approx. 0.8" to provide visual indication of "lock out". THRUST REVERSER -DEACTIVATION Thrust Reverser - Maintenance Cascades – Removal & Installation • The cascades are not all interchangeable. • The cascade positions are identified by numbering. • No 1 position is the top position on the RH C-duct. • numbering is clock-wise when viewed from the rear. • The cascades (deflector boxes) are identified by part No, as shown below. VIEW LOOKING FORWARD LH NACELLE VIEW LOOKING FORWARD RH NACELLE # BLANK CASCADE THRUST REVERSER CASCADE POSITIONS Thrust Reverser - Maintenance Cascades - Removal/Installation The procedure to remove/install a cascade is outlined below. Reference must be made to the M.M. Task 78-32-19-000010. • deploy the thrust reverser, hydraulically or manually • de-activate the hydraulic control unit (HCU). • remove the cascade (deflector box) Note blank deflector boxes have 12 bolts, all others have eight bolts • installation of cascade boxes is a . reversal of the above procedure. LH AFT SUPPORT RING THESE TWO (2) HOLES ARE 0.10 INCH FURTHER AFT TO PROVIDE BULKING FEATURE THRUST REVERSER CASCADE - INSTALLATION Thrust Reverser - Electrical Harness The thrust reverses electrical harness connections are as shown below. HYDRAULIC CONTROL UNIT 1 Directional control valve 2 Isolation valve 3 Hydraulic pressure switch LH LVDT RH LVDT LH LOCK SENSOR RH LOCK SENSOR RH CDUCT 4005VC (9O0J) THRUST REVERSER - HARNESS V2500 GENERAL PART THREE - SECTION 2 COMPONENT LOCATION GUIDE HYDRAULIC PUMP NOSE CONE FAN CASE FUEL COOLED OIL COOLER HP COMPRESSOR SECTION REAR ENGINE MOUNT FUEL FILTER OIL TANK HYDRAULIC PUMP COMBUSTION SECTION OIL PUMP STAGE 10 BLEED VALVE COMMON NOZZLE FUEL PUMP GEARBOX ENGINE- LEFT HAND SIDE- COMPONENT LOCATION RELAY BOX STAGE 7 BLEED VALVES LP COMPRESSOR (FAN) EEC TURBINE SECTION No. 4 BEARING COMPARTMENT AIR COOLER STARTER DE-OILER AIR COOLED OIL COOLER IDG BLEED VALVE CONTROL VALVES ENGINE- RIGHT HAND SIDE- COMPONENT LOCATION V2500 GENERAL PART THREE - SECTION 1 TROUBLE SHOOTING TROUBLESHOOTING A320 Introduction Menu Mode The A3 20 utilises Electronic systems to detect, categorise and display faults which occur during aircraft operation. This mode is only available when the aircraft is on the ground and the engines are not running. One of these systems is the Centralised Fault Display: System (CFDS). The CFDS consists basically of;- It allows maintenance crews to establish an interactive dialogue through one MCDU with any of' the connected aircraft systems. • all (Bite) Built In Test Equipment portions of. the electronic control systems • the central computer - the Centralised Fault Display Interface Unit (CFDIU) • two Multi-Purpose Control and Display Units (MCDU) The CFDS has two operating modes:• normal mode (or reporting mode) • menu mode (or interactive mode) Normal Mode The CFDS continuously receives failure and status information from the Bite portions of all the connected systems. The location of the units is shown below. MULTI PUHPOSE PRINTER 2 MCDU TO MCDU's AND PRINTER CFDIU bite bite bite bite ELECTRONIC SYSTEM AVIONICS BAY - 80 VU A 320 - CFDS - COMPONENTS LOCATION A320 TROUBLESHOOTING Centralised Fault Display System (CFDS) The CFDS is used to provide maintenance information and can also be used to initiate various functional tests, e.g. thrust reverser, ignition system, etc. The diagram below shows how the CFDS interfaces with the Engine Electronic Control (EEC) and the Electronic Centralised Aircraft Monitoring System (ECAM). Engine Systems which are controlled by the EEC are continuously monitored by fault detection logic circuitry (BITE), within the EEC. When the EEC detects a fault it generates fault data which is then transmitted to three user systems:• the Flight Warning Computer, which generates ECAM displays to alert the flight crew and provide advice on recovery / handling procedures. • through the Central Maintenance System (CMS) to the CFDIU. • to its own non volatile memory Maintenance personnel can interrogate this memory through the CFDS to obtain more detailed fault data. Printed copies of all fault data can be obtained from the EEC memory, using the CFDS printer. ATA Ref.' Language Message MCDU Prinloul of In-Hight messages 761100 EGAM ENG W.D. TRA sense/HC/EEC Ptinlout of EEC fault memory CFDIU Maintenance crew can use CMS menu mode to obtain more Information Flight Warning Compuler ARINC output bus CMS logic Faull delectfon logic Memory EEC Faulted componanl E.G. TRA crosscheclt FAULT MESSAGE ANNUNCITION ECAM/CFDS A3 20 CFDS Failure Classification Three classes of failures have been defined, according to:• how critical the failure is • its operational impact • the scheduled maintenance policy The failures are categorised as:Class 1 These are failures which have consequence on the flight in progress. an operational Class 2 These are failures which do not have an operational consequence on the flight in progress but may affect subsequent flights - refer to Minimum Equipment List (MEL) for functions lost. Class 3 Failures which have no operational effect and can be left until next scheduled maintenance check. The failure classification is summarized below. Note: The EEC does not transmit class three messages. 1 2 Operational consequences on the current flight YES NO NO Indicated to the Pilots YES YES NO Failure Classes Dispatch consequences Maintenance information Warnings / Flags System pages REFER TO MMEL may be "GO" "GO IF" "NO GO" On the system Display "STATUS"page FUNCTIONS LOST INDICATED IN MMEL "GO" without conditions HAVE TO BE REPORTED BY THE PILOTS IN THE LOG BOOK INDICATED AT THE END OF EACH FLIGHT LEG : "LAST LEG REPORT FAILURE CLASSIFICATION 3 NO REFERENCE IN MMEL AVAILABLE ON REQUEST CAN BE LEFT UNCORRECTED UNTIL THE NEXT ADEQUATE MAINTENANCE OPPORTUNITY A320 TROUBLESHOOTING CFDS Normal Mode - Health Monitoring Wraparound Checks When operating in the Normal mode the Bite portion of the EEC continuously monitors the operation of the systems, under its control, by carrying out:- This checks out the system electrical circuitry, detects faults such as - loose connectors, chaffed harness, bent pins, high resistance, broken cable etc. 1. track checks 2. cross checks 3. wraparound checks Track Check The EEC compares the commanded position (output signal) and the actual position (feedback signal) of the output device. An error between these two signals shows that the output device has not gone to the commanded position. If the error is in one channel only ;the fault is most likely in the output device, if the fault is in both channels it indicates either a failure or a mechanical problem, fouling, bent control rod etc* Cross Check The ESC compares the outputs of the Channel A and Channel B feedback devices (LVDT's, microswitches). A difference indicates and internal unit fault. Fault Indications A fault detected by the Bite generates two messages:• a clear language message displayed on CFDS • a unique alphanumeric fault code CHANNEL A DATA LINK CHANNEL B CONDITlON (FEEDBACK) SIGNAL EEC 1. TRACK CHECK √ or X 2. CROSS CHECK √ or X 3. WSAPAROUND CHECK √ or X LVDT CONTROL SIGNAL CFDS- HEALTH MONITORING A320 TROUBLESHOOTING CFDS - Menu Mode Detailed information about failures can be obtained from the CFDS, operating in the Menu mode, through either of the MCDU's. Menus are displayed on the MCDU. Selection of the appropriate item is made by pushing the keys located alongside the display. The CFDS menu mode functions are:Last Leg Report: Displays all the class 1 and class 2 failures of the last leg, up to a maximum of 20. Last Leg ECAM Report: Displays the ECAM warnings experienced during the last leg - up to a maximum of 20. Previous Leg Report This is a copy of the 63 previous legs maximum capacity of 200 failures, (whichever is first). Avionics Status Provides a real time display of all systems affected by internal or external failures. Post Flight Print Prints a copy of the Last Leg Report plus the Last Leg ECAM Report. System Report/Test Allows maintenance personnel to interact with the chosen system, through the CFDIU. CFDS- MENU MODE A3 20 TROUBLESHOOTING Fault Diagnosis The following pages describe the use of the CFDS to confirm and diagnose faults. For example, a fault message which reads:ENG 1 COMPRESSOR VANE has appeared on the ECAM during flight and has been written in the tech. log by the pilot. Several faults could generate the same message. By utilising the CFDS the specific problem can be identified and further fault data, to assist in fault diagnosis, can be obtained. The troubleshooting sequence is as follows:Gain access to the flight deck, then: 1 turn on the FADEC power using the switch on the overhead panel - RH side. 2 press MCDU menu key 3 select CFDS 4 select POST FLIGHT PRINT The post flight report, see the example below, records the ECAM messages and the fault messages. The fault message is identified by cross checking the times, and is a clear language message - in this case 2.5 BLEED ACT/HC/EECl This tells us that the specific fault is in the 2.5 (BSBV) bleed system. POST FLIGHT REPORT A320 TROUBLESHOOTING Fault Diagnosis (Coiit) The clear language message obtained from the post flight report provides the basis for further troubleshooting. At this point reference is Troubleshooting Manual (TSM). made to the A320 An extract from the TSM is shown below. When using the TSM use the clear language CFDS message to locate the correct page. The TSM presents 3 options for the cause of this particular fault:• • • the 2.5 bleed actuator a harness or connector fault the EEC. Additional data, to identify which option to take, can be obtained from the CFDS. This data is in the form of an alpha / numeric fault code which comprises two letters/digits e.g. D9, 35, 3C or 5F. The procedure used to obtain the fault code is detailed, step by step, in the following pages. A 320 TSM EXTRACT A320 TROUBLESHOOTING Fault Codes To obtain a fault code for a CFDS fault indication proceed as follows:Gain access to the flight deck, then:1. turn on the FADEC power using the switch on the overhead panel RH side 2. Press MCDU menu key 3. select CFDS 4. select SYSTEMREPORT/TEST 5. press NEXT PAGE key on MCDU 6. select ENG 7. select FADEC I A Note: Always interrogate FADEC A and B for the appropriate engine 8. select LAST LEG REPORT The sequence is shown below and is continued on next page. CFDS- FAULT CODE ACCESS A320 TROUBLESHOOTING Fault Codes (Cant) 9. an example of a LAST LEG REPORT is shown below. Look for the clear language message which was shown on the CFDS i.e. 2.5 BLD ACT/HC/EECl identify, and note, the fault cell number (30) 10. press return key to go back to FADEC IA menu 11. (cont on next page) FAULT CELL-IDENTIFICATION A320 TROUBLESHOOTING Fault Codes (Cont) 11 select TROUBLESHOOTING 12 select FLIGHT DATA 13 look for a fault cell which has the same number as the fault in question (found in step 9) i.e. CELL 30 NB there may be more than one page, if the fault cell number does not appear, use the NEXT PAGE key on the MCDU to move through all the fault cells 14 identify word Number 1, as shown below, and note the two right hand letter / digit(s) i.e. 3C Note On EECS with a software code prior to SCN11E (approx January 1991) word six was used to obtain this data. FAULT CODE IDENTIFICATION A320 TROUBLESHOOTING Fault Code Identification The fault code 3C troubleshooting data. is used to provide further Reference to fault code lists shows that the fault code 3C means that the 2.5 bleed actuator has failed the track check. This indicates a mechanical fault, the actuator has not gone to the commanded position. An examination of the 2.5 bleed actuator may reveal an actuator foul or mechanical damage, if this is not the case, change the 2.5 bleed master actuator. Note • Three other faults could have produced the same clear language message, these fault codes are:- • 5F 2.5 bleed cross check failed. To rectify this fault change the 2.5 bleed valve master actuator • 35 2.5 bleed torque motor wrap around failure usually indicates a harness or connector problem • D9 Local 2.5 LVDT latched failed change actuator FAULT CODE A3 20 TROUBLESHOOTING Harness / Electrical Faults These faults usually cause the system to fail the Wraparound check. Reference to the troubleshooting manual will detail the investigation (method and sequence) to be carried out. This can be seen below. In our example the fault code was 3C. The TSM details the specific harness checks to be carried out for this fault code and also provides the locator reference for the Aircraft Schematic Manual (ASM>, i.e. ASM 73-25-00 The ASM provides a schematic circuit diagram which identifies and locates the connectors, cables, etc for this circuit. The Aircraft Maintenance Manual (AMM) locator reference is also quoted i.e. AMM 71-50-00, This details the specific harness checks to be carried out i.e. continuity checks, visual checks, resistance checks etc and also explains how these should be accomplished. Should any defective circuits or damaged parts be found the AMM will provide the locator reference for the Aircraft Wiring Manual (AWM) i.e. 20-71-00. The AWM provides explicit instructions on how to carry out the necessary repairs/replacements. DATA PATH - HARNESS FAULT A3 20 TROUBLESHOOTING FADEC Test The FADEC system self test if carried out from the flight deck through the MCDU using the CFDS in the MENU mode. The procedure is shown, step by step, below. With regard to the FADEC self test there are two points which must be clearly understood:1. if fault rectification has involved a component change or a repair the FADEC self test alone does not satisfy the test requirements. Additional test, some of which involve an engine ground run, may also be required. This is explained in more detail on the next page - see Engine Testing After Fault Rectification. 2. if the FADEC self test shows "Test Failed", it does not mean the fault is within the EEC. For example, part of the FADEC self test are wraparound checks, and a harness fault, a loose plug, for instance, could register a "Test Failed" indication. FADIC SELF TEST A320 TROUBLESHOOTING Engine Testing After Fault Rectification On completion of Power Plant, Module, Component, repairs or replacement procedures some further tests may be required before the aircraft is returned to service. The tests required after specific actions are detailed in the Aircraft Maintenance Manual (AMM) in:CH 71-00-00 Page Block 200 Maintenance Practices as shown below. In our example the 2.5 Bleed, master actuator has been changed. Reference to AMM 71-00-00 PB 200 shows that tests 3 or 1, and 11 are required. test 1 : Dry motor leak check. test 3 : Idle leak check. test 11: High Power assurance test () The test procedures are also detailed under the same AMM reference. An extract from the AMM is shown below and on the next page. AMM EXTRACT (I) AMM EXTRACT (2) V2500 NACELLE PART TWO - SECTION 4 NACELLE VENTILATION / OVERHEAT & FIRE PROTECTION NACELLE Nacelle Ventilation Ventilation is provided for the fan compartment (Zone 1), and the core compartment (Zone 2), to: • prevent accessory and component overheating • prevent the accumulation of flammable vapours Zone 1 Ventilation Ram air enters the zone through an inlet located on the upper LH side of the air intake cowl. The air circulates through the fan compartment and exits at the exhaust located on the bottom rear centre line of the fan cowl doors. Zone 2 Ventilation The ventilation of Zone 2 is provided by air exhausting from the Active Clearance Control (ACC.) system around the turbine area. The air circulates through the core compartment and exits through the lower bifurcation of the 'C ducts. Ventilation during Ground Running During ground running local pockets of natural convection exist providing some ventilation of the fancase - Zone 2. Zone 2 ventilation is still effected in the same way as when the engine is running. PNEUMATIC DUCT LEAKS FAN CASE COMPARTMENT VENTILATION EXIT VENTILATION INL£T PRESSURE RELIEF DOOR (TYPICAL 2 PLACES) THRUST REVERSER SEAL LEAKS ZONE 1 ZONE 2 CORE ENGINE COMPARTMENT PRESSURE RELIEF DOOR VENTILATION EXIT FAN CASE AND CORE COMPARTMENT VENTILATION NACELLE Fire Detection System The fire detection system monitors the air temperature in Zone 1 and Zone 2. When the air temperature increases to a predetermined level the system provides flight deck warning by: • master warning light • audible warning • specific fire indications Zone 1 and Zone 2 fire detectors function independently of each other. Each zone has two detector units which are mounted as a pair, each unit gives an output signal when a fire or overheat condition occurs. The two detector units are attached to support tubes by clips. ENGINE FIRE DETECTION SYSTEM - OVERALL VIEW NACELLE Fire Detection System Detector Units Responder Each detector unit comprises:- The responder has two pressure switches, one normally open, the other normally closed. • a hollow sensor tube • a responder assembly Sensor Tube The sensor tube is closed and sealed at one end, the other open end is connected to the responder. The tube is filled with Helium and carries a central core of ceramic material impregnated with Hydrogen. An increase in the air temperature around the sensor tube causes the Helium to expand and increase the pressure within the tubes, further increases in temperature cause the core material to expel Hydrogen to increase the pressure within the tube. Both pressure switches sense the gas pressure in the sensor tube. The responder is connected via its electrical receptacle and the harness to an electronic fire detectLon module. ELECTRICAL CONNECTOR NORMAL OPEN SWITCH RESPONDER CORE ELEMENT CERAMIC IMPREGNATED WITH HYDROGEN SENSOR TUBE NORMALLY CLOSED SWITCH TEST AIRCRAFT ELECTRICAL SYSTEM 28V DC FIRE DETECTOR UNIT - SCHEMATIC GAS-HELIUM NACELLE Fire Detection System Sensor Setting Limits The operating temperature limits for Zone 1 and Zone 2 are shown below. ACCESSORY ZONE CORE ZONE 12 INCHES IMMERSED 371 ± 55° C 620 ± 55° C FULLY IMMERSED 235 ± 14° C 370 ± 22° C MARGIN MARGIN 125° C ZONE MAXIMUM TEMPERATURE 260° C ZONE MAXIMUM TEMPERATURE FIRE DETECTOR UNIT - SETTING RANGE NACELLE Fire Detection System Indications and Controls The fire indications and controls are located as shown below. FIRE PROTECTION-INDICATIONS/CONTROLS NACELLE Fire Detection and Nacelle Temperature Monitoring Electrical Harness The fire detection nacelle temperature indication electrical harness connections are as shown below. NACELLE TEMPERATURE SENSING AND FIRE DETECTION HARNESS V2S00 NACELLE PART TWO - SECTION 5 ENGINE REMOVAL/INSTALLATION EngIne Change The arrangements for slinging/hoisting the engine are shown below. NOTE During this operation the C-ducts are supported by (GSE) Ground Service Equipment rods which are positioned between the C-duct and the aircraft pylon. ENGINE REMOVAL AND INSTALLATION V2500 GENERAL PART THREE - SECTION 4 BORESCOPING LP COMPRESSOR OUTLET GUIDE VANE STAGE 1 FAN BLADE FAN FRAME STRUT FLEXIBLE BORESCOF FAN OUTLET INNER VANE LP COMPRESSOR STAGE 1,5 BLADE Examine the Front Surfaces of the Stage ? LP Compressor Blades Fig 602/TASK 72-00-00-991-156 INSPECTION/CHECK R EFFECTIVITY: V2500 72-00-00 Page 612 INLET GUIDE— VANE LP COMPRESSOR STAGE 2.5 BLADE FAN FRAME STRUT BORESCOPE FLEXIBLE IP Compressor Stage 2.5 Blades - Inspection/Check Fig 603/TASK 72-O0-OO-991-157 INSPECTION/CHECK R EFFECTIVITY: V2500 72-00-00 Page 613 INTERNATIONAL AERO ENGINES V2500 Propulsion System ENGINE MAINTENANCE MANUAL