A330 ATA 71-80 RR Trent 700 L3 e.pdf

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Lufthansa Technical Training GmbH Lufthansa Base Frankfurt D-60546 Frankfurt/Main Tel. +49 69 / 696 41 78 Fax +49 69 / 696 63 84 Lufthansa Base Hamburg Weg beim Jäger 193 D-22335 Hamburg Tel. +49 40 / 5070 24 13 Fax +49 40 / 5070 47 46

TABLE OF CONTENTS ATA 71-80 700 . . . .

MAIN ROTATING ASSEMBLIES . . . . . . . . . . . . . . . . . . . . . ENGINE MAIN BEARING ARRANGEMENT . . . . . . . . . . . MODULE BREAKDOWN OF ENGINE . . . . . . . . . . . . . . . . L.P. COMPRESSOR MODULE . . . . . . . . . . . . . . . . . . . . . . FAN BLADE, ANNULUS FILLER, SPINNER, FAIRING AND MAKE–UP PIECE INSPECTION/CHECK . . . . . . . . AIR INTAKE FAIRING/SPINNER AND MAKE–UP PIECE REMOVAL/INSTALLATION . . . . . . . . . . . . . . . . . . . LP COMPRESSOR (FAN)BLADE REMOVAL / INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I.P. COMPRESSOR MODULE . . . . . . . . . . . . . . . . . . . . . . . INTERMEDIATE CASE MODULE . . . . . . . . . . . . . . . . . . . . COMPRESSOR FAIRINGS / ’A’ FRAME STRUTS . . . . . H.P. SYSTEM MODULE . . . . . . . . . . . . . . . . . . . . . . . . . . . . I.P. TURBINE MODULE . . . . . . . . . . . . . . . . . . . . . . . . . . . . L.P. TURBINE MODULE . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPRING LOADED L.P. TURBINE BEARING . . . . . . . . . . H S GEARBOX MODULE . . . . . . . . . . . . . . . . . . . . . . . . . . L.P. COMPRESSOR CASE MODULE . . . . . . . . . . . . . . . .

ENGINE RR TRENT 1

ATA 71- 80 ENGINE RR TRENT 700 . . . . . . . . . . . .

2

GENERAL INFORMATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RB 211 FAMILY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .................................................. ENGINE LEFT HAND VIEW . . . . . . . . . . . . . . . . . . . . . . . . ENGINE RIGHT HAND VIEW . . . . . . . . . . . . . . . . . . . . . . .

2 2 2 4 6

ENGINE SPECIFIC DATA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING PARTICULARS . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE OPERATING LIMITS AND GUIDELINES . . . . . GENERAL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . .

8 8 8 10

72-00

ENGINE PRESENTATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . MAJOR UNITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIR INTAKE COWL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . REVERSER COWL ( C-DUCT ) . . . . . . . . . . . . . . . . . . . . . LEFT AND RIGHT HAND FAN COWL DOORS . . . . . . . . FAN COWL DOOR OPENING AND CLOSING . . . . . . . . C - DUCT OPENING AND CLOSING . . . . . . . . . . . . . . . . . C-DUCT LATCH NO.1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C-DUCT LATCH 3 AND 4 . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MOUNTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMMON NOZZLE ASSEMBLY ( CNA ) . . . . . . . . . . . . . ACCESSIBILITY (LEFT SIDE) . . . . . . . . . . . . . . . . . . . . . . . ACCESSIBILITY (RIGHT SIDE) . . . . . . . . . . . . . . . . . . . . .

12 12 14 16 18 20 22 24 24 26 30 32 34

71-70

ENGINE DRAINS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPERATION OF DRAINS TANK . . . . . . . . . . . . . . . . . . . . . COMPONENT LOCATION . . . . . . . . . . . . . . . . . . . . . . . . . .

72-00

ENGINE PRESENTATION . . . . . . . . . . . . . . . . . . . . . . . . . . . .

42 44 46 48 50 52 54 56 58 60 62 64 66 68 70 72

72-00

ENGINE BORESCOPING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BORESCOPE ACCESS PORTS . . . . . . . . . . . . . . . . . . . . . IP HAND TURNING TOOL . . . . . . . . . . . . . . . . . . . . . . . . . . IP BORESCOPE PLUGS . . . . . . . . . . . . . . . . . . . . . . . . . . . VIGV ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TURNING THE HIGH PRESSURE (H.P.) SYSTEM . . . . COMBUSTION LINER BORO PLUGS . . . . . . . . . . . . . . . . HP-TURBINE BORO PLUGS . . . . . . . . . . . . . . . . . . . . . . . . HP AND IP TURBINE BORO PLUG . . . . . . . . . . . . . . . . . . LP TURBINE BORO PLUG . . . . . . . . . . . . . . . . . . . . . . . . .

74 74 76 78 80 82 86 86 88 90

36 36 38 40

77-00

ENGINE INDICATION PRESENTATION . . . . . . . . . . . . . . . . ENGINE/WARNING DISPLAY . . . . . . . . . . . . . . . . . . . . . . . SYSTEM DISPLAY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESSURE AND TEMPERATURE STATIONS . . . . . . . .

92 92 94 96

42

SHAFT SPEED . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENT LOCATION . . . . . . . . . . . . . . . . . . . . . . . . . .

98 100

Page: i

TABLE OF CONTENTS N1 INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N2 INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N3 INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

102 102 104

POWER MEASUREMENT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPR INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWER MEASUREMENT . . . . . . . . . . . . . . . . . . . . . . . . . . EXHAUST GAS TEMPERATURE . . . . . . . . . . . . . . . . . . . .

106 108 110 112

VIBRATION MONITORING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE INTERFACE AND VIBRATION MONITORING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIBRATION MONITORING . . . . . . . . . . . . . . . . . . . . . . . . .

118 118

EIVMU .

126 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIVMU CONTINUED . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN UNBALANCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MAX FLIGHT VIBRATION DISPLAY . . . . . . . . . . . . . . . . . FREOUENCY ANALYSIS . . . . . . . . . . . . . . . . . . . . . . . . . . . DISCRETE INPUTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DISCRETE OUTPUTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE CONDITION MONITORING . . . . . . . . . . . . . . . .

120 122

THROTTLE CONTROL LEVER MECHANISM . . . . . . . . . BASIC CONTROL LOOP–STEADY STATE . . . . . . . . . . . THRUST MODES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST SETTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE RATING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CMS EEC INTERACTIVE TESTS . . . . . . . . . . . . . . . . . . . .

ATA 79 OIL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192 79-00

OIL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SYSTEM DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . FEED OIL, LUBRICATION AND COOLING . . . . . . . . . . . VENTING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL TANK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SCAVANGE FILTER ASSEMBLY . . . . . . . . . . . . . . . . . . . . OIL PUMP / SCAVENGE FILTER . . . . . . . . . . . . . . . . . . . . OIL PUMP / MCD HOUSINGS . . . . . . . . . . . . . . . . . . . . . . CENTRIFUGAL BREATHER . . . . . . . . . . . . . . . . . . . . . . . . HEAT MANAGEMENT SYSTEM . . . . . . . . . . . . . . . . . . . . . FUEL COOLED OIL COOLER . . . . . . . . . . . . . . . . . . . . . . . AIR OIL HEAT EXCHANGER DEACTIVATION . . . . . . . . I.D.G. OIL COOLING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIR COOLED OIL COOLER . . . . . . . . . . . . . . . . . . . . . . . . FILLING THE ENGINE OIL SYSTEM . . . . . . . . . . . . . . . . . M.C.D. REMOVAL, INSPECTION AND REPLACEMENT PRESSURE OIL FILTER . . . . . . . . . . . . . . . . . . . . . . . . . . .

192 192 194 196 198 200 202 204 206 208 212 214 216 218 220 222 224

79-30

OIL INDICATING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

226 226

126 128 130 132 134 136 138 140

ATA 73 ENGINE FUEL AND CONTROL . . . . . . . . 142 73-20

FADEC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DEDICATED ALTERNATOR . . . . . . . . . . . . . . . . . . . . . . . . POWER CONTROL UNIT . . . . . . . . . . . . . . . . . . . . . . . . . . ELECTRONIC ENGINE CONTROL . . . . . . . . . . . . . . . . . . E.E.C. INTEGRITY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DEP PROGRAMMING UNIT . . . . . . . . . . . . . . . . . . . . . . . . OVERSPEED PROTECTION SYSTEM (OPU) . . . . . . . . TURBINE OVERSPEED PROTECTION SYSTEM . . . . . P20 / T20 PROBE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWER MANAGEMENT . . . . . . . . . . . . . . . . . . . . . . . . . . .

142 144 146 150 154 158 160 164 168 170

172 176 178 180 180 182

ATA 73 ENGINE FUEL AND CONTROL . . . . . . . . 238 73-00

FUEL SYSTEM PRESENTATION . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

238 238

Page: ii

TABLE OF CONTENTS FUEL FUEL FUEL FUEL

COMPONENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPRAY NOZZLES . . . . . . . . . . . . . . . . . . . . . . . . . . .

240 242 246 248

ATA 75 AIR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 258 75-33

IP/HP COMPRESSOR AIRFLOW CONTROL . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIGV / VSV SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . V.I.G.V./V.S.V. OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . COMPRESSOR BLEED CONTROL SYSTEM . . . . . . . . . COMPRESSOR BLEED VALVES . . . . . . . . . . . . . . . . . . . .

258 258 260 262 268 270

75-20

TURBINE CASE COOLING . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TCC ACTUATOR AND VALVE . . . . . . . . . . . . . . . . . . . . . . . TCC COOLING SOLENOID . . . . . . . . . . . . . . . . . . . . . . . . . TCC MANIFOLD AND COOLING LINER . . . . . . . . . . . . . . ENGINE COOLING AND SEALING SYSTEM . . . . . . . . . BEARING COMPARTMENT COOLING SYSTEM . . . . . . COMPONENT DESCRIPTION AND LOCATION . . . . . . . TURBINE OVERHEAT DETECTION SYSTEM . . . . . . . .

280 280 282 282 284 286 288 292 294

75-00

ACCESSORY COOLING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SYSTEM DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . NACELLE TEMPERATURE INDICATION . . . . . . . . . . . . .

296 296 298 300

ARTIFICIAL FEEL UNIT . . . . . . . . . . . . . . . . . . . . . . . . . . . . THROTTLE CONTROL UNIT . . . . . . . . . . . . . . . . . . . . . . .

310 310

ATA 30 ICE AND RAIN PROTECTION . . . . . . . . . . 312 30-20

ENGINE AIR INTAKE ICE PROTECTION . . . . . . . . . . . . . . . GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ANTI–ICING SYSTEM OPERATION . . . . . . . . . . . . . . . . .

312 312 314

ATA 78 EXHAUST . . . . . . . . . . . . . . . . . . . . . . . . . . . 320 78-00

THRUST REVERSER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER INDICATION . . . . . . . . . . . . . . . . . . SYSTEM OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HYDRAULIC CONTROL . . . . . . . . . . . . . . . . . . . . . . . . . . . . HYDRAULIC OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . PRIMARY LOCKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ACTUATORS AND SECONDARY LOCKS . . . . . . . . . . . . MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER CMS TEST . . . . . . . . . . . . . . . . . . .

320 320 322 324 326 328 332 334 338 340 352

ATA 74 IGNITION SYSTEM . . . . . . . . . . . . . . . . . . . 356 74–00

IGNITION SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITER PLUG REMOVAL/INSTALLATION . . . . . . . . . . .

356 360

ATA 76 ENGINE CONTROLS . . . . . . . . . . . . . . . . . 306 76-00

GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .................................................. HP FUEL SHUTOFF VALVE . . . . . . . . . . . . . . . . . . . . . . . . LP FUEL SHUTOFF VALVE . . . . . . . . . . . . . . . . . . . . . . . . .

306 306 308 308

ATA 80 STARTING . . . . . . . . . . . . . . . . . . . . . . . . . . . 364 80-00

STARTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTING SYSTEM INTRODUCTION . . . . . . . . . . . . . . AIR STARTER MOTOR . . . . . . . . . . . . . . . . . . . . . . . . . . . .

364 364 368

Page: iii

TABLE OF CONTENTS START CONTROL VALVE . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE START CONTROL AND INDICATION . . . . . . . . START PROCEDURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

370 376 378

Page: iv

TABLE OF FIGURES Figure 1 Figure 2 Figure 3 Figure 4 Figure 5 Figure 6 Figure 7 Figure 8 Figure 9 Figure 10 Figure 11 Figure 12 Figure 13 Figure 14 Figure 15 Figure 16 Figure 17 Figure 18 Figure 19 Figure 20 Figure 21 Figure 22 Figure 23 Figure 24 Figure 25 Figure 26 Figure 27 Figure 28 Figure 29 Figure 30 Figure 31 Figure 32 Figure 33 Figure 34 Figure 35

The RB 211 Family . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Left Hand View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Right Hand View . . . . . . . . . . . . . . . . . . . . . . . . . . . External Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Propulsion System Outline . . . . . . . . . . . . . . . . . . . . . . . . . Engine Major Units . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Nose Cowl . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Assembly . . . . . . . . . . . . . . . . . . . . . . . . . Fan Cowl Doors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Cowl Door Latches and Rods . . . . . . . . . . . . . . . . . . C - Duct Latches / Hydraulic Manifold / Open Rods . . . C - Duct Latches 1, 3 and 4 . . . . . . . . . . . . . . . . . . . . . . . Front Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rear Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Common Nozzle Assembly . . . . . . . . . . . . . . . . . . . . . . . . Left Side Accessibility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Right Side Accessibility . . . . . . . . . . . . . . . . . . . . . . . . . . . Drains System Simplified . . . . . . . . . . . . . . . . . . . . . . . . . . Drains Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drains Tank / Drains Mast . . . . . . . . . . . . . . . . . . . . . . . . . Main Rotating Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . Engine Bearing Arrangement . . . . . . . . . . . . . . . . . . . . . . Trent Modular Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . LP Compressor Module . . . . . . . . . . . . . . . . . . . . . . . . . . . Fan Blade, Spinner Fairing, Make-up Piece . . . . . . . . . . Support Ring, Slider Assembly and Annulus Fillers . . . Fan Blade Replacement Sequence . . . . . . . . . . . . . . . . . IP Compressor Module . . . . . . . . . . . . . . . . . . . . . . . . . . . Intermediate Case Module . . . . . . . . . . . . . . . . . . . . . . . . Compressor Fairings / A Frame Struts . . . . . . . . . . . . . . HP System Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP Turbine Case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LP Turbine Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spring Loaded LP Turbine Bearing . . . . . . . . . . . . . . . . . External Gearbox Module . . . . . . . . . . . . . . . . . . . . . . . . .

3 5 7 9 11 13 15 17 19 21 23 25 27 29 31 33 35 37 39 41 43 45 47 49 51 53 55 57 59 61 63 65 67 69 71

Figure 36 Figure 37 Figure 38 Figure 39 Figure 40 Figure 41 Figure 42 Figure 43 Figure 44 Figure 45 Figure 46 Figure 47 Figure 48 Figure 49 Figure 50 Figure 51 Figure 52 Figure 53 Figure 54 Figure 55 Figure 56 Figure 57 Figure 58 Figure 59 Figure 60 Figure 61 Figure 62 Figure 63 Figure 64 Figure 65 Figure 66 Figure 67 Figure 68 Figure 69 Figure 70

LP Compressor Case Module . . . . . . . . . . . . . . . . . . . . . . Borescope Access Ports . . . . . . . . . . . . . . . . . . . . . . . . . . IP Hand Turning Tool . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP Borescope Plugs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIGV Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hand Turning of the HP System . . . . . . . . . . . . . . . . . . . . HP Borescope Plugs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Combustion Liner / HPT Boro Plugs . . . . . . . . . . . . . . . . HP / IP Turbine Boro Plugs . . . . . . . . . . . . . . . . . . . . . . . . LP Turbine Boro Plugs . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Primary Display . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine System Display . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressure and Temperature Stations . . . . . . . . . . . . . . . . Shaft Speed Components . . . . . . . . . . . . . . . . . . . . . . . . . Shaft Speed Components . . . . . . . . . . . . . . . . . . . . . . . . . N1 Indication / N2 Indication . . . . . . . . . . . . . . . . . . . . . . . N3 Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPR System P20 / P50 Probes . . . . . . . . . . . . . . . . . . . . EPR Indication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPR Trimming . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Simplified EGT Diagram / ECAM Indication . . . . . . . . . . EGT Thermocouples . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGT Trimming . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Simplified Vibration Diagram . . . . . . . . . . . . . . . . . . . . . . . EIVM System Architecture . . . . . . . . . . . . . . . . . . . . . . . . Vibration Transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remote Charge Converter . . . . . . . . . . . . . . . . . . . . . . . . EIVMU System Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIVMU Specific Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIVMU Fan Unbalance . . . . . . . . . . . . . . . . . . . . . . . . . . . EIVMU Max Flight Vibration Display . . . . . . . . . . . . . . . . Frequency Analysis Readout . . . . . . . . . . . . . . . . . . . . . . Discrete Inputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Discrete Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thermocouples ( T30 / T25 ) . . . . . . . . . . . . . . . . . . . . . .

73 75 77 79 81 83 85 87 89 91 93 95 97 99 101 103 105 107 109 111 113 115 117 119 121 123 125 127 129 131 133 135 137 139 141 Page: v

TABLE OF FIGURES Figure 71 Figure 72 Figure 73 Figure 74 Figure 75 Figure 76 Figure 77 Figure 78 Figure 79 Figure 80 Figure 81 Figure 82 Figure 83 Figure 84 Figure 85 Figure 86 Figure 87 Figure 88 Figure 89 Figure 90 Figure 91 Figure 92 Figure 93 Figure 94 Figure 95 Figure 96 Figure 97 Figure 98 Figure 99 Figure 100 Figure 101 Figure 102 Figure 103 Figure 104 Figure 105

FADEC Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dedicated Alternator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC Electrical Power Configuration . . . . . . . . . . . . . Power Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electronic Engine Controller . . . . . . . . . . . . . . . . . . . . . . . EEC Suitcase . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . High Integrity Computer Schematic . . . . . . . . . . . . . . . . . Data Entry Plug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DEP Programming Unit and Printouts . . . . . . . . . . . . . . . Overspeed Protection Logic . . . . . . . . . . . . . . . . . . . . . . . Overspeed Protection Unit . . . . . . . . . . . . . . . . . . . . . . . . Overspeed Pretection Simplified Diagram . . . . . . . . . . . Turbine Overspeed Protection . . . . . . . . . . . . . . . . . . . . . P20 / T20 Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Combined TRA Relationship . . . . . . . . . . . . . . . . . . . . . . . Power Setting - Basic Control Loop . . . . . . . . . . . . . . . . . Forward Thrust - Throttle Legend . . . . . . . . . . . . . . . . . . . Flat Rating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ground Scanning / Class 3 Faults / Ground Report . . . EEC System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EEC Probe Heater Test . . . . . . . . . . . . . . . . . . . . . . . . . . . CMS Engine Runningn Test . . . . . . . . . . . . . . . . . . . . . . . Specific Data Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Oil Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Scavenge Filter Assembly . . . . . . . . . . . . . . . . . . . . . . . . Oil Pump and Scavenge Filter . . . . . . . . . . . . . . . . . . . . Oil Pump / MCD Housings . . . . . . . . . . . . . . . . . . . . . . . Centrifugal Breather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Heat Management System . . . . . . . . . . . . . . . . . . . . . . . AOHE Air Modulating Valve . . . . . . . . . . . . . . . . . . . . . .

143 145 147 149 151 153 155 157 159 161 163 165 167 169 171 173 175 177 179 181 183 185 187 189 191 193 195 197 199 201 203 205 207 209 211

Figure 106 Figure 107 Figure 108 Figure 109 Figure 110 Figure 111 Figure 112 Figure 113 Figure 114 Figure 115 Figure 116 Figure 117 Figure 118 Figure 119 Figure 120 Figure 121 Figure 122 Figure 123 Figure 124 Figure 125 Figure 126 Figure 127 Figure 128 Figure 129 Figure 130 Figure 131 Figure 132 Figure 133 Figure 134 Figure 135 Figure 136 Figure 137 Figure 138 Figure 139 Figure 140

Fuel Cooled Oil cooler . . . . . . . . . . . . . . . . . . . . . . . . . . . Air / Oil Heat Exchanger . . . . . . . . . . . . . . . . . . . . . . . . . Integrated Drive Generator . . . . . . . . . . . . . . . . . . . . . . . Air cooled Oil Cooler . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Oil Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Master MCD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Quantity Transmitter . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Temperature Sensors . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Transmitters . . . . . . . . . . . . . . . . . . . . . . . . Low Oil Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . Oil Pressure Filter Differential Pressure Switch . . . . . . Oil Scavenge Filter Differential Pressure Switch . . . . . Fuel System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . FCOC and Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Pump Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Metering Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Spray Nozzle and Manifold . . . . . . . . . . . . . . . . . . . Fuel Spray Nozzle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Spray Nozzle Removal / Installation . . . . . . . . . . . Fuel Temperature Thermocouples . . . . . . . . . . . . . . . . . LP Fuel Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HP Fuel Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIGV / VSV Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIGV / VSV System Schematic . . . . . . . . . . . . . . . . . . . VIGV / VSV Actuation Schematic . . . . . . . . . . . . . . . . . VSV Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CMS VSV Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IP and HP Bleed Valves . . . . . . . . . . . . . . . . . . . . . . . . . . IP and HP Bleed Valves . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Valve System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Valve Solenoid Pack . . . . . . . . . . . . . . . . . . . . . . . Bleed Valve Open . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Valve Closed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbine Case Cooling Schematic . . . . . . . . . . . . . . . . . .

213 215 217 219 221 223 225 227 229 231 233 235 237 239 241 243 245 247 249 251 253 255 257 259 261 263 265 267 269 271 273 275 277 279 281 Page: vi

TABLE OF FIGURES Figure 141 Figure 142 Figure 143 Figure 144 Figure 145 Figure 146 Figure 147 Figure 148 Figure 149 Figure 150 Figure 151 Figure 152 Figure 153 Figure 154 Figure 155 Figure 156 Figure 157 Figure 158 Figure 159 Figure 160 Figure 161 Figure 162 Figure 163 Figure 164 Figure 165 Figure 166 Figure 167 Figure 168 Figure 169 Figure 170 Figure 171

TCC Actuator / Valve and Solenoid . . . . . . . . . . . . . . . . TCC Manifold and Cooling Liner . . . . . . . . . . . . . . . . . . Cooling and Sealing Airflows . . . . . . . . . . . . . . . . . . . . . ACAC Open Position . . . . . . . . . . . . . . . . . . . . . . . . . . . . ACAC Closed Position . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Cooled Air Cooler . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbine Overheat Detection Probes . . . . . . . . . . . . . . . Fireproof Bulkheads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Accessory cooling Zones . . . . . . . . . . . . . . . . . . . . . . . . . Zone 3 Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . Fire Seals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EEC Unit Cooling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Throttle Control System . . . . . . . . . . . . . . . . . . . . . . . . . . HP and LP Fuel Shutoff Valve . . . . . . . . . . . . . . . . . . . . Throttle Control / Artificial Feel Unit . . . . . . . . . . . . . . . . Anti Icing System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anti Icing Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anti Ice Valve Schematic . . . . . . . . . . . . . . . . . . . . . . . . Anti Ice Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Assembly . . . . . . . . . . . . . . . . . . . . . . . Thrust Reverser Control Indication . . . . . . . . . . . . . . . . Thrust Reverser Actuation Diagram . . . . . . . . . . . . . . . Hydraulic Control Fwd Thrust Position . . . . . . . . . . . . . Hydraulic Control Commanding Deploy . . . . . . . . . . . . Hydraulic Control Deploy Position . . . . . . . . . . . . . . . . . Primary Lock Mechanism . . . . . . . . . . . . . . . . . . . . . . . . TR Actuator Commanding Extending . . . . . . . . . . . . . . TR Actuation System Schematic . . . . . . . . . . . . . . . . . . TR Activation/ De–Activation . . . . . . . . . . . . . . . . . . . . . TR In–Flight Lock Out . . . . . . . . . . . . . . . . . . . . . . . . . . . Primary, Secondary and Tertiary Locks Release Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 172 Pivoting Door RVDT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 173 Pvoting Door Stow Switch . . . . . . . . . . . . . . . . . . . . . . . . Figure 174 Tertiary Lock/ Power Conditioning Module . . . . . . . . . .

283 285 287 289 291 293 295 297 299 301 303 305 307 309 311 313 315 317 319 321 323 325 327 329 331 333 335 337 339 341

Figure 175 Figure 176 Figure 177 Figure 178 Figure 179 Figure 180 Figure 181 Figure 182 Figure 183 Figure 184 Figure 185 Figure 186 Figure 187 Figure 188 Figure 189 Figure 190 Figure 191

Tertiary Lock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CMS Thrust Reverser Test 1 . . . . . . . . . . . . . . . . . . . . . CMS Thrust Reverser Test 2 . . . . . . . . . . . . . . . . . . . . . CMS Thrust Reverser Test 3 . . . . . . . . . . . . . . . . . . . . . Ignition System Components . . . . . . . . . . . . . . . . . . . . . Igniter Plug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Igniter Plug Depth of Immersion Check . . . . . . . . . . . . CMS Ignition Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starting System Overview . . . . . . . . . . . . . . . . . . . . . . . . Starter Motor and Air Duct . . . . . . . . . . . . . . . . . . . . . . . Air Starter Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Start Control Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Start Valve: Closed Position . . . . . . . . . . . . . . . . . . . Air Start Valve: Open Position . . . . . . . . . . . . . . . . . . . . Engine Start Control and Indications . . . . . . . . . . . . . . . Auto Start Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . Manual Start Procedure . . . . . . . . . . . . . . . . . . . . . . . . . .

351 353 354 355 357 359 361 363 365 367 369 371 373 375 377 379 381

343 345 347 349

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A 330 Trent 700

71-00

ATA 71-80

ENGINE RR TRENT 700

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

FRA US-T TH NOV 99

Page: 1

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ENGINE GENERAL

A 330 Trent 700

71-00

ATA 71- 80 ENGINE RR TRENT 700 GENERAL INFORMATION RB 211 FAMILY The Engine identification is built up from 6 basic blocks: S RB 211 - The family designation S TRENT - Is now used instead of a numerical designation e.g. 524 or 535 S 7 - Series S 68 - Approximate thrust S -A - A letter coding indicating minor performance changes S -60 - Indicates the aircraft the engine is fitted to

For Training Purposes Only

In the above examples the engine would be written as RB 211 TRENT 768-A-60 THRUST LEVELS TRENT- 890 TRENT- 884 TRENT- 877 TRENT- 875 TRENT- 775 TRENT- 772 TRENT- 768 RB 211- 524H RB 211- 524G RB211- 524D4 RB211- 524C2 RB211- 524B4

FRA US-T TH NOV 99

90,000 84,400 77,000 74,600 75,150 71,100 67,500 60,600 58,000 53,000 51,500 50,000

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A 330 Trent 700

71-00

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Lufthansa Technical Training

ENGINE GENERAL

Figure 1 FRA US-T TH NOV 99

The RB 211 Family Page: 3

A 330 Trent 700

71-00 ENGINE LEFT HAND VIEW The following diagram shows the left hand side of the engine.

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ENGINE GENERAL

FRA US-T TH NOV 99

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A 330 Trent 700

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ENGINE GENERAL

Figure 2 FRA US-T TH NOV 99

Engine Left Hand View Page: 5

A 330 Trent 700

71-00 ENGINE RIGHT HAND VIEW The following diagram shows the right hand side of the engine.

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ENGINE GENERAL

FRA US-T TH NOV 99

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A 330 Trent 700

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ENGINE GENERAL

Figure 3 FRA US-T TH NOV 99

Engine Right Hand View Page: 7

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ENGINE GENERAL

71-00 ENGINE SPECIFIC DATA LEADING PARTICULARS Take off thrust S (S.L. Static), Trent 768 67,500 lbs Trent 772 71,100 lbs L.P. System S (N1 Indication) 4 Stage turbine, Single stage Fan I.P. System S (N2 Indication), 8 stage axial flow compressor, single stage turbine H.P. System S (N3 Indication), 6 stage axial flow compressor, single stage turbine Flat rated Temperature S ISA +150C By–pass ratio S Trent 768 4.9 : 1, Trent 772 4.66 : 1 Overall pressure at take off S Trent 768 35.9 : 1, Trent 772 37.42 : 1

For Training Purposes Only

A 330 Trent 700

Overall length S 221 inches / 5613mm Fan diameter S 97.4 inch / 2474mm Powerplant weight S 14350 lbs / 6500Kg

R.P.M’s S 100 % N1 - 3900 RPM S 100 % N2 - 7000 RPM S 100 % N3 - 10611 RPM

ENGINE OPERATING LIMITS AND GUIDELINES EGT Limits S Starting 700_ C S Take Off 920_ C S Max Cont. 850_ C Oil Pressure Limits S Minimum Oil Pressure 25 PSI Oil Temperature Limits S Minimum Temp above Idle 50_ C S Maximum Temp Steady State 170_C S Maximum Temp Transient 190_ C Starter Limits S Cycle: 3 min...............3 min...............1 min S Cooldown: ...........30 sec................30 sec S Extended Cycle: 5 min then 30 min cooldown Vibration Guidelines S N1 Advisory 3.3 Units S N2 Advisory 2.6 Units S N3 Advisory 4.0 Units

Direction of rotation all shafts S Anti clockwise viewed from the rear

FRA US-T TH NOV 99

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A 330 Trent 700

71-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 4 FRA US-T TH NOV 99

External Dimensions Page: 9

Lufthansa Technical Training For Training Purposes Only

ENGINE GENERAL

A 330 Trent 700

71-00 GENERAL DESCRIPTION MAINTENANCE PRACTICES The Maintenance practices explained throughout these course notes may not be a word for word copy of what is described in the maintenance manual, but rather an understanding of the principles of a particular maintenance task. It is therefore essential that these course notes are never used when working on an engine.The official Maintenance Manual must be used at all times. The maintenance manual will list all warnings and cautions at the beginning of each task. In the course notes they will be listed at the beginning of each maintenance section of each chapter. Take time to read these warnings and cautions so that you become familiar with them. Also listed at the end of each chapter is a list of precautions under the heading of ETOPS AWARENESS. ETOPS stands for Extended Range Twin Engine Operation. In the past the only aircraft with the capability of flying long range operations were four engined or three engined. With the development of more powerful turbo fan engines it became possible to operate twin engined aircraft with equivalent long range capability. For the first time the regulatory authorities and manufacturers had to consider the airworthiness and safety implications of twin–engined aircraft operating a long way from any en–route airfield. Typically these routes are over water but there are some over land. The certification rules of all regulatory authorities prevent operations of twin–engined aircraft more than a specific distance (expressed as flying time) from an alternate airfield, en–route. In the early 1980’s the design of twin engine aircraft with long range capability and with reliable engines and systems demanded a rational evaluation of the technical and operation factors involved. It was deemed by the major certification authorities that a threshold of 60 minutes flying time at single engine cruise speed from a suitable alternate airfield was acceptable.

FRA US-T TH NOV 99

PROPULSION SYSTEM OUTLINE Engine power to operate the Airbus A330 is provided by two propulsion systems located through pylons to the underside of the wings. Looking forward from the rear of the aircraft they are numbered one and two, the left hand engine being number one. The Rolls–Royce Trent engine is a 3 shaft high by–pass ratio turbo fan engine with Low Pressure (L.P.), Intermediate Pressure (I.P.) and High Pressure (H.P.) compressors driven by turbines through the co–axial shafts. All the air entering the engine through the air intake cowl passes through the L.P. Compressor (fan) and is then directed into two main flows by the splitter fairing, the cold airflow and the hot gas flow. The cold airflow passes through the fan outlet guide vanes (O.G.V’s) into the by–pass casing and enters the common nozzle assembly (C.N.A.). The air passing through the gas generator also enters the C.N.A. Both flows are exhausted through the CNA to atmosphaere Exhausting both flows through the C.N.A. results in a low velocity jet efflux producing high propulsive efficiency. The L.P. system consists of a single stage wide chord hollow fan blade compressor driven by a 4 stage turbine. The I.P. system consists of an 8 stage axial flow compressor driven by a single stage turbine. The H.P. system consists of a 6 stage axial flow compressor driven by a single stage turbine. The Combustion system is of annular construction incorporating spray nozzles through which fuel is supplied from a fuel – system in accordance with the setting of the engine throttle and aircraft operating conditions.

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ENGINE GENERAL

Figure 5 FRA US-T TH NOV 99

Propulsion System Outline Page: 11

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00

72-00

ENGINE PRESENTATION

MAJOR UNITS The propulsion system comprises of the following items: S S S S S S

Air intake cowl Right and left fan cowl doors Engine, associated fairings, front and rear mounts Common nozzle assembly (C.N.A.) Pylon mounted, left and right hand Thrust reverser halves (’C’ ducts)

For Training Purposes Only

The left fan side is the so called dry side ( EEC, power control unit...) of the engine. The right fan side is the so called wet side ( Oil and Fuel components ) of the engine.

FRA US-T TH NOV 99

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A 330 RR Trent 700

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Lufthansa Technical Training

ENGINE GENERAL

Figure 6 FRA US-T TH NOV 99

Engine Major Units Page: 13

A 330 RR Trent 700

72-00 AIR INTAKE COWL The air intake cowl provides a smooth airstream into the engine. On the upper side is an access panel for the P2/T2 probe. There are several Pip Pin positions around the nose cowl for installing the engine cover. S The intake cowl uses stage 3 HP air for anti icing purposes S The NACA inlet provides ambient air to enter the fan case for venting S EEC cooling is also supplied by air taken from the nose cowl

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ENGINE GENERAL

FRA US-T TH NOV 99

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A 330 RR Trent 700

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ENGINE GENERAL

Figure 7 FRA US-T TH NOV 99

Engine Nose Cowl Page: 15

A 330 RR Trent 700

72-00 REVERSER COWL ( C-DUCT ) DESCRIPTION The function of the thrust reverser is to supply reverse thrust when the aircraft has made a landing. Reverse thrust helps decrease the speed of the aircraft after landing. This is done by the movement of the pivot doors into the L.P. compressor outlet duct, which seals off the duct and makes the L.P. compressor airflow go outboard and in a forward direction. The equal and opposite reaction to the air moving forward creates the reverse thrust. The thrust reverser is an assembly of primary parts, a left and right C–duct. The C–ducts are connected to the pylon by 5 hinges and are latched together at the bottom by 7 latches. When they are closed they make a cover over the core engine and an exit for the L.P. compressor air. The pivot doors are operated by hydraulic pressure from the engine driven hydraulic pump. There are two actuators installed in the C–ducts they allow the C–ducts to be opened for access to the engine. The actuators are operated by hydraulic pressure. A hand operated pump (ground support equipment) is used to supply pressure. NOTE: The A330 uses aircraft hydraulic systems to operate the thrust reverser pivoting doors. Engine No.1 uses the blue hydraulic system, Engine No.2 uses the yellow hydraulic system of the aircraft.

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FRA US-T TH NOV 99

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ENGINE GENERAL

Figure 8 FRA US-T TH NOV 99

Thrust Reverser Assembly Page: 17

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ENGINE GENERAL

A 330 RR Trent 700

72-00 LEFT AND RIGHT HAND FAN COWL DOORS DESCRIPTION The fan cowl doors are manufactured from a lightweight composite material. Each door is attached to the aircraft pylon by four hinges and are closed around the LP compressor cases.They can be opened during ground maintenance to give access to the components installed on the cases and to enable the thrust reverser halves (’C’ Ducts) to be opened. Each fan cowl door has a number of access doors and outlets as follows: S Left Fan Cowl Door – Starter control valve and thrust reverser ground safety switch access – IDG oil fill, sight glass and reset lever access – IDG oil cooler air outlet

For Training Purposes Only

S Right Fan Cowl Door – Oil fill and sight glass access – Hydraulic filter contamination indicator and master MCD access – Air oil heat exchanger air outlet – Zone 1 airflow outlet – Engine breather outlet mast Each fan cowl door has four hinges, four latches or latch keepers, one deflection restraint and two hold open rods.

FRA US-T TH NOV 99

Page: 18

A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 9 FRA US-T TH NOV 99

Fan Cowl Doors Page: 19

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 FAN COWL DOOR OPENING AND CLOSING DESCRIPTION The fan cowl doors are held together by 4 adjustable latches, and 2 deflection restraints. It is not permitted to open the fan cowl doors when the wind speed is more than 60 mph. The engine operation is also limited to idle power only with opened fan cowl doors. There are two deflection restraints which engage in a hole on the nose cowl. OPENING SEQUENCE S release the deflection restraints on the leading edge of the fan cowl doors S release the four fan cowl latches in the sequence 1, 3, 2, 4 numbered from the front S get access to the hold open rods and attach them on support brackets on the fan case

For Training Purposes Only

CLOSING PROCEDURE S hold the fan cowl doors and disengage the hold open rods to stow them back on the fan cowl door S close the fan cowl doors and engage the four latches in the sequence 1, 3, 2, 4 S engage the deflection restraints

FRA US-T TH NOV 99

Page: 20

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ENGINE GENERAL

A 330 RR Trent 700

72-00

For Training Purposes Only

DEFLECTION RESTRAINTS

Figure 10 FRA US-T TH NOV 99

Fan Cowl Door Latches and Rods Page: 21

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ENGINE GENERAL

A 330 RR Trent 700

72-00 C - DUCT OPENING AND CLOSING DESCRIPTION The C - duct contains all the reverser pivoting doors and actuators and is mounted on the pylon by 5 hinges. The sleeves are hydraulicaly opened by two hydraulic actuators pressurized by a handpump . There are 7 reverser latches which are adjustable to keep both c - ducts together. To keep the c-duct in the open position each c-duct has two hold open rods

S engage the take-up device and pull the c-duct together S engage the pin latches 3 and 4 and remove the take-up deivice and stow it back S engage the hook latches in this sequence numbers 1, 2, 5, 6, 7, S close the pressure relief door and remove the hand pump S close the fan cowl doors

For Training Purposes Only

NOTE: Prior opening and closing a Take - up Device has to be installed to take the load from the latches. The latches 1, 2, 5, 6, 7, are hook latches. The latches 3 and 4 are pin latches and located underneath of a pressure relief door which is also the access panel. OPENING PROCEDURE S open the fan cowl doors S open the thrust reverser lacth access and overpressure relief door S operate the take-up device to take the load off the latches S release the hook latches in this sequence, numbers 7, 6, 5, 2, 1 NOTE: The hook latch 1 is operated by a remote lever S then insert a speed brace into pin latch 3 and 4. Push up and turn to disengage the pin S remove the take-up device S connect the hand pump on the manifold and open the c-duct sleeve until both hold open rods can be engaged CLOSING PROCEDURE S connect the hand pump to the manifold and take the load off the hold open rods S disengage the hold open rods and stow them back on the c-duct S lower down the c-duct

FRA US-T TH NOV 99

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A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 11 FRA US-T TH NOV 99

C - Duct Latches / Hydraulic Manifold / Open Rods Page: 23

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 C-DUCT LATCH NO.1 DESCRIPTION The c-duct latch No. 1 is operated by a remote handle and trigger. To open the remote handle the trigger must be released first. The trigger is spring loaded and must hold the handle in the latched position. The force of the handle can be adjusted by use of a flat blade screwdriver to turn the adjusting nut.

C-DUCT LATCH 3 AND 4

For Training Purposes Only

DESCRIPTION The pin latches 3 and 4 can be operated by use of a 3/8 inch square . To engage the pin into the keeper the direction is clockwise, to disengage the pin counterclockwise is the direction of turn.

FRA US-T TH NOV 99

Page: 24

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00

LATCHES 3 AND 4

For Training Purposes Only

HOUSING ASSEMBLY

LATCH 1

Figure 12 FRA US-T TH NOV 99

C - Duct Latches 1, 3 and 4 Page: 25

Lufthansa Technical Training For Training Purposes Only

ENGINE GENERAL

A 330 RR Trent 700

72-00 ENGINE MOUNTS DECRIPTION The engine is attached to the aircraft pylon with two engine mounts. The front mount is attached at the top of the intermediate case.The rear mount is attached at the top of the exhaust case. The mounts support the weight of the engine and transmit loads to the aircraft structure. Spherical bearings in each mount permit thermal expansion and some movement between the engine and the aircraft pylon, the two mounts are made to fail safe. FRONT MOUNT The engine front mount transmits engine thrust, side and vertical loads to the aircraft pylon. The thrust and side loads are transmitted from the intermediate case through a split spherical bearing – which is mounted on the intermediate case – to the cylindrical trunnion. These loads are now transmitted through the main attachment bracket to the aircraft pylon. The vertical loads are transmitted from the intermediate case through the vertical load links to the vertical load support beam.They are then transmitted through the front horizontal trunnion to the main attachment bracket to the aircraft pylon. The main attachment bracket is in two halves to give more than one route for the thrust and side loads. If there is a failure of a primary component that affects the vertical loading the engine would drop and the fail safe catcher link would contact the rear trunnion (on the main attachment bracket) and support the vertical loads.

FRA US-T TH NOV 99

Page: 26

A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 13 FRA US-T TH NOV 99

Front Mount Page: 27

A 330 RR Trent 700

72-00 REAR ENGINE MOUNT The engine rear mount transmits engine torque loads, vertical loads and side loads to the aircraft pylon. The torque loads and side loads are transmitted from the lug at the top of the exhaust to the fail safe link. They are then transmitted through the left hand inboard fork lugs (when viewed from the rear of the engine) to the intermediate fitting to the aircraft pylon. Vertical loads are transmitted through the engine mount links through the intermediate fitting to the aircraft pylon. If a failure of the fail safe link occurs, the fail safe link pin will transmit torque and side loads. The fail safe pin is in a clearance hole. Therefore the pin will only transmit loads if a failure occurs. If there is a failure of the fail safe link where it attaches to the exhaust case, the torque pads and shoulder pads engage and transmit side and torque loadings. If there is a failure of an engine mount link the fail safe link will transmit side, torque and vertical loads, the remaining engine mount link will transmit vertical and torque loads.

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ENGINE GENERAL

FRA US-T TH NOV 99

Page: 28

A 330 RR Trent 700

72-00

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Lufthansa Technical Training

ENGINE GENERAL

Figure 14 FRA US-T TH NOV 99

Rear Mount Page: 29

A 330 RR Trent 700

72-00 COMMON NOZZLE ASSEMBLY ( CNA ) DESCRIPTION The CNA which is bolted to the engine exhaust case mixes the primary and the secondary airstreams together. There are 4 different CNA’s available from the manufacturer. Each of the four CNA’s has slightly different airstreams through it. Therefore the engine data plate shows a EPR trim value on it which results from those manufacturer tolerances.

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ENGINE GENERAL

FRA US-T TH NOV 99

Page: 30

A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

TRIM PLATE

Figure 15 FRA US-T TH NOV 99

Common Nozzle Assembly Page: 31

A 330 RR Trent 700

72-00 ACCESSIBILITY (LEFT SIDE) DESCRIPTION Detachable or hinged panels are provided in the propulsion system outer surfaces where necessary to allow for access to the following: S Thermal anti–icing air outlet S Interphone socket S Starter control valve and thrust reverser ground safety switch access door S I.D.G. oil fill sight glass and reset lever access door S I.D.G. oil cooler air outlet Thrust reverser pivot doors

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ENGINE GENERAL

FRA US-T TH NOV 99

Page: 32

A 330 RR Trent 700

72-00

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Lufthansa Technical Training

ENGINE GENERAL

Figure 16 FRA US-T TH NOV 99

Left Side Accessibility Page: 33

A 330 RR Trent 700

72-00 ACCESSIBILITY (RIGHT SIDE) DESCRIPTION Detachable or hinged panels are provided in the propulsion system outer surfaces where necessary to allow for access to the following: S P20/T20 probe access panel S Engine breather outlet S Hydraulic filter contamination indicator and master M.C.D. access door S Oil filter and sight glass access door S Engine air oil heat exchanger (A.O.H.E.) air outlet S Zone 1 airflow outlet S Thrust reverser pivot doors

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ENGINE GENERAL

FRA US-T TH NOV 99

Page: 34

A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 17 FRA US-T TH NOV 99

Right Side Accessibility Page: 35

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POWER PLANT ENGINE DRAINS

A 330 RR Trent 700

71-70

71-70

ENGINE DRAINS

INTRODUCTION The powerplant drains system is provided to fulfil the following functions: S To collect fuel which has not been burned because of engine shut down or failure to start. S To remove and discard fuel and/or oil if a leak occurs across an internal seal in certain primary components. This also provides the means of monitoring the condition of these seals. S To remove unwanted liquids which can collect in the pylon, cowls and fairings. DESCRIPTION If component seals are leaking stainless steel tubes convey the fluid to an overboard drain (drains mast). Most of the tubes connect to specified primary components as follows: S Air Oil Heat Exchanger (A.O.H.E.) S Hydraulic Pumps S L.P./H.P. Fuel Pump S Fuel Metering Unit (F.M.U.) S Starter Motor S Integrated Drive Generator (I.D.G.) S (V.I.G.V./V.S.V.) Actuator S Drains Collector Tank S Oil Tank Filler Scupper Other tubes in the drains system remove unwanted fluids from specified areas of the powerplant and they are:– S The Pylon Primary Structure S The Core Engine Fairings S The L.P. Turbine Area A small sump is installed in many of the components drain tubes. The sump will hold some of the fluid if leaks occur from the components. This fluid in the sumps can then be used to identify the defective component.

FRA US-T TH NOV 99

Page: 36

A 330 RR Trent 700

71-70

For Training Purposes Only

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POWER PLANT ENGINE DRAINS

Figure 18 FRA US-T TH NOV 99

Drains System Simplified Page: 37

A 330 RR Trent 700

71-70 OPERATION OF DRAINS TANK DESCRIPTION When the engine is shut down, or after failure to start fuel is drained from the fuel manifold. As fuel flows into the tank air is released through the outlet tube. After a number of failed starts, the tank can become full of drained fuel, this fuel is then discharged through the outlet tube and to the drains mast. During normal operation fuel in the drains tank lifts the float valve and moves it to the open position. During engine starting L.P. fuel flows through the ejector, this will lower the fuel pressure in the ejector to less than that in the tank and the non–return valve opens, fuel is now removed from the tank to be routed to the inlet side of the L.P. pump. When the fuel falls to a certain level the float valve closes this prevents air being introduced into the L.P. fuel supply.

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POWER PLANT ENGINE DRAINS

FRA US-T TH NOV 99

Page: 38

A 330 RR Trent 700

71-70

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POWER PLANT ENGINE DRAINS

Figure 19 FRA US-T TH NOV 99

Drains Tank Page: 39

A 330 RR Trent 700

71-70 COMPONENT LOCATION DRAINS TANK The drains tank is attached to the front face of the external gearbox. DRAINS MAST The drains mast is located on a bracket and fitted to the front face of the external gearbox. Provision is made for it to protrude out of the bottom of the hinge side cowls.

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POWER PLANT ENGINE DRAINS

FRA US-T TH NOV 99

Page: 40

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POWER PLANT ENGINE DRAINS

A 330 RR Trent 700

71-70

For Training Purposes Only

DRAINS TANK

DRAINS MAST

Figure 20 FRA US-T TH NOV 99

Drains Tank / Drains Mast Page: 41

A 330 RR Trent 700

72-00

72-00

ENGINE PRESENTATION

MAIN ROTATING ASSEMBLIES The three rotating assemblies comprise: S Low Pressure (L.P.) compressor (Fan) connected by a shaft to a four stage turbine. S Intermediate pressure (I .P.) compressor connected by a shaft to a single stage turbine. S High Pressure (H.P.) compressor connected by a shaft to a single stage turbine. S Each shaft is supported by roller bearings and ball (location) bearings. S The external gearbox is driven from the H.P. shaft through an internal gearbox and an intermediate (step–aside) gearbox.

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FRA US-T TH NOV 99

Page: 42

A 330 RR Trent 700

72-00

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ENGINE GENERAL

Figure 21 FRA US-T TH NOV 99

Main Rotating Assemblies Page: 43

A 330 RR Trent 700

72-00 ENGINE MAIN BEARING ARRANGEMENT DESCRIPTION The LP and IP rotor assemblies are each supported by three bearings. The HP rotor is supported by two. Two types of bearings are used in this engine: – deep seated ball bearings for shaft location – roller bearings providing shaft radial support whilst allowing axial thermal movement The location bearings for all three spools are positioned in the intercase (module 33 ) All 3 bearings are ball bearings which are used for thrust transmission. They are located in line with the front engine mount. The roller bearings are in the respective bearing housings i.e., the LP compressor and IP compressor shaft roller bearings are in the module 32 front bearing housing, the HP turbine and IP turbine roller bearings are in the module 05 bearings housing and the LP turbine roller bearing is in the module 51 tail bearing housing.

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FRA US-T TH NOV 99

Page: 44

A 330 RR Trent 700

72-00

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Lufthansa Technical Training

ENGINE GENERAL

Figure 22 FRA US-T TH NOV 99

Engine Bearing Arrangement Page: 45

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 MODULE BREAKDOWN OF ENGINE DESCRIPTION The engine modules are corresponding to ATA chapters. S Module 31 LP Fan Shaft and Rotor S Module 32 IP Compressor S Module 33 Intermediate Case S Module 34 LP Compressor Case S Module 41 HP system S Module 51 IP Turbine S Module 52 LP Turbine S Module 61 External Gearbox

For Training Purposes Only

NOTE: For Splitship transportation the module 32 front part has to detached from the LP compressor case.

FRA US-T TH NOV 99

Page: 46

A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 23 FRA US-T TH NOV 99

Trent Modular Breakdown Page: 47

A 330 RR Trent 700

72-00 L.P. COMPRESSOR MODULE DESCRIPTION The L.P. compressor is a one stage rotor with 26 wide chord fan blades which engage in axial ”dovetail” slots. Each blade is held in the disc with two shear keys, radial movement is prevented by a slider assembly. Installed between adjacent blades and held by the front support ring are annulus fillers to give smooth contour to prevent air turbulence. The L.P. shaft is attached to the disc with a curvic coupling. Behind the curvic coupling is a roller bearing and to the rear of the bearing is a machined phonic wheel. This is used with an electrical pick–up to measure L.P. compressor speed. Bolted to the front of the disc is a spinner which gives a smooth contour to the air entering the L.P. compressor. At the rear end of the L.P. compressor shaft there are internal splines which engage the L.P. turbine shaft. A coupling, a splined locking ring and a nut that can be adjusted for setting the turbine blades, in relation to the turbine static assembly, connect the two shafts together. The rear of the L.P. compressor shaft is held in location by a ball bearing which maintains it in the correct axial position.

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FRA US-T TH NOV 99

Page: 48

A 330 RR Trent 700

72-00

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ENGINE GENERAL

Figure 24 FRA US-T TH NOV 99

LP Compressor Module Page: 49

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 FAN BLADE, ANNULUS FILLER, SPINNER, FAIRING AND MAKE–UP PIECE INSPECTION/CHECK DESCRIPTION S Examine blades for cracks, if cracks are found reject blade. S Examine blade tips for blueing or heat discolouration, if found the maintenance manual explains the accept/reject standard.

S Examine the make up piece inner and outer surface for scores, the Maintenance Manual explains the accept/reject standard. S If there are shank nuts with damaged threads reject.

S Examine blade surface for arc burns, if found reject blade. S Examine blade for nicks and bends, if either are found the Maintenance Manual explains the accept/reject standard. In some cases there are ”fly on” limits for a maximum of 125 hours or 25 flight cycles. The Maintenance Manual splits the blade up into zones and in some cases it depends which zone a nick or bend is in as to whether to accept or reject. In the diagram below is a typical example. S Examine the annulus filler, if cracks, bends or distortion are found reject filler. If nicks, scores, dents, loss of surface protection are found the Maintenance Manual explains the accept/reject standard. If the airseals are damaged or missing reject filler. S Examine spinner rubber tip, if it is not there or there is not a good bond between tip and spinner reject spinner.

For Training Purposes Only

S If the filler is not all there but there is a good bond between tip and spinner accept it. S Examine the spinner for grooves, scores, cracks and delamination, see Maintenance Manual for accept/reject standard. S Examine fairing for cracks and bulges, if found reject fairing. S If fairing ”P” seal is damaged, loose, or not there reject fairing. ”Fly on” limits maximum of 125 hours or 25 flight cycles.

FRA US-T TH NOV 99

Page: 50

A 330 RR Trent 700

72-00

SHEAR KEYS

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ENGINE GENERAL

Figure 25 FRA US-T TH NOV 99

Fan Blade, Spinner Fairing, Make-up Piece Page: 51

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 AIR INTAKE FAIRING/SPINNER AND MAKE–UP PIECE REMOVAL/INSTALLATION DESCRIPTION WARNING Make sure CNA rear cover FK22421 is installed. Movement of air through the engine can cause the LP compressor to turn very quickly and cause injury. S Get access to air intake cowl. S Install work mats in air intake. Removal Procedure S Mark a line from the annulus filler surface across the make–up piece, fairing and spinner using the OMat 262 marker. S Remove fairing retaining bolts. S Using extractor HU 29255 and adapter HU 35451 remove the fairing. S Remove the spinner. S Hold ma’ke–up piece and remove bolts. NOTE : The make–up piece weighs 17 lbs (8 kgs).

For Training Purposes Only

Installation Procedure S Position make–up piece on the support ring. Make sure it is correctly aligned with the line on the annulus filler and fit bolts. S Torque load the bolts to 285 lbf/in (3,22 mdaN).

FRA US-T TH NOV 99

Page: 52

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00

For Training Purposes Only

SUPPORT RING ALIGNMENT PIN

Figure 26 FRA US-T TH NOV 99

Support Ring, Slider Assembly and Annulus Fillers Page: 53

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 LP COMPRESSOR (FAN)BLADE REMOVAL / INSTALLATION NOTE: When handling the LP compressor blades use suitable gloves to avoid injury from the edges of the blade. Removal Procedure S Remove air intake fairing/spinner and make–up piece. S Hold support ring and remove bolts. Usea applicable bolts and install them in the extractor bushes.Tighten bolts in increments until support ring is released S Using the OMat 262 marker, mark the blade and the annulus fillers either side of the blade to be removed. Correlate each annulus filler to the disc for refit purpose Pull annulus filler forward to disengage it from the disc and then turn the annulus filler in the direction of its curve to clear the blade and remove it. Make sure the blade to be removed is at the bottom S Use extractor HU 29255 and adapter HU 37954 to remove the slider assembly. S Hold the blade and lift it radially until the two shear keys disengage from their safety slots: then pull the blade forward approximately lin (25mm) and lower the blade back to the bottom of disc groove.

For Training Purposes Only

S Pull the blade slowly forward until the rear shear key engages in the front safety slot. S Hold the blade and lift it radially until the rear shear key disengages from the front safety slot: then pull the blade forward approximately lin (25mm) and lower the blade back to the bottom of the disc groove. S Pull the blade forward and remove it. S Make a note of the moment weight of the removed blade.

FRA US-T TH NOV 99

Page: 54

A 330 RR Trent 700

72-00

For Training Purposes Only

Lufthansa Technical Training

ENGINE GENERAL

Figure 27 FRA US-T TH NOV 99

Fan Blade Replacement Sequence Page: 55

Lufthansa Technical Training

ENGINE GENERAL

72-00 I.P. COMPRESSOR MODULE DESCRIPTION The I.P. compressor module is an eight stage axial assembly consisting four main sections: S Front bearing housing S Variable stator vane case S The I.P. compressor case S The I.P. compressor rotor FRONT BEARING HOUSING The front bearing housing, most of which is made of titanium, includes a hub, which locates the L.P. and I.P. compressor bearings and an oil sump, also the L.P. and I.P. shaft speed probes. Connected to the hub are 58 titanium engine section stator vanes. The vanes are welded together as one unit and there are lugs on the outer ring. These lugs align with the titanium outlet guide vanes at the torsion ring to make the FBH/OGV joint. The mating parts of this joint are aligned with dowels and are connected with bolts that pass through the dowels. This FBH/OGV joint holds the L.P. compressor case to the core engine. Two electrical cables pass internally through two of the E.S.S. vanes to transmit signals from the shaft speed probes. Six more vanes contain tubes to supply oil to and from the roller bearings. Behind the E.S.S. vanes are 58 variable inlet guide vanes.

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A 330 RR Trent 700

I.P. COMPRESSOR ROTOR The I.P. compressor rotor is an assembly of eight rotor discs made of titanium and welded together to form a drum, in between the discs there are spacers which have interstage seal fins. The discs at stages 1 to 6 have axial dovetail slots into which the rotor blades are installed. Retaining plates and lock plates keep the blades in position. At stages 7 and 8 the blades are installed in circumferential dovetail slots. These blades are locked in position with nut and screw lock assemblies. The I.P. front stubshaft is attached to the stage 1 disc with bolts, the forward end of the stubshaft has a phonic wheel. The rear stubshaft is attached to the stage 6 disc with a curvic coupling.Splines in the stubshaft engage with splines on the I.P. turbine shaft.

VARIABLE STATOR VANE CASE The I.P. compressor case can be divided into two cylindrical parts. The front part which is made of titanium contains the first two stages of the compressor. This is the V.S.V. case which can be divided into two semi–circular half cases. Stage 1 and 2 stator vanes, which are variable, are installed in these half casings and are connected to the V.I.G.V. mechanism. I.P. COMPRESSOR CASE The I.P. compressor case is flanged and bolted to the rear of the V.S.V. case and is made of steel and contains stages 3 to 8 of the compressor. It also can be divided into two semi–circular half cases. The stage 8 stator vanes, also known as the I.P. compressor outlet guide vanes are contained in a case which is flanged and bolted to the rear of the I.P. compressor case. FRA US-T TH NOV 99

Page: 56

A 330 RR Trent 700

72-00

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Lufthansa Technical Training

ENGINE GENERAL

Figure 28 FRA US-T TH NOV 99

IP Compressor Module Page: 57

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 INTERMEDIATE CASE MODULE DESCRIPTION The intermediate case is made from two titanium cylindrical casings which are welded together. In the rear case there are ten vanes which support the internal gearbox housing. These aerofoil shaped vanes are hollow and some contain tubes which supply oil to and from the internal gearbox. Other vanes will supply I.P. compressor air for cooling and sealing bearingchambers. The internal gearbox contains the three location bearings for the three compressor shafts and provides the drive for the H.S. external gearbox drive housing. The front part of the casing has a strengthened top section to include the front engine mount. Above and below the centre line are the positions for the installation of the ’A’ frame struts. These struts connect to the inside of the L.P. case. The front part of the intermediate case is installed around the rear part of the I.P. compressor case. The flange connecting to a flange at the rear of the V.S.V. case.

For Training Purposes Only

The rear part of the intermediate case is installed around the H.P. compressor case. The flange connecting to a flange of the combustion chamber outer case.

FRA US-T TH NOV 99

Page: 58

A 330 RR Trent 700

72-00

A-FRAME ATTACHMENT

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Figure 29 FRA US-T TH NOV 99

Intermediate Case Module Page: 59

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 COMPRESSOR FAIRINGS / ’A’ FRAME STRUTS COMPRESSOR FAIRINGS To ensure a smooth airflow over the parts of the gas generator not covered by the inner fixed structure ( C-ducts ), 6 removable fairings are fitted around the front part o fthe IP compressor case. Each fairing is made of carbon fibre with a honeycomb core. the inner surface has a fireproof protection. At the front edge bolts attach the fairings to the LP compressor OGV torsion ring with floating anchor nuts. The rear edge is attached to mounting brackets on the rear support diaphragm with bolts and floating anchor nuts.

For Training Purposes Only

’A’ FRAME STRUTS The ’A’ Frames make a connection between the intermediate case module and the LP case inner side. There are two attachment positions on each side of the intermediate case module just above and below the centre line. On the inner side of the LP case there is one attachment position on each side ( 03:00 and 09:00 position ). The struts are made of titanium and strengthen the connection between the LP case and the intermediate case compressor.

FRA US-T TH NOV 99

Page: 60

A 330 RR Trent 700

72-00

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Figure 30 FRA US-T TH NOV 99

Compressor Fairings / A Frame Struts Page: 61

Lufthansa Technical Training

ENGINE GENERAL

A 330 RR Trent 700

72-00 H.P. SYSTEM MODULE DESCRIPTION The system comprises: S H.P. compressor S Combustion chamber and outer case S H.P. turbine

a stubshaft. On the front face of the disc there are two sets of seal fins which control the flow of cooling air. The disc has fir tree roots into which fit 92 turbine blades.

For Training Purposes Only

H.P. COMPRESSOR The H.P. compressor rotor is a six stage assembly of titanium discs welded together to form one drum. The first stage blades are installed in axial dovetail slots and are locked with retaining plates and lock plates. Stages 2 to 6 are installed in circumferential dovetail slots and locked with nuts and screws. The rotor blades 1 to 3 are made of titanium and the others of a heat resistant alloy. Welded to the rear of the stage 6 disc is a titanium cone which tapers rearwards. At the rear of this cone is a mini disc to which the H.P. turbine is connected. The H.P. compressor case is an assembly of six flanged, cylindrical casings, bolted together. The flanged joints are also the location for the rotor path abradable linings. There are slots in this assembly for the installation of the stator vanes.The stage 6 stator vanes are also the H.P. compressor outlet guide vanes (O.G.V’s).These are installed at the entrance of the combustion chamber inner case.

COMBUSTION CHAMBER AND OUTER CASE The outer case is flanged and bolted to the rear of the intermediate case and to the front of the I.P. turbine module. There are 24 openings through which the fuel spray nozzles are installed. There are also two igniter plugs installed through bosses in the combustion outer case. The combustion chamber is fully annular and consists of a liner which is located inside the combustion chamber inner case. At the front of the inner case are the H.P. compressor O.G.V’s and at its rear are 40 H.P. turbine nozzle guide vanes (N.G.V’s). H.P. TURBINE The H.P. turbine is a single stage disc connected to a mini disc to the rear of the H.P. compressor drum. On the rear of the disc is a flange which attaches to FRA US-T TH NOV 99

Page: 62

A 330 RR Trent 700

72-00

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ENGINE GENERAL

Figure 31 FRA US-T TH NOV 99

HP System Module Page: 63

A 330 RR Trent 700

72-00 I.P. TURBINE MODULE DESCRIPTION The I.P. turbine is a single stage disc which has a flange on its rear face. Attached to this with bolts are the flanges of two shafts which go forward through the centre of the disc. One of these is a stubshaft which is supported by the inner race of a roller bearing to hold the I.P. turbine in position. The other shaft passes through the stubshaft and the H.P. system and connects to the I.P. compressor shaft with splines. The disc has fir tree roots into which fit 126 turbine blades. In front of the I.P. turbine blades are 26 hollow N.G.V’s. In 13 of these N.G.V’s is a strut which is attached to the turbine case by a bolt. The inner end of each strut is connected to the structure which holds the H.P./I.P. bearing support assembly. Through some of the other N.G.V’s are tubes to supply oil to and from the bearings and I.P. 8 cooling air to cool the housing. The I.P. turbine case is flanged and bolted between the combustion chamber outer case at the front, and the L.P. turbine case at the rear. Adjacent to the rear flange is a turbine case cooling (T.C.C.) air manifold and location bosses for eleven thermocouples. To the rear of the turbine blades are the L.P. 1 N.G.V’s.

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11 EGT THERMOCOUPLES

Figure 32 FRA US-T TH NOV 99

IP Turbine Case Page: 65

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A 330 RR Trent 700

72-00 L.P. TURBINE MODULE DESCRIPTION The L.P. turbine has four discs which are bolted together to form a drum. The discs have fir tree roots into which fit the turbine blades. In front of each stage of turbine blades there is a stage of N.G.V’s. The firststage of N.G.V’s, which are hollow, are installed as 3 vane sets in the outlet from the I.P. turbine case. One vane in each set of eleven sets contains a thermocouple and another set includes an overheat detector. Stages 2, 3 and 4 N.G.V’s which are solid are installed in the L.P. turbine case. At the inner ends of the N.G.V’s are honeycomb liners which touch the fins of the interstage seals between the rotor discs. The stage 3 turbine disc has a flange on the front which is attached to the turbine shaft with a curvic coupling. This shaft goes forward through the centre of the I.P. shaft and connects with the L.P. compressor shaft with splines. Also connected to the rear of the stage 3 disc flange is a stubshaft. This is connected to the inner race of the L.P. roller bearing to hold the L.P. turbine in position at the rear.

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The L.P. turbine case is a one piece cylinder flanged and bolted between the I.P. turbine case at the front, and the tail bearing housing at the rear. Around the case is a cooling duct through which cooling air flows. On the inner surface between the N.G.V. locations there are seal segments which touch the turbine blade shrouds. The tail bearing housing support structure includes a hub which is held concentric in an outer case by twelve radial hollow vanes. Some of the vanes contain tubes which supply oil to and from the bearing housing. There is also a supply of I.P. 8 air to cool and seal the bearings. Five of the vanes have pressure inlets in the leading edge to measure L.P. turbine outlet pressure ( P50 ). These pressure values are used as part of the engine pressure ratio ( EPR ) system. The front flange of the case is attached with bolts to the rear flange of the L.P., turbine case, and at the rear flange to the common nozzle assembly (C.N.A.). Around the case are two flanges to increase the strength. Attached to these flanges, at the top, is the rear mount.

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Figure 33 FRA US-T TH NOV 99

LP Turbine Module Page: 67

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A 330 RR Trent 700

72-00 SPRING LOADED L.P. TURBINE BEARING DESCRIPTION It is possible when the engine is running within a certain power range that the L.P. location bearing may suffer damage from skidding due to load cross over. Load cross over is when the L.P. location bearing is not axially loaded, either forwards or rearwards. It is known that ball bearing skidding can be eliminated if the bearing is always loaded axially.

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On the Trent 700 engine the L.P. turbine rear stubshaft has, in addition to a roller bearing, a spring loaded ball bearing. The bearing is part of a spring pack assembly housed within the tail bearing housing.This spring pack loads the L.P. shaft in a rearwards direction, eliminating cross over under all conditions.

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Figure 34 FRA US-T TH NOV 99

Spring Loaded LP Turbine Bearing Page: 69

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A 330 RR Trent 700

72-00 H S GEARBOX MODULE DESCRIPTION The High Speed (H.S.) external gearbox is a one piece aluminium gearcase and is mounted on the lower part of the L.P. compressor case. The H.S. 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 air starter motor to the engine. The gearbox also provides a means of hand turning H.P. rotor system for maintenance purposes. COMPONENTS MOUNTED ON THE FRONT FACE S Air Starter Motor S No2 Hydraulic Pump S Centrifugal Breather S Dedicated Alternator

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COMPONENTS MOUNTED ON THE REAR FACE S Nol Hydraulic Pump S Integrated Drive Generator (I.D.G.) S Oil Pumps S L.P./H.P. Fuel Pumps

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Figure 35 FRA US-T TH NOV 99

External Gearbox Module Page: 71

A 330 RR Trent 700

72-00 L.P. COMPRESSOR CASE MODULE DESCRIPTION The L.P. compressor case is made up of two aluminium cases flanged and bolted together. The front case contains the L.P. compressor and the rear case includes the Outlet Guide Vanes (O.G.V.). The outside construction of the front case is made up of 72 axial ribs equally spaced between the front and rear flange, between these ribs are diagonal ribs that form a triangle, this type of construction is known as isogrid. Kevlar material which has great strength is wrapped around the front casing. This is to contain any failed blade released from the disc. Opposite the blade track there is an attrition lining and at each side of this there are acoustic linings. The rear case includes an O.G.V. outer ring to which are attached, with bolts, 58 hollow titanium O.G.V’s. The inner ends of the vanes are attached to a torsion ring. The titanium blades and the two outer titanium supports (”A” frames) connect the case to the core engine.

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A 330 RR Trent 700

72-00

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Figure 36 FRA US-T TH NOV 99

LP Compressor Case Module Page: 73

A 330 RR Trent 700

72-00

72-00

ENGINE BORESCOPING

BORESCOPE ACCESS PORTS DESCRIPTION IP COMPRESSOR S There are 4 IP compressor access ports HP COMPRESSOR S There are 5 HP compressor access ports COMBUSTION LINER S There are 8 combustion liner borescope ports behind the fuel nozzles HP TURBINE S There is 1 borescope access port for the HP turbine LP TURBINE S There are 4 LP turbine access ports, one can also be used for HP turbine

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Figure 37 FRA US-T TH NOV 99

Borescope Access Ports Page: 75

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A 330 RR Trent 700

72-00 IP HAND TURNING TOOL TURNING THE INTERMEDIATE PRESSURE (IP) SYSTEM The IP system is turned using tool (HU38122). Note:The same warnings and cautions apply as for turning the LP system Before the IP system turning tool can be used, it is necessary to open the VIGV’s as follows: S Open the fan cowl doors S Open thrust reverser ’C’ ducts S Drain the HP fuel system into a clean container S Get access to the VIGV actuators by removing the gas generator fairings S Attach the correct spanner to the spanner flats on the VIGV bellcrank and pull the actuator rams to the retracted position (VIGV’s open) S Remove spanner and replace gas generator fairings NOTE : More fuel may drain as the actuator rams are retracted Turning the IP system is as follows: S Before installing the turning tool turn the clamp using the lever through 180 and lock the thumb nut S Put the turning tool (HU38122) through the LP compressor blades, inlet guide vanes and variable inlet guide vanes at approximately top dead centre S Position the turning tool on the forward edge of the inlet guide vanes (see diagram) S Loosen thumb nut and turn the clamp back through 180_ S Ensure clamp is engaged on the trailing edge of the inlet guide vane and tighten thumb nut S Turn hand knob clockwise to turn the IP system S On completion loosen thumb nut and turn clamp through 180 to disengage clamp from trailing edge of inlet guide vane and remove turning tool S Carry out fuel leak test lAW the MM

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Figure 38 FRA US-T TH NOV 99

IP Hand Turning Tool Page: 77

A 330 RR Trent 700

72-00 IP BORESCOPE PLUGS There are 4 borescope plugs for the IP compressor these are: S IP1S S IP3S S IP5S S IP7S

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Figure 39 FRA US-T TH NOV 99

IP Borescope Plugs Page: 79

A 330 RR Trent 700

72-00 VIGV ACTUATOR The VIGV actuator on the left hand side needs to be opened for IP borescoping. The bellcrank has a opening lug to connect an opening tool. The movement of the actuator opens the VIGV’s and the VSV’s .

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72-00

For Training Purposes Only

PULL HERE TO OPEN

Figure 40 FRA US-T TH NOV 99

VIGV Actuator Page: 81

A 330 RR Trent 700

72-00 TURNING THE HIGH PRESSURE (H.P.) SYSTEM As can be seen from the diagram below, hand turning the H.P. system is done using a special tool installed through the breather housing. This position can be used to turn the H.P. system clockwise or counter–clockwise. The procedure is as follows: S Open the fan cowl doors. S Remove the two bolts and remove breather cover. S Remove and discard seal. S Carefully install H.P. system turning tool E2J52189. S Using breather cover bolts and washers secure turning tool to breather housing. S NOTE: The maximum torque to be applied to the tool is 70 lbf/ft (9,49mdaN). S When hand turning is completed remove the two bolts and washers and carefully remove tool. S Install new sealing ring to breather cover and fit breather cover to breather housing. S Torque load the bolts to 100 lbf/in (1,15m,daN).

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Figure 41 FRA US-T TH NOV 99

Hand Turning of the HP System Page: 83

A 330 RR Trent 700

72-00 HP BORESCOPING CONTINUED The following grafic shows the borescope plugs used on the HP compressor.

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Figure 42 FRA US-T TH NOV 99

HP Borescope Plugs Page: 85

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A 330 RR Trent 700

72-00 COMBUSTION LINER BORO PLUGS There are 8 borescope plugs for the combustion liner and the Nozzle guide vanes. They are located just behind the fuel nozzles. NOTE: 3 of the 8 plugs are using the T30 thermocouples for their installation

HP-TURBINE BORO PLUGS

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There is only one borescope plug for the high pressure turbine. The location is about 05:00 position.

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Figure 43 FRA US-T TH NOV 99

Combustion Liner / HPT Boro Plugs Page: 87

A 330 RR Trent 700

72-00 HP AND IP TURBINE BORO PLUG There is only one HP / IP turbine borescope plug which has a cover plate on it. The location is about 02:00 position on the HP/IP turbine case.

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A 330 RR Trent 700

72-00

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Figure 44 FRA US-T TH NOV 99

HP / IP Turbine Boro Plugs Page: 89

A 330 RR Trent 700

72-00 LP TURBINE BORO PLUG There are 3 LP turbine borescope plugs in order to check the LP turbine. The location is about 05:00 position.

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Figure 45 FRA US-T TH NOV 99

LP Turbine Boro Plugs Page: 91

A 330 RR Trent 700

77-00

77-00

ENGINE INDICATION PRESENTATION

ENGINE/WARNING DISPLAY There are six identical full colour D.U’s located in the flight deck as shown below. Engine parameters are displayed on the two ECAM (Electronic Centralised AircraftMonitoring) display units. The engine and warning display is normally on the upper ECAM display unit which isdivided into an upper and lower area. The upper area displays the following primary engine parameters: S Engine Pressure Ratio (E.P.R.) S Exhaust Gas Temperature (E.G.T.) S N1 Speed S N3 Speed S Fuel Flow Per Engine S Condition of Ignition System S Warning (Cautions and Memo Messages)

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Figure 46 FRA US-T TH NOV 99

Engine Primary Display Page: 93

A 330 RR Trent 700

77-00 SYSTEM DISPLAY The following secondary engine parameters are displayed on the lower ECAM display unit: S N2 Speed S Fuel Used Per Engine S Oil Pressure, Temperature and Quantity S Vibration Level N1, N2, N3 S Nacelle Temperature S Starting Information The lower D.U. also provides system pages, (aircraft and engine system synoptic diagrams and data) text pages (aircraft status and maintenance messages). The engine system parameters will be displayed when the engine system page is called up either automatically or manually.

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Figure 47 FRA US-T TH NOV 99

Engine System Display Page: 95

A 330 RR Trent 700

77-00 PRESSURE AND TEMPERATURE STATIONS The diagram shows the internal system of numbering of pressures and temperatures throughout the engine. Station numbers are assigned to identify specific positions along the aerodynamic flowpath of an engine. A station is a position at the engine, where thermodynamically changes (Pressure, temperature or airspeed) starts or ends.

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Figure 48 FRA US-T TH NOV 99

Pressure and Temperature Stations Page: 97

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A 330 RR Trent 700

77-00

SHAFT SPEED INTRODUCTION There are three primary compressor shafts in the engine. These are the Low Pressure (L.P.) compressor shaft, the Intermediate Pressure (I.P.) compressor shaft and the High Pressure (H.P.) compressor shaft. The speeds at which the shafts turn are measured independently and shown as a percentage equivalent (N1 and N3 rotor speeds) on the upper ECAM display unit, (N2 rotor speed) on the lower ECAM display unit.

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DESCRIPTION The system uses speed probes together with phonic wheels to measure N1 and N2 shaft speeds. The dedicated alternator supplies the speed of the H.P. shaft N3. The outputs from the L.P. and I.P. speed probes are sent to the Overspeed Protection Unit (O.P.U.). The O.P.U. uses two of these signals from each shaft and transmits them to the E.E.C. The dedicated alternator supplies the speed of the H.P. shaft to the E.E.C. via the Power Control Unit (PCU).

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Figure 49 FRA US-T TH NOV 99

Shaft Speed Components Page: 99

A 330 RR Trent 700

77-00 COMPONENT LOCATION The following components are fitted in the system: S The L.P. shaft phonic wheel S The I.P. shaft phonic wheel S Three L.P. speed probes S Three I.P. speed probes S The dedicated alternator The L.P. shaft phonic wheel is at the rear of the L.P. compressor from the roller–bearing inner–race assembly. The I.P. shaft phonic wheel is on the I.P. compressor front stub shaft. The two phonic wheels each have 60 teeth. There are three speed probes for each phonic wheel. Each speed probe has a magnet and a coil which when operational produces a magnetic field. During operation of the engine the shaft and hence the phonic wheels turn. As the teeth of the phonic wheels pass through the magnetic field of the speed probe, it causes an electrical pulse in that probe. The frequency of the pulses is in proportion to the speed of the shaft. The speed probes and the trim balance probe are attached to the support block inside the front bearing housing. Conduits carry wires from the probe assemblies through the front bearing support case and through the fixed I.P. inlet guide vanes to a junction block. The dedicated alternator is installed on the front face of the external gearbox.

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Figure 50 FRA US-T TH NOV 99

Shaft Speed Components Page: 101

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A 330 RR Trent 700

77-00 N1 INDICATION GENERAL The N1 indication is displayed on the EWD in percent in analog and digital forms. The N1 needle and the N1 digital indications are: S in green colour in normal operation S in red steady colour if the N1 actual exceeds the N1 RED LINE value. The master warning light comes on together with the CRC and the ECAM message ENG N1 OVERLIMIT is displayed to the flight crew. Degraded data are displayed in case of failure of the direct N1 measurement system, in this case, the EEC computes a theoretical value through the other engine parameters. The last digit is then displayed in amber dashes across. N1 limit is displayed when EPR indication is lost and the engine has to be operated in the N1 reverionary mode ( N1MODE ). The EPR limits are then replaced by the specific N1 limit indications in the white box and the N1 MODE shows that the engines have to be controlled in the N1 mode. NOTE The N1 exceedance must be erased by maintenace action on the MCDU or by next engine start.

NOTE The N2 exceedance must be erased by maintenance action on the MCDU or by next engine start.

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N2 INDICATION GENERAL The N2 indication is displayed in percent on the SD in digital form only. Normally the N2 actual is displayed in green colour. When N2 actual exceeds N2 red limit value ( 103,3% ): S the indication changes from geen to red S the red cross appears next to the digital indication to show the flight crew and the maintenance that N2 exceedance occured The master warning light comes on together with the CRC and the ECAM message ENG N2 OVERLIMIT. Degraded data is displayed with two amber dashes acrossthe last digit.

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77-00

N1 INDICATION

N2 INDICATION

For Training Purposes Only

N2 ACTUAL ( NORMALLY GREEN )

DEGRADED DATA ( AMBER ) N1 EXCEEDANCE ( RED ) N1 ACTUAL ( NORMALLY GREEN )

N2 EXCEEDANCE

N1 RED LINE ( RED ) DEGRADED DATA ( AMBER )

Figure 51 FRA US-T TH NOV 99

N1 Indication / N2 Indication Page: 103

A 330 RR Trent 700

77-00 N3 INDICATION The N3 indication is displayed in percent on the EWD in digital form only. Normally the N3 actual digital display is in green colour. When the N3 actual exceeds N3 red limit value ( 100% ): S the indication changes from green to red S the the red cross appears next to the digital indication to show the flight crew and the maintenance that an exceedance has occured The master warning light comes on together with the CRC and the ECAM message ENG N3 OVERLIMIT. Degraded data is displayed with two amber dashes across the last digit. NOTE: The exceedance must be erased by the maintenance through the MCDU or next engine start.

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N3 INDICATION

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N3 ACTUAL ( GREEN )

DEGRADED DATA ( AMBER )

N3 EXCEEDANCE ( RED )

Figure 52 FRA US-T TH NOV 99

N3 Indication Page: 105

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A 330 RR Trent 700

77-00

POWER MEASUREMENT INTRODUCTION Actual thrust can only be measured in a test cell, therefore when the engine is fitted to the aircraft some other method of measuring power that can equate to thrust must be used. The power measuring device used on the Trent is called the engine pressure ratio (E.P.R.) system. The air intake pressure P20 and the turbine outlet pressure P50 are compared and expressed as a ratio therefore: EPR = P50 : P20 E.P.R. is related to thrust and as a result can be used as a parameter for its control. The E.P.R. indication is displayed in the flight deck. The P20 accumulator collects the air intake flow from the P20T20 probe and smoothly supplies the air to the EEC. The accumulator prevents sudden changes of the P20 which can have an unsatisfactory effect on engine performance. The accumulator is downstream of the P20T20 probe. You can find the accumulator in a position lower than the electronic–unit protection–box.

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DESCRIPTION The E.P.R. indicating system consists of three primary components: S The L.P. turbine bearing housing support vanes pressure inlets S The E.E.C. S The P20/T20 probe OPERATION The E.E.C. collects values of P50 from the L.P. turbine bearing support vanes and P20 from the P20/T20 probe. The E.E.C. compares these values to form the actual E.P.R. and transmits the E.P.R. value to the electron centralised aircraft monitoring system (E.C.A.M.) for display on the upper screen engine/warning display (E/WD). Each of the two E.E.C. channels perform this operation independently.

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Figure 53 FRA US-T TH NOV 99

EPR System P20 / P50 Probes Page: 107

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A 330 RR Trent 700

77-00 EPR INDICATION EPR - ENGINE PRESSURE RATIO EPR the primary parameter used to set the thrust, is the ratio of exhaust gas total pressure to front compressor inlet pressure ( P50/P20 ). The EPR system has – a combination P20/T20 probe to measure the front compressor inlet pressure and temperature – five probes used to measure exhaust gas pressure Front compressor inlet pressure and exhaust gas pressure are routed to the EEC where transducers change the pneumatic pressures to electrical signals. The signals are used to calculate the actual engine pessure ratio which is transmitted to the flight deck for display.

1

EPR ACTUAl digital

5

EPR MAX – The EPR max is displayed by means of a thick amber mark across the EPR scale – corresponds to the EPR limit of the TOGA mode – not displayed in reverse mode or when the engine is off

6

REVERSE INDICATION – green indication when reverser is deployed – amber indication when at least one sleeve is unstowed – reverser failed stowed in flight amber indication flashes for 9 sec

– green steady in normal operation – replaced by amber XX in the back-up mode

2

EPR ACTUAL analog – green steady in normal operation – disappears in back-up mode

For Training Purposes Only

3

EPR TREND – four green arcs indicate the EPR trend – only displayed if auto thrust is active – not displayed when reverse is active or engine is off

4

EPR THROTTLE – EPR throttle sysmbol ( small white circle ) – not displayed in reverse mode or when the engine is off ( FADEC deenergized )

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4

EPR THROTTLE

3 For Training Purposes Only

2

5

EPR MAX

EPR TREND

ACTUAL ANALOG EPR

6 1

ACTUAL DIGITAL EPR

1.282

Figure 54 FRA US-T TH NOV 99

REVERSE INDICATION

1.282

EPR Indication Page: 109

A 330 RR Trent 700

77-00 POWER MEASUREMENT E.P.R. TRIMMING During pass off testing of an engine it is run at a designed thrust. However the E.P.R. indication of different engines being run at the same thrust will be different.This is due to the manufacturing tolerances of the tail bearing housing. Therefore to enable the E.P.R. indication for all engines to be the same, for a given thrust, E.P.R. trimming is required. This is done by entering any one of 32 different trims into the E.E.C. by means of a data entry plug. Each entry is allocated a trim code which is etched on to a data plate see FIG 1. This data plate is located on the right hand side of the tail bearing housing. This code is also etched onto the main data slip plate (Fig 2). The areas of the C.N.A. can also affect the E.P.R. indication for a given thrust. The C.N.A. has been grouped into 4 bands depending on area and each band is given a code. This code is etched onto the main data slip plate (Fig 2) and onto the CNA. The engine name plate and main data slip plate (Fig 2) are attached to the left side of the rear fancase. During engine repair or overhaul of the tail bearing housing or a complete C.N.A. or core engine is replaced, then the main data slip plate must be checked to ensure that the E.P.R. trim code is consistent with the ones for the C.N.A. and tail bearing housing. If it is incorrect it must be changed for one that has the correct code. A check must also be carried out to ensure that the data entry details align with this data.

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FIG. 1

FIG. 2

Figure 55 FRA US-T TH NOV 99

EPR Trimming Page: 111

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A 330 RR Trent 700

77-00 EXHAUST GAS TEMPERATURE INTRODUCTION Analog Electrical signals are sent from the 11 dual thermocouples in the E.G.T.indicating system to the EEC. The EEC transmits the signal to the E.C.A.M. to be displayed on the upper screen E/WD. The voltage from each thermocouple is proportional to the temperature at its thermocouple. ECAM INDICATION

1

EGT actual – digital and analog indication in green color

2

EGT amber line – is a variable value – 700_ C for eng start – 920_ C for take-off

3

EGT red line exceedance – appears when the EGT red line has exceeded – stays at the maximum value which has been reached – reset after new start on ground, or through the EEC via the MCDU

For Training Purposes Only

4

EGT red line – is represented by an arc shaped red ribbon located at the end of the scale beginning at the red limit value ( 920_ C )

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1

Figure 56 FRA US-T TH NOV 99

EGT Actual

720

2

Amber Line

3

Red Line Exceedance

4

Red Line Band

Simplified EGT Diagram / ECAM Indication Page: 113

A 330 RR Trent 700

77-00 DESCRIPTION The E.G.T. indicating system uses 11 thermocouple assemblies to measure E.G.T. The thermocouples are installed in the L.P. turbine stage 1 nozzle guide vanes. A thermocouple assembly consists of two thermocouples each contained in a tube. Each tube has an inlet and an outlet hole. These holes align with inlet and outlet holes in the L.P.1 N.G.V.. The two tubes are brazed together to make a rounded thermocouple probe. Each thermocouple probe is installed in the L.P.1 N.G.V. An end plug at the end of each probe helps it stay rigid and also gives it some wear protection. Each thermocouple probe passes through a transfer tube assembly. The tube assembly isolates the case from the hot gas in the thermocouple N.G.V. The thermocouple assemblies are connected in parallel through two wires. One wire is made of Nickel Chromium (Chromel) and the other is made of Nickel Aluminium (Alumel). A wire harness connects the, thermocouple set to the E.E.C.

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A 330 RR Trent 700

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Figure 57 FRA US-T TH NOV 99

EGT Thermocouples Page: 115

A 330 RR Trent 700

77-00 EGT TRIMMMING The following grafic shows the engine data plate which is located on the left hand side of the engine. The EGT trim factors the actual engine EGT to a lower value for display in the cockpit. The EGT trim is calculated from data obtained during the engine manufacturers test to align approved EGT levels with the cockpit indications. The EGT trim is calculated at three temperatures equivalent to the aircraft cockpit ECAM limits for crew warnings S max continuous 850° S max take-off 900° S max overtemperature 920°

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Figure 58 FRA US-T TH NOV 99

EGT Trimming Page: 117

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A 330 RR Trent 700

77-00

VIBRATION MONITORING INTRODUCTION The purpose of the engine vibration monitoring system is to provide the flight crew with continuous indication of the state of balance of the engine main rotating assemblies during steady state running conditions.This information can alert operators to existing or impending engine problems and assist in planning module renewal, with minimum disruption to aircraft operation. The transducer converts vibration signals into electrical signals which are sent to the dual junction box. The signal is then amplified by the remote charge converter to be processed by the engine interface and vibration monitoring unit (E.I.V.M.U.) and is displayed on the lower screen of the E.C.A.M.

3.3 UNITS 2.6 UNITS 4.0 UNITS

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VIBRATION GUIDELINES N1 ADVISORY N2 ADVISORY N3 ADVISORY

FRA US-T TH NOV 99

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Figure 59 FRA US-T TH NOV 99

Simplified Vibration Diagram Page: 119

A 330 RR Trent 700

77-00 ENGINE INTERFACE AND VIBRATION MONITORING SYSTEM The system consists in a computer EIVMU and in a separate Remote Charge Converter, RCC. The EIVMU is installed in the avionic compartment, interfaces with the aircraft computers and controls, and with the associated propulsion system to perform the following functions: – Data concentration from the cockpit panels and various aircraft computers to the associated engine control – Engine to Engine segregation – Airframe electrical supplies to engine control – Internal processing of some engine related status signals needed by airframe systems – internal processing of some airframe status signals needed by the related engine control system – Engine vibrations signals processing and monitoring

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Figure 60 FRA US-T TH NOV 99

EIVM System Architecture Page: 121

A 330 RR Trent 700

77-00 VIBRATION MONITORING DESCRIPTION The engine vibration monitoring system contains three primary components: S The vibration transducer S The dual junction box S The remote charge converter The vibration transducer is a dual output accelerometer and is attached to the intermediate case. It contains two piezo–electric crystal stack elements. Each stack element is subjected to a mechanical load from an electrically insulated seismic mass. During operation of the engine, vibration causes the seismic mass to apply pressure to the crystal stack elements to generate electrical signals which are in proportion to the engine vibration frequency. These signals are then sent to the dual junction box and on to the remotecharge converter this electronic unit amplifies the electrical signal to indicate the vibration level on the lower screen of the E.C.A.M.

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A 330 RR Trent 700

77-00

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Figure 61 FRA US-T TH NOV 99

Vibration Transducer Page: 123

A 330 RR Trent 700

77-00 REMOTE CHARGE CONVERTER The RCC is located under the left hand fan cowl close to the starter air valve. This RCC is an amplifier interfacing with the EIVMU and the engine mounted dual accelerometer.

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77-00

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Figure 62 FRA US-T TH NOV 99

Remote Charge Converter Page: 125

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A 330 RR Trent 700

77-00

EIVMU GENERAL 1 LAST LEG REPORT The purpose of this item is to present in plain language the list of the class 1 and 2 internal and external faults detected by the EIVMU during the last flight.

7 TEST This item enables the initiation of the EIVMU test from the MCDU.

8 GROUND REPORT 2 PREVIOUS LEGS REPORT The purpose of this item is to present in plain language the list of the class 1 and 2 internal and external faults detected by the EIVMU during the last 63 flights ( exluding the last flight ).

The purpose of this item is to present in plain language the list of the internal faults detected by the EIVMU and which occured after landing. This item enables access to the TROUBLE SHOOTING DATA fault by fault.

3 LRU IDENTIFICATION The purpose of this item is to present the hardware and software status of the EIVMU ( part number and serial number ).

4 GROUND SCANNING The purpose of this item is to analyze on the ground only the faults that occured during the lasat flight. All the faults detected while using this function ( internal, external, class 1, 2, and 3 ) are shown in real time on the MCDU and are not memorized in NVM.

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5 TROUBLE SHOOTING DATA The purpose of this item is to present the encoded data ( hexadecimal ) associated to each fault detected by the EIVMU and displayed in the LLR, PLR and GR.

6 CLASS 3 REPORT The pupose of this item is to present in plain language all class 3 internal and external faults detected by the EIVMU during the last flight

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77-00

1

6

2

7

3 4

8

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5

Figure 63 FRA US-T TH NOV 99

EIVMU System Tests Page: 127

A 330 RR Trent 700

77-00 EIVMU CONTINUED 1

SPECIFIC DATA

This item can only be activated on the ground. The specific data is composed of:

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Figure 64 FRA US-T TH NOV 99

EIVMU Specific Data Page: 129

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A 330 RR Trent 700

77-00 FAN UNBALANCE Acquisition ot 8 sets of points. Acquisition ot phase and displacement in stabilized flight conditions allows to rebalance the engine when aircraft is on the ground. 8 N1 speeds shall be selected through the MCDU in the following speed ranges: – 16 – 50 percent – 50 – 65 percent – 65 – 75 percent – 75 – 80 percent – 80 – 85 percent – 85 – 90 percent – 90 – 95 percent – 95 – 99 percent These selected speeds are memorized in NVM until next change on the MCDU. Default values are: 50, 60, 67, 76, 84, 88, 92 and 96 percent rpm. Corresponding to these speeds, 8 sets of points shall be stored for each of the Iast two flights with the conditiones here under. – Flight phase 6 – N1 speed range – N1 speed accuracy ( variation allowed  2 percent ) – Stabilization time 15 seconds A set is defined as: – N1 phase – N1 vibration – N2 vibration – N3 vibration – N1 speed – N2 speed – N3 Speed – Date and UTC – Engine S/N FRA US-T TH NOV 99

Acquisition during ground run This menu allows to acquire vibration data set updated every 3 seconds

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A 330 RR Trent 700

77-00

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Figure 65 FRA US-T TH NOV 99

EIVMU Fan Unbalance Page: 131

A 330 RR Trent 700

77-00 MAX FLIGHT VIBRATION DISPLAY Max flight vibration acquisition The EIVMU shall store for each of the Iast 8 flights for N1, N2 and N3 vibrationes, and for each accelerometer the max vibration Ievel reached. For each flight these max Ievels shall be processed during the whole flight phase 6 and shall be stored in NVM with the following data sampled at the same time: – associated N1, N2 and N3 speeds – N1, N2 and N3 vibrationes from the accelerometer – N1 phase angle from the accelerometer – broadband vibrationes from accelerometer – engine S/N – date and UTC

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77-00

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Figure 66 FRA US-T TH NOV 99

EIVMU Max Flight Vibration Display Page: 133

A 330 RR Trent 700

77-00 FREOUENCY ANALYSIS A frequency analysis giving measured vibrationes from 1 to 500 HZ by steps ot 40 HZ for one engine condition shall be processed once per flight – if the broadband threshold exceeded or – for the programmed accelerometer at a given flight phase and shaft speed A specific procedure is used between the printer and the EIVMU in order to print the frequency analysis after the flight transition from flight phase 9 to 10, in phases 1 and 10 with only one successfull print allowed per accelerometer. Comments can be added in the print destination field using the specific comment Iines via the scratchpad.

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Figure 67 FRA US-T TH NOV 99

Frequency Analysis Readout Page: 135

A 330 RR Trent 700

77-00 DISCRETE INPUTS This page gives the state of each EIVMU discrete inputs. These discrete inputs are used by the EIVMU and /or by the EEC which is Iinked to the EIVMU by an ARINC 429 bus.

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Figure 68 FRA US-T TH NOV 99

Discrete Inputs Page: 137

A 330 RR Trent 700

77-00 DISCRETE OUTPUTS This page gives the state of each EIVMU discrete outputs. These discrete outputs are the result of some EIVMU Iogics. The discretes are updated every 3 seconds.

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Figure 69 FRA US-T TH NOV 99

Discrete Outputs Page: 139

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A 330 RR Trent 700

77-00 ENGINE CONDITION MONITORING DESCRIPTION Ground based Engine Health Monitoring (EHM) computer programmes are used to analyse data generated by the engine and aircraft during revenue operation. The engine condition monitoring programme must be able to predict when an engine is no longer capable of providing, within certified limits, the maximum thrust required for single engine diversion. The basic instrumentation and full gas path instrumentation are compulsory for Extended Twin Engine Operations (ETOPS) and airlines must use the data output by this instrumentation. Provision is made in the basic engine design for instrumentation and EEC software to facilitate health monitoring and gas path performance analysis down to component level. The Airplane Condition Monitoring System (ACMS) is used to record, process and store airplane system data for report generation. The reports generated by the ACMS are then used by ground analysis programmes to monitor engine and other airplane systems for performance and trend analysis. Engine Health Monitoring ground based software programmes such as COMPASS – Condition Monitoring and Performance Analysis, analyse data generated by the engine during operation, parameters are selected from those already processed for engine control and output on the data bus by the EEC, others are obtained by installing additional tappings/sensors on the engine with some parameters transmitted via the EEC and others hardwired direct from the sensor to the aircraft ACMS.

flame–out is prevented during bad weather conditions (such as heavy rain and/or hail).

IP Compressor Exit Thermocouple ( T25 ) S The T25 probe is a single thermocouple which is used for Engine Condition Monitoring HP Compressor Exit Thermocouple ( T30 ) S The 3 thermocouples are located behind the fuel nozzles and viewed from the rear in 1:00, 4:00 and 8:00 o’clock position. They are also used for combuster borescoping The EEC monitors the temperature of the air supplied to the combustion chamber (through the engine compressors) with three HP compressor exit thermocouples The temperature (T30) is used as a control parameter to make sure an engine FRA US-T TH NOV 99

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T30 THERMOCOUPLE

T25 THERMOCOUPLE

Figure 70 FRA US-T TH NOV 99

Thermocouples ( T30 / T25 ) Page: 141

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A330 RR Trent 700

73-20

ATA 73

ENGINE FUEL AND CONTROL

73-20

FADEC

GENERAL A Full Authority Digital Engine Control system (FADEC) controls the RB211–Trent engine. The FADEC system is made of sub–systems working together to form a closed loop control system, maintaining efficient engine operation at a selected condition ranging from engine start through the take–off flight/landing operation envelope to engine shut–down. NOTE: Shut–down is effected by the H.P. fuel shut–off valve torque motor (see Fuel System), which is hardwired directly from the cockpit, giving the pilot the ability to operate the shut–off valve at any time. This function has priority over any automatic H.P. fuel shut–off valve command. The sub–systems are: S Power generation and conditioning control. – Dedicated Alternator – Power Control Unit S Engine Control – Electronic Engine Control Unit ( EEC ) – Discrete Sensors – Torque Motor Devices – Solenoids S Rotor Integrity Protection. – Overspeed Protection Unit (OPU) – Turbine Overspeed Protection

FRA US-T TH NOV 99

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Figure 71 FRA US-T TH NOV 99

FADEC Overview Page: 143

A330 RR Trent 700

73-20 DEDICATED ALTERNATOR The Dedicated Alternator supplies primary power to the FADEC system and provides a speed reference signal of the H.P. shaft speed (N3). The unit is mounted on the external gearbox and driven by direct drive from the H.P. shaft N3. The Alternator consists of two separate three phase stator windings and two separate single phase stator windings. The associated rotor magnets are connected to a common cantilever shaft. (The shaft does not require bearings). The three phase circuits provide power to the E.E.C. in the speed range 8% to 125% N3. One of the phase windings in each three phase circuit provides the E.E.C. with referencing to the H.P. shaft rotational speed. The two separate single phase circuits provide power to the overspeed protection unit. For This power supply only one single phase is used and the other is unused. NOTE: A swapover can be initiated by the maintenance in case of OPU power supply problems.

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PCU CHANNEL A SUPPLY

OPU POWER SUPPLY OPU SPARE SUPPLY PCU CHANNEL B SUPPLY

Figure 72 FRA US-T TH NOV 99

Dedicated Alternator Page: 145

A330 RR Trent 700

73-20 POWER CONTROL UNIT GENERAL The P.C.U. is a dual channel unit, it rectifies, filters and regulates power supply to the respective channel of the Electronic Engine Control Unit to 22V DC and also conditions power supply to the P20/T20 probe heater, the high energy ignition circuits and cabin air bleed H.P.V. The primary source of FADEC power is provided by the Dedicated Alternator. A secondary back–up electrical supply is provided by two aircraft 115V supplies. The back–up supply provides electrical power: – during engine start up – system failure – during ground test The FADEC power supply is fed into the Power Control Unit (P.C.U.). The Dedicated Alternator power or aircraft power as appropriate is conditioned by the P.C.U. The unit is mounted on the fan case inside the suitcase.

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Figure 73 FRA US-T TH NOV 99

FADEC Electrical Power Configuration Page: 147

A330 RR Trent 700

73-20 POWER CONTROL UNIT CONTINUED The PCU is a digital unit which has two Channels of operation. The two Channels are identified as Channel A and Channel B. You can find the PCU installed in an electronic–unit protection–box. This box is for the protection of the electronic units in the FADEC system. The PCU outer cover is made of metal and is specially prepared with a layer of high emissivity paint. This type of paint gives better heat radiation to help keep the unit cool. Convection and radiation of heat from the internal power supply modules is also helped by external fins on the case surfaces. The PCU contains internal temperature sensors which are continuously monitored by the EEC. If the internal temperature increases more than a specified limit, the EEC transmits a status message to the cockpit. On the front face of the outer cover are six electrical receptacles. The top row of three receptacles are for the Channel A inputs/outputs. The bottom row of three receptacles are for the channel B inputs/outputs. The mating electrical connectors for each row of receptacles connect their related PCU Channel to: S The EEC S The EEC Dedicated Alternator S The aircraft power interfaces.

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Figure 74 FRA US-T TH NOV 99

Power Control Unit Page: 149

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A330 RR Trent 700

73-20 ELECTRONIC ENGINE CONTROL DESCRIPTION The Electronic Engine Control unit is the heart of the FADEC system. It is located on the fancase and is shielded and grounded as protection against (E . M. I.) Electromagnetic Interference The E.E.C. unit is a dual channel digital unit.The channels are identified as channel A and channel B with a communication link between each. Each channel uses the high integrity computer (H.I.C.) concept to perform software instructions and utilises dual interfaces to provide a high degree of fault tolerance. In normal operation only one channel controls, in the event of certain failurescontrol is transferred to the alternative channel. The E.E.C. also transmits engine performance data and system test data to the aircraft which is used in flight deck display, thrust management and condition monitoring systems. The E.E.C. automatically exercises full control of the engine to achieve or maintain a command signal, it automatically controls : S Engine Starting (Ground & Flight) S Relighting following flame–out detection S Control of fuel flow (thrust) in both forward and reverse thrust S V.I.G.V./V.S.V. position S The Airflow Handling Bleed Valves S The engine oil/fuel Heat Management System S Internal gearbox cooling airflow S I.P. Turbine case cooling airflow S Aircraft cabin air H.P. valve control S S P20/T20 probe heater control S The F.M.U. in the event of L.P. shaft breakage S The thrust reverser S I.P. turbine overheat data

FRA US-T TH NOV 99

EEC Air Pressure Module The EEC air pressure module contains a number of cylinder pressure transducers which give S An AC signal of frequency that is in proportion to the air pressure S A DC temperature diode signal for air temperature correction. Each air connector on the EEC front face is identified to show its related air pressure input. These inputs are: S PO, which is measured from engine Zone 1 S P20, LP compressor inlet pressure S P25, IP compressor outlet pressure S P30, HP compressor outlet pressure S P50, LP turbine outlet pressure S P160, LP compressor outlet pressure. The PO air is supplied and filtered through a special cap on the protection–box external connection. The other air pressures are supplied to the protection–box external connections by rigid tubes which have routings from the different air sources.

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Figure 75 FRA US-T TH NOV 99

Electronic Engine Controller Page: 151

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ENGINE FUEL AND CONTROL CONTROLLING

A330 RR Trent 700

73-20 EEc DESCRIPTION CONTINUED The EEC had been designed as a single unit, mounted using anti–vibration mountings high on the left hand side of the engine fan case with, and alongside the PCU and OPU. All three units are contained within a hinged cover– screw sealed box unit. The two electronic circuits or channels of the EEC are almost the same, each channel contains circuit boards or cards which match the cards in the other channel. Each card has a specific function related to it. The cards within an individual channel are inter–connected by a mother board, each mother board is connected by electronic wiring to the engine and the aircraft. A wiring loom within the EEC unit connects Channel A to Channel B. Also within the EEC unit is a pressure module which contains a number of pressure transducers and transducer interface circuits. Each transducer is connected to pressure signal pipes from the engine. On the side of the EEC unit is a connector for a Data Entry Plug (see 73–21–12) which is attached when the EEC is installed on an engine. The EEC computer reads the data and applies it to adjust the engine operation characteristics. An external test socket is incorporated in the EEC. To allow discrete external test signals to communicate with the EEC computers. EEC input data is received from the aircraft and engine sensors. All input data including signals from pressure transducers is checked and conditioned on entry into the EEC ie. all inputs are multiplexed and therefore, as necessary, converted to digital format. The inputs are also subject to radio frequency interference (RFI) filtering and lightning strike voltage protection. Temperature System S The EEC monitors the temperature of the air supplied to the combustion chamber (through the engine compressors) with three HP compressor exit thermocouples.The temperature (T30) is used as a control parameter to make sure an engine flame–out is prevented during bad weather conditions (such as heavy rain and/or hail). S The EEC monitors the temperature of the air around the IP turbine disk with IP turbine overheat thermocouples. If the air used to keep the disk cool becomes too hot (forward or rearward of the disk) the EEC transmits a warning for display at the cockpit.

FRA US-T TH NOV 99

S The EEC monitors the temperature of the fuel downstream of the FMU with two fuel temperature thermocouples. The fuel temperature and the engine oil temperature are used as the primary control parameters in the control and removal of unwanted heat from the engine S The EEC monitors the fuel flow to the combustion system with a fuel flow transmitter and calculates the fuel used. Fuel flow and fuel used data is then transmitted for display at the cockpit. S The EEC monitors the fuel pressure downstream the LP fuel pump with a fuel low pressure switch installed adjacent to the fuel outlet connection. When fuel low pressure is detected a maintenance message is sent by EEC to the CMS. Air Data Selection The EEC initially does a check that each engine or aircraft air data input (P20, T20 and P0) is correct before it is included in the selection procedure. Allowance is made for the effects of probe heater operation in relation to the P20 and the T20 inputs. The selection procedure is then as follows: Condition 1 If the engine and ADIRU 1 parameters agree then the ADIRU 1 parameter will be used to calculate the engine ratings. If they do not agree but the engine and ADIRU 2 parameters agree then the ADIRU 2 parameters will be used. Condition 2 If the engine and aircraft parameters are at almost the same values then the aircraft air data parameter will be used (as in condition 1). Condition 3 If the engine and aircraft parameters are at different values (but their difference is satisfactory) then a value between these parameters will be used. Condition 4 If the engine and aircraft parameters are at different values (but their difference is unsatisfactory) then the engine air data is used. In this condition the EPR control (in forward thrust) will be continued in relation to the engine P0 and T20 values. But the engine P20 value will have the effect that follows: S If engine P20 is more than aircraft P20 the EEC will stay in EPR control S If engine P20 is less than aircraft P20 the EEC will change to N1 reversionary control (rated).

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Figure 76 FRA US-T TH NOV 99

EEC Suitcase Page: 153

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A330 RR Trent 700

73-20 E.E.C. INTEGRITY DESCRIPTION The EEC utilises advanced computer technology features. The dual channels, both interfacing with the aircraft provide a high level of fault tolerance. Each channel of the EEC contains a High Integrity Computer (HIC). This consists of a control computer and a monitor computer. Both control and monitor computers access memory, internal data and input data from the airframe and engine sensors. Both computers process the data independently and should produce identical output. Output from both computers is fed into a comparator (parity memory) which is also within the HIC, any discrepancy results in a request for a channel change and a reset of the faulty channel. The channel change is achieved without interruption to engine control. The EEC as a computer, has multi–programmes of engine model logic. Aircraft engine input data is continuously fed into these programmes for comparison. Disagreements are processed to generate corrective output signals which are fed to engine mounted units which control engine operation. Contingency programmes are built into the EEC to cater for channel failure. In addition the EEC is programmed with logic to synthesise any primary control signal loss from available data. The FADEC system has been designed to provide a high level of fault tolerance ie. following a signal failure, loss of redundancy occurs rather than loss of function. This provides a system which is safe and reliable and secondly allows tolerance of faults for a limited period of time. This is achieved by using duplicated engine hardware, and EEC dual channel software.

FRA US-T TH NOV 99

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Figure 77 FRA US-T TH NOV 99

High Integrity Computer Schematic Page: 155

A330 RR Trent 700

73-20 ENGINE DATA PLUG The E.E.C. has been designed to control all possible configurations of the engine regardless of individual characteristics. To provide interchangeability of the unit, specific engine information must be made available to the E.E.C. i.e.: S Engine Serial Number S Thrust Rating S E.P.R./Thrust Trim Relationship S Turbine Gas Temperature Trim S Engine Standard S Intermix/Retrofit S Engine Health Monitoring S Idle Trim This is achieved by a device known as the Data Entry Plug which plugs into the E.E.C. The plug is a dual channel memory device which is programmed with relevant engine data which is used by the E.E.C. to enable correct engine operation control. The Data plug remains with the engine throughout its operational life not with the E.E.C.

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Figure 78 FRA US-T TH NOV 99

Data Entry Plug Page: 157

A330 RR Trent 700

73-20 DEP PROGRAMMING UNIT DESCRIPTION The purpose of the Programming Unit is the reprogramming of the DEP. Both EEC channels must be programmed. The following diagram shows an example of those four charts which can partially be seen on the CMS readout as well.

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Figure 79 FRA US-T TH NOV 99

DEP Programming Unit and Printouts Page: 159

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ENGINE FUEL AND CONTROL CONTROLLING

A330 RR Trent 700

73-20 OVERSPEED PROTECTION SYSTEM (OPU) SYSTEM DESCRIPTION The OPU is an independent system to the E.E.C. It is designed to prevent severe L.P. and I.P. shaft overspeed in the event of severe mal–scheduling V.S.V’s and/or fuel upward runaway due to fuel metering valve failing open. In the event of overspeed the unit indirectly signals the fuel shut–off valve to close. The primary function of the OPU is to give protection from an N1 or N2 overspeed. These engine shaft speeds are usuallykept in safe limits by ’red–line’ limiters in the EEC. But if a failure occurs such that these limiters can not prevent a shaft overspeed the OPU will shut down the engine independently of the EEC. The OPU also makes the selection of two satisfactory N1 and N2 compressor speed signals (from the three available N1 and N2 speed signals) and supplies the same satisfactory speed signals for each EEC Channel. You can find the OPU installed to the rear face of the PCU. The OPU is a two channel digital control unit. It has interfaces with: S Three speed probes for N1 speed S Three speed probes for N2 speed S The EEC for operation of BITE and to transmit N1 and N2 speeds S The overspeed valve in the FMU. The OPU outer cover is made of metal. On the front face of the outer cover are four electrical receptacles. These are used to electrically connect the OPU to: S The EEC dedicated alternator S The speed probes S Each channel of the EEC S The FMU overspeed torque motor. The OPU contains a power supplies circuit board, signal conditioning circuit board and BITE. The power supplies board is used to change the one–phase AC input from the alternator to DC supplies for the OPU. The signal conditioning board is used for the OPU functions and contains two Channels (A and B) of logic. The BITE does the selection of two satisfactory N1 and N2 speed signals from the three available N1 and N2 speed signals. And will transmit data to the EEC FRA US-T TH NOV 99

for failures of the speed probes, overspeed valve torque motor and OPU circuits. Signal Conditioning / Overspeed Detection The speed probe inputs to the OPU are initially sent through probe selection circuits. BITE controlled relays in these circuits independently supply an N1 and N2 signal to each OPU Channel: but only if they are found to be satisfactory. The satisfactory signals are also isolated and then transmitted to the EEC for usual engine control. Each OPU Channel changes its N1 and N2 analog signal input to a digital signal. These signals are then read by an ASIC which continuously monitors their values and uses logic to find an overspeed condition. If an ASIC finds an overspeed condition it will energize its output circuit. If the ASIC in the other Channel also energizes its output circuit, at the same time, then the condition is read as ’TRUE’. An electrical signal is then sent to energize the FMU overspeed torque motor. This causes: S The overspeed valve to operate and close the FMU PRSOV S The engine to shutdown independently of the usual shutdown control selections. If one ASIC energizes its output circuit and it reads that the other has not, then the OPU is automatically disabled. And a failure indication is transmitted (through the EEC) to the cockpit.

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Figure 80 FRA US-T TH NOV 99

Overspeed Protection Logic Page: 161

A330 RR Trent 700

73-20 OPU continued The OPU is made cool by convection and radiation to the air in the protectionbox. Air flows through the box when the LP compressor turns. The OPU is attached with six bolts to the rear face of the PCU. A bonding strap is permanently installed to the OPU and connects to one of the PCU mount brackets. Thus the OPU is grounded through the PCU. A unit label and modification label are on the top of the unit. An electrostatic sensitive–device warning–label and an insulation–test warning label are also on the unit.

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Figure 81 FRA US-T TH NOV 99

Overspeed Protection Unit Page: 163

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ENGINE FUEL AND CONTROL CONTROLLING

A330 RR Trent 700

73-20 TURBINE OVERSPEED PROTECTION SYSTEM DESCRIPTION The EEC monitors the LP shaft breakage. The LP TOS (Low Pressure Turbine OverSpeed) system is designed to accommodate turbine overspeed due to shaft breakage. A shaft breakage is detected by a comparison between the speed measured by the compressor speed probes mounted in the front compressor bearing housing and the turbine speed probes, mounted in the turbine bearing housing. For LP shaft breakage at any power: S the fuel flow is automatically shut off via the HPSOV overspeed torque motor.

If one logic lane ’A’ or ’B’ becomes defective the turbine overspeed circuit is disarmed. This prevents incorrect operation of the system.

L P SHAFT BREAKAGE Three L.P. compressor speed probes send signals of shaft speed to the overspeed protection unit (O.P.U.). This unit makes a selection of two satisfactory N1 signals and transmits them to the Electronic Engine Controller (E.E.C.). One N1 signal is supplied to each of the logic lanes ’A’ and ’B’ on the turbine overspeed circuit board in the E.E.C. lane ’A’. Three L.P. turbine speed probes send signals directly to the L.P. Turbine Overspeed (L.P.T.O.S.) circuit board of the E.E.C. lane ’A’. Each logic lane is supplied with one N1 signal. If one of these signals is not satisfactory then the applicable logic lane makes the selection of the alternative signal. When the L.P. rotor system reaches a speed higher than 1000 R.P.M. the turbine overspeed protection system is armed. Each logic lane compares the L.P. compressor speed with its L.P. turbine speed. If the two logic lanes detect a speed difference between L.P. compressor and L.P. turbine in a specified time limit it is accepted as a true failure condition. If an L.P. shaft failure is accepted as true the system will signal a closure of the Pressure Raising and Shut–off Valve (P.R.S.O.V.). The engine is immediately shut down. Once the fuel flow has been shut off the P.R.S.O.V. is latched in the fuel off position. If inadvertent shut down occurs the pilot has a reset facility in the flight deck. If there is a failure of a compressor speed signal, which shows that overspeed of the turbine is not possible, the related overspeed protection circuits are disarmed. FRA US-T TH NOV 99

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Figure 82 FRA US-T TH NOV 99

Overspeed Pretection Simplified Diagram Page: 165

A330 RR Trent 700

73-20 Turbine Overspeed BITE The E.E.C. does a test of the L.P. turbine overspeed function during each ground start – pre light up.The B.I.T.E. provides the necessary turbine speed difference to the turbine overspeed protection circuits to momentarily shut off the fuel during start sequences. Movement of the P.R.S.O.V. in the F.M.U. to the closed position is monitored by the E.E.C. Almost immediatel the E.E.C. cancels the B.I.T.E. test signal to cause the P.R.S.O.V. to become open again. Therefore the engine start sequence is not stopped by the test. Defects found during the B.I.T.E. test are stored in the E.E.C. The defects are subsequently transmitted to the Central Maintenance Computer (C.M.C.).

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A330 RR Trent 700

73-20

EEC

N1C N1T N1T

OVERSPEED

SERVO

LOGIC A

VALVE

PRSOV TO

PROBE DRIVE

SELECT

N1T

OVERSPEED

ENGINE

TORQUE OVERSPEED

N1C

MOTOR

LOGIC B

FMU

TURBINE EXHAUST CASE

LPTOS COVER PLATE LP TURBINE BEARING HOUSING COVER

TERMINAL UPPER SPEED PROBE

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TERMINAL CENTER SPEED PROBE

TERMINAL LOWER SPEED PROBE LP TURBINE OVERSPEED ELECTRICAL HARNESS

Figure 83 FRA US-T TH NOV 99

Turbine Overspeed Protection Page: 167

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ENGINE FUEL AND CONTROL CONTROLLING

A330 RR Trent 700

73-20 P20 / T20 PROBE DESCRIPTION The P20/T20 probe is mounted inside the air intake cowl at 15_ to right of top dead centre when viewed from rear. The probe measures both engine intake pressure and temperature. Temperature is measured by two independent platinum resistance elements. A small amount of air passes over the elements, whilst the rest of the air passes straight through the probe. The pressure signal offtake is just above where the main airstream flows through the probe. A pipe passes through the body to the pressure connector on the base plate. P20 / T20 PROBE HEATER CONTROL The EEC autimatically makes the selection of the T20 / P20 probe heater elements to prevent ice on the probe air inlets. The elements are set to ON when the engine speed is higher than 10% N1. And OFF when the engine speed is lower than 10% N1 ( or aircraft is on ground and the engine speed is lower than 45% N3) The EEC software allows to perform a probe heater test via the CMC on the MCDU. This is a active test where the power supply and the temperature increment is checked.

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WARNING If you remove the inner row of nuts the probe will fall into the engine air intake. This can cause injury and/or damage.

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Figure 84 FRA US-T TH NOV 99

P20 / T20 Sensor Page: 169

A330 RR Trent 700

73-20 POWER MANAGEMENT DESCRIPTION A Full Authority Digital Engine Control System (FADEC) controls the RB211–Trent engine.The FADEC system is described further in the FADEC Section’. The FADEC schematic outlines: S Sub–systems – For specific system refer to respective section, i.e. Airflow Control, Fuel Control, Propulsion System. S E.E.C. input and output signals S The interface with the aircraft systems

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Figure 85 FRA US-T TH NOV 99

FADEC Schematic Page: 171

A330 RR Trent 700

73-20 THROTTLE CONTROL LEVER MECHANISM DESCRIPTION Movement of the flight deck throttle control lever generates a command signal for the Electronic Engine Control (E.E.C.) unit, which converts the signal to an E.P.R. value or N1 value. (see FADEC section). Each throttle lever is mechanically connected to two throttle resolvers which convert Throttle Lever Angle (T.L.A.) into a Throttle Resolver Angle (T.R.A.) command signal. The resolvers are independent but produce the same output signal. Each resolver is dedicated to one channel of the E.E.C. The throttle control lever moves a total arc path of approximately 55 degrees. Forward thrust – is selected by moving the lever through an arc > 4.5 degrees. Reverse thrust – is selected by moving the reverse thrust lever through an arc of 96 degrees.

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Figure 86 FRA US-T TH NOV 99

Thrust Control Page: 173

A330 RR Trent 700

73-20 THROTTLE CONTROL LEVER MECHANISM S Forward thrust Three detents are provided in the forward thrust range i.e.: – maximum climb (MCLB) at 30 degrees – maximum continuous (MCT) at 42 degrees – maximum take off (TOGA) at 55 degrees S Reverse thrust Lifting the reverse thrust lever allows the throttle to operate in the reverse thrust range. Maximum reverse power – TLA at 96 degrees.

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Figure 87 FRA US-T TH NOV 99

Combined TRA Relationship Page: 175

A330 RR Trent 700

73-20 BASIC CONTROL LOOP–STEADY STATE DESCRIPTION The Electronic Engine Control (E.E.C.) receives a command signal from the flight deck Throttle Resolver (T.R.A.) which is converted by the E.E.C. into an Engine Pressure Ratio (E.P.R.) or N1 demand. Alternatively the E.E.C. receives a F.M.G.E.C./auto–thrust computed E.P.R.signal. The E.E.C. also receives signals from engine mounted and aircraft (air data) sensors. The command signal and other relevant input signals are processed within the E.E.C. Output control signals are transmitted to engine accessory mounted control units. The primary engine control unit being the Fuel Metering Unit (F.M.U.). Insidethe F.M.U., a torque motor receives the E.E.C. output signal.The torque motor modulates fuel servo pressur to move the Fuel Metering Valve (F.M.V.) which is integral of the F.M.U. Indirectly the F.M.V. is adjusted to control fuel flow to match an E.P.R./thrust demand signal. Once operating, many of the engine accessories feed status signals back to the E.E.C. e.g. movement of the F.M.V. is sensed by a double resolver which feeds back a position signal to the E.E.C. The E.E.C. uses these feedback signals to make comparisons with software logic and as a consequence of any disagreement, process output signals to engine control units as necessary i.e. adjust fuel schedule accordingly.

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Figure 88 FRA US-T TH NOV 99

Power Setting - Basic Control Loop Page: 177

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ENGINE FUEL AND CONTROL CONTROLLING

A330 RR Trent 700

73-20 THRUST MODES Manual Mode When in manual mode the flight deck throttle levers are used exclusively to control engine thrust. Automatic Thrust Control The auto thrust function provides automatic computation of the thrust level to be set in order to achieve the desired aircraft flight characteristics. NOTE: During take–off the automatic thrust control is not active. The automatic thrust control function is part of the auto flight system: i.e. the auto thrust system interfaces with the aircraft flight management guidance envelope computers which receive and provide output signals to control the aircraft flight services via the electronic flight controls system (E.F.C.S.) and to the engines via the FADEC. When the auto thrust function is active, moving the throttle lever (T.L.A.) into idle, climb or maximum continuous detents, command the auto thrust function accordingly

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Memo Mode This is a transitive mode of thrust control between the autothrust mode and manual mode of the autothrust function. When the autothrust mode is deactivated and the throttle levers are set on the max continuous or max climb detent points, the E.E.C. will enter the memo thrust mode. In this mode the thrust demand is locked by the E.E.C prior to exiting autothrust mode.This is to prevent potential thrust step changes which may occur when reverting from autothrust to manual mode.

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Figure 89 FRA US-T TH NOV 99

Forward Thrust - Throttle Legend Page: 179

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ENGINE FUEL AND CONTROL CONTROLLING

73-20 THRUST SETTING E.P.R. Control The E.E.C. controls the engine to an E.P.R. schedule during manual operating conditions. If the autothrust function is active then the E.E.C. controls the engine to an E.P.R. target supplied by the F.M.G.E.C. N1 Reverse Thrust Control The E.E.C. controls the engine to an N1 reverse schedule if the T.R.A. indicates that the pilot has selected the reverse thrust lever. N1 Reversionary Control The E.E.C. controls the engine to an N1 reversionary schedule e.g. as a result of pilot command or loss of actual E.P.R. parameters. There are 2 forms of N1 reversionary control : S Rated N1 Reversionary Mode The E.E.C. calculates an E.P.R. command as in the E.P.R. control mode. The E.E.C. then converts this E.P.R. command into an N1 command using a simple look–up table and the engine is controlled using this N1 command. S Unrated N1 reversionary Mode The E.E.C. sets the forward idle detent position equal to idle N1 and the max take–off detent position equal to red–line N1. The E.E.C. then interpolates between the two N1 speeds so as to maintain an approximately linear thrust vs T.R.A. relationship between the two detent positions. The engine is then controlled using this N1 command.

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A330 RR Trent 700

Idle Control At idle the EEC controls the engine so to prevent the engine being operated below certain minimum operating limits. The limits are: S minimum P30 limiter – minimum P30 pressure demand necessary to maintain nacelle, wing and ECS bleed air requirements S minimum N1 limiter – minimum N1 shaft speed to prevent icing of the spinner FRA US-T TH NOV 99

S minimum N3 limiter – minimum N3 shaft speed necessary to maintain aircraft services and allow acceleration to go-around thrust setting mode within a predefined time limit S minimum fuel flow limiter – minimum fuel flow limiter to maintain combustion integrity S minimum T30 – minimum T30 necessary to protect against adverse weather conditions

ENGINE RATING To cater for FADEC E.E.C .interchangeability, the E.E.C’s memories stored with rating information to cater for up to ten possible different rated engines. The rating information is stored in the form of tables. The table to be used by the E.E.C. is selected using rating index data stored within the Data Entry Plug (D.E.P.). Basic E.P.R. Rating The E.E.C. calculates the E.P.R. value corresponding to: S Maximum Take–off/Go–Around S Flexible Take–off S Maximum Continuous S Maximum Climb S Maximum Take–off/Go–Around This is the maximum thrust which the engine can give for take–off under the ambient conditions. Flexible Take–off/Derated Take–off The Trent engine is flat–rated i.e. thrust versus ambient temperature/pressure relationship up to a kink–point temperature. After this kink–point the thrust that the engine can produce decreases due to T.E.T. limitations. The flexible take–off system allows the pilot (or by derate selection) to specify a modified ambient temperature greater than the maximum take–off kink point temperature which reduces the engine thrust output as a factor of aircraftweight.

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A330 RR Trent 700

73-20

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ISA + 15

Figure 90 FRA US-T TH NOV 99

Flat Rating Page: 181

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A330 RR Trent 700

73-20 CMS EEC INTERACTIVE TESTS GROUND SCANNING The purpose of this item is to analyze on the ground only the failts that occured during the last flight. All the faults detected while using this function are shown in real time on the MCDU and are not memorozed in non-volatile memory. CLASS 3 REPORT This function shows the internal or external class 3 faults detected by the EEC during the last flight.

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GROUND REPORT This function gives access to the internal faults detected by the EEC which are occured after landing.

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Figure 91 FRA US-T TH NOV 99

Ground Scanning / Class 3 Faults / Ground Report Page: 183

A330 RR Trent 700

73-20

EEC SYSTEM TEST This function enables the maintenance the initiation of the EEC test from the MCDU. Upon selection of this function the EEC performs a power up test plus 20 sec of ground scanning.

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A330 RR Trent 700

73-20

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Figure 92 FRA US-T TH NOV 99

EEC System Test Page: 185

A330 RR Trent 700

73-20 CMS PROBE HEATER TEST This function gives the ability for the maintenance to perform an active P20/T20 probe heater test. CAUTION: P20/T20 probe will get hot during test!

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Figure 93 FRA US-T TH NOV 99

EEC Probe Heater Test Page: 187

A330 RR Trent 700

73-20 ENGINE RUNNING TEST When selecting ENGINE RUNNING TEST the maintenance has the ability to simulate the engine running relay switch closure in engine running condition for the aircraft systems.

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73-20

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Figure 94 FRA US-T TH NOV 99

CMS Engine Runningn Test Page: 189

A330 RR Trent 700

73-20 SPECIFIC DATA READOUT S EEC CONFIGURATION – This readout is used to compare or readout the programmed engine EEC data from the Data Plug and the engine Name Plate. S EEC EXCEEDANCE – This function gives the readout of the exceedance values from N1, N2, – N3 and EGT which are memorized since last engine operation.

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Figure 95 FRA US-T TH NOV 99

Specific Data Readout Page: 191

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POWER PLANT OIL SYSTEM

79-00

ATA 79

OIL

79-00

OIL SYSTEM

SYSTEM DESCRIPTION The oil system is a full flow recirculatory system and its function is to supply oil to the engine internal drives, gears and bearings.The oil is used to lubricate these locations and remove unwanted heat throughout all operating conditions. Components within the system must ensure that the oil supplied to these drives, gears and bearings is in the correct condition with regard to cleanliness, pressure, temperature and quantity. The complete system is divided into three main areas: S Feed Oil, lubrication and cooling (pressure side) S Return oil (Scavenge side) S Breather system (Vent)

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A 330 RR Trent 700

Vane type pumps are used to move the oil around the system, a total of 8 vane elements are assembled on to two rotors: S Pressure Pump Element S L.P. Turbine Bearing Chamber Scavenge Element S H.P./I.P. Turbine Bearing Chamber Scavenge Element S Internal Gearbox Scavenge Element S Front Bearing Chamber Scavenge Element S Intermediate Gearbox and Gearbox Input Drive Assembly Scavenge Element S External Gearbox Scavenge Element S Centrifugal Breather Scavenge Element

essary filtration. Location’ for magnetic chip detectors (M.C.D’s) are provided in the scavenge lines. The system is vented through a centrifugal breather, located in a housing on the front of the external gearbox. A self contained oil tank is mounted on the right hand side of the fan case. It incorporates a quantity sight glass and provision is made for pressure and gravity oil filling. The following indications are provided on the flight deck: S Oil quantity in the tank S Oil temperature S Oil pressure S Pressure Filter Impending Blockage S Scavenge Filter Impending By–pass

Two oil coolers are used in the system, a fuel cooled oil cooler (F.C.O.C.) and an air oil heat exchanger (A.O.H.E.). A pressure filter, scavenge filter and line filters (last chance) provide the nec-

FRA US-T TH NOV 99

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Figure 96 FRA US-T TH NOV 99

Oil System Schematic Page: 193

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POWER PLANT OIL SYSTEM

79-00 FEED OIL, LUBRICATION AND COOLING General Feed oil is circulated by a single pressure pump which draws oil from the oil tank through a gauze strainer. The oil system is protected against pressures exceeding 635 psi by a pump relief valve which relives excess pressure back to the pump inlet. Excessive pressure may be due to very cold oil or system blockage. Feed oil is cleaned by a 125 micron filter. A differential pressure switch monitors filter condition and provides a flight deck indication that the filter is becoming clogged, this switch is set to operat at a differential pressure of 13 psi. The A.O.H.E. combined with the F.C.O.C. will keep the oil and fuel temperatures within specified limits. Fan air (cool) is supplied to the A.O.H.E. to decrease the oil temperature when significant oil cooling is required. The fan air is shut–off when the F.C.O.C. can control the oil/fuel temperature on its own. From the A.O.H.E. the oil is supplied to the F.C.O.C., this component has two functions. The primary function is to decrease the temperature of the oil. The secondary function is to increase the temperature of the fuel. This will prevent the water content in the fuel from forming into ice particles and blocking the filter. The fuel filter is fitted in the bottom of the F.C.O.C.

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A 330 RR Trent 700

The F.C.O.C. has two by–pass valves. One is the oil pressure by–pass to give protection to the cooler core. The other is fuel filter by–pass which operates when the filter becomes clogged. An anti–syphon tube prevents oil suction from the F.C.O.C. during engine shut down.

Oil is returned from each of the six primary lubricated locations of the engine and the breather ( air / oil separator ) The lubrication locations are as follows: S The front bearing chamber S The internal gearbox S The HP / IP bearing chamber S The LP bearing chamber S The intermediate gearbox assembly S The external gearbox Oil from each location is drawn by its own vane type scavenge element, fitted in the oil pump assembly. Provision is made for magnetic chip detectors to be fitted in each of the oil return lines for trouble shooting. Oil outlet from the scavange pumps join to form a combined scavange return flow and is sampled by a master MCD before passing through a 30 micron filter. The filter has a bypass valve that will function at 20 PSI pressure differential. Filter condition is monitored by a pressure differential switch set at 13 PSI to provide flight deck indication of impending bypass. Temperature sensors are located in the scavenge return line between the scavenge filter and the tank to provide flight deck indication of oil temperature. The oil returned to the tank is discharged over a deaerator tray to release the entrained air prior to circulation.

From the F.C.O.C. the feed oil is supplied through external tubes to the main engine bearings, gears and drives.

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Figure 97 FRA US-T TH NOV 99

Oil System Diagram Page: 195

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POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 VENTING SYSTEM General Oil loss from the main bearing chambers is prevented by the use of grooved labyrinth seals pressurised by air. To contain the oil within the bearing chambers air enters the annular space between the stationary and rotating parts of the seal. The airflow inwards across the seal opposes any escaping oil and carries it back into the bearing chambers. To maintain the pressure drop across the seals, the bearing chambers – with the exception of the L.P. turbine bearing chamber – are vented by external tubes to the centrifugal breather. The pressure drop across the seals is controlled by restrictors in the vent return tubes. For the L.P. bearing housing the scavenge pump is able to maintain the pressure drop. The centrifugal breather separates the air and oil before directing the air to atmosphere, the oil is scavenged from the breather housing back to the oil tank.

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The remaining sealing air which is returned to the oil tank with the scavenge oil is separated from the oil by the de–aerator cone in the oil tank. The separated air is vented by an external tube to the centrifugal breather.

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Figure 98 FRA US-T TH NOV 99

Oil System Diagram Page: 197

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POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 OIL TANK DESCRIPTION The oil tank is attached to the front flange of the L.P. compressor case on the right hand side. The tank is a magnesium casting to which other components attach to make up the oiltank assembly. These components are as follows: S Oil quantity transmitter S Sight glass S Pressure fill and overflow connection S Oil filler assembly S Scavenge filter assembly S Outlet tube S Vent tube To help release the air from the scavenge oil returning to the tank there is a deaerator fitted inside the tank, the released air passing out of the vent tube. There is also a filter in the tank to prevent contamination of the oil pressure system supply.

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The oil filler assembly has a quick release ’cap. Internally the filler has a valve to prevent opposite flow if the cap was inadvertently left off.

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Figure 99 FRA US-T TH NOV 99

Engine Oil Tank Page: 199

A 330 RR Trent 700

79-00 SCAVANGE FILTER ASSEMBLY The diagram below shows the positions of the: S Scavenge filter assembly S Master M.C.D. S Scavenge filter differential pressure switch S Oil temperature sensors S Scavenge filter by–pass valve

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Figure 100 FRA US-T TH NOV 99

Scavenge Filter Assembly Page: 201

A 330 RR Trent 700

79-00 OIL PUMP / SCAVENGE FILTER The diagram below shows the positions of the: S Oil pump S Pressure filter housing S M.C.D. S Scavenge filter housing Master M.C.D.

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FRA US-T TH NOV 99

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POWER PLANT OIL SYSTEM

Figure 101 FRA US-T TH NOV 99

Oil Pump and Scavenge Filter Page: 203

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 OIL PUMP / MCD HOUSINGS The folowing diagram shows the provision made for fitting the M.C.D. into the six scavenge line positions.

For Training Purposes Only

Normally the M.C.D. probes and housings are not fitted and blanks cover the aperture. M.C.D. and housing can be fitted for trouble shooting individual scavenge lines, this would occur if the master M.C.D. was found to be loaded.

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POWER PLANT OIL SYSTEM

Figure 102 FRA US-T TH NOV 99

Oil Pump / MCD Housings Page: 205

A 330 RR Trent 700

79-00 CENTRIFUGAL BREATHER DESCRIPTION The centrifugal breather has a rotor that contains retimet segments and is driven by the external gearbox. Aerated oil from the bearing chamber vent system and the oil tank de–aerator tray is delivered to the centrifugal breather. The aerated oil tries to pass through the retimet segments but is centrifuged out. The air can pass through the retimet segments into the hollow rotor and is vented overboard. The centrifuged oil is scavenged back to the oil tank by its own scavenge pump element.

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FRA US-T TH NOV 99

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POWER PLANT OIL SYSTEM

Figure 103 FRA US-T TH NOV 99

Centrifugal Breather Page: 207

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 HEAT MANAGEMENT SYSTEM DESCRIPTION The function of the heat management system is to keep the engine oil and fuel temperature within specified limits to contribute to achieving the best engine performance. To do this an Air Oil Heat Exchanger (A.O.H.E.) and a Fuel Cooled Oil Cooler (F.C.O.C.) are used.

For Training Purposes Only

A.O.H.E. The A.O.H.E. is fitted on the right hand side of the L.P. compressor case and consists of two main components: S Heat Exchanger Assembly S Air Modulating Valve The operation of the A.O.H.E. is controlled by the E.E.C. The E.E.C. uses data collected from oil temperature sensors and fuel thermocouples to control the A.O.H.E. The E.E.C. monitors these temperatures and makes sure they remain within limits. If they are not then the E.E.C. will send a signal to the torque motor in the servo valve. Signals from the E.E.C. move the torque motor which causes the spool valve to move. As the spool valve moves it allows servo fuel to the applicable side of the piston and moves it. The movement of the piston will, through a linkage, turn the air modulating valve to give the best airflow necessary through the heat exchanger.

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POWER PLANT OIL SYSTEM

Figure 104 FRA US-T TH NOV 99

Heat Management System Page: 209

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 HEAT MANAGEMENT CONTINUED The A.O.H.E. has two modes. Mode 1 S Air modulating valve fully closed. This mode is for usual conditions. No airflow through the heat exchanger. Temperature of oil and fuel can be controlled by the F.C.O.C. operation only.

For Training Purposes Only

Mode 2 S Air modulating valve turns to L.P. compressor airflow open position. This mode is used when cool air is required to lower the temperature of the oil. S In mode two the air modulating valve is fully adjustable between its minimum and maximum open position.The position of the valve is controlled by the E.E.C .adjusting the valve to the best position.This position is when sufficient cool air is used for the A.O.H.E. to function, with minimum decrease in engine performance. An oil spring attached to the pinion gear will move the valve to mode two, if there is a system failure.

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POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00

MODE 1

For Training Purposes Only

MODE 2

Figure 105 FRA US-T TH NOV 99

AOHE Air Modulating Valve Page: 211

A 330 RR Trent 700

79-00 FUEL COOLED OIL COOLER DESCRIPTION The F.C.O.C. has two functions: The primary function is to reduce the temperature of the oil and the secondary function is to increase the temperature of the fuel, this will prevent the water content in the fuel from turning to ice. The oil flow through the core is made slower by many baffle plates around the steel tubes. The slower oil enhances the exchange of heat. If the oil pressure in the F.C.O.C. becomes more than a specified limit a by–pass valve will open and relieve the pressure back to the oil inlet. An anti–syphon hole connects the inlet to the outlet to prevent oil suction from the F.C.O.C. during engine shut down. The fuel filter is fitted in the bottom of the F.C.O.C. this filter has a by– pass valve which will operate if the filter becomes clogged.

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POWER PLANT OIL SYSTEM

Figure 106 FRA US-T TH NOV 99

Fuel Cooled Oil cooler Page: 213

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 AIR OIL HEAT EXCHANGER DEACTIVATION The following diagram shows the air / oil heat exchanger which is located on the right hand side of the fan case. AIR / OIL HEAT EXCHANGER The MEL deactivation procedure requires a locked open valve. This makes sure that under all circumstances a maximum cooling is provided to keep the engine oil cool for safe engine operation.

For Training Purposes Only

PROCEDURE S open the right hand fan cowl door S turn the manual turning device in clockwise direction until the valve indicates fully open S remove the cotter pin from the lockpin S remove the lockpin from the heat exchanger S remove the lockpin spacer which is part of the lockpin and store it for later installation S install the lockpin less spacer into the heat exchangerunit S install the cotter pin into the lockpin S release the exchanger air valve manual turning device S close the fan cowl door

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POWER PLANT OIL SYSTEM

Figure 107 FRA US-T TH NOV 99

Air / Oil Heat Exchanger Page: 215

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 I.D.G. OIL COOLING INTRODUCTION The Integrated Drive Generator (I.D.G.) is the primary source of A.C. electrical power supply to the aircraft. Each engine has an I.D.G. mounted on the left hand side rear face of the external gearbox. An oil system which is an integral part of the I.D.G., lubricates the I.D.G. bearings and keeps it cool. The system is connected to an external Air Cooled Oil Cooler (A.C.O.C.) to keep the oil temperature at a satisfactory level.

For Training Purposes Only

DESCRIPTION The I.D.G. includes a Constant Speed Drive (C.S.D.) and an A.C. generator installed in the one housing. The C.S.D. maintainsa constant output speed of 24,000 R.P.M., thus the generator turns at a constant speed of 24,000 R.P.M. The I.D.G. has an oil system which lubricates the generator bearings and keeps it cool. The I.D.G. has a pressure filling point, a drain point and an oil level sight glass. In the scavenge part of the oil system there is a filter that can be removed. A pop–out button gives visual indication that the filter is clogged. But when the oil is cold a bimetal element prevents this function.

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POWER PLANT OIL SYSTEM

Figure 108 FRA US-T TH NOV 99

Integrated Drive Generator Page: 217

A 330 RR Trent 700

79-00 AIR COOLED OIL COOLER DESCRIPTION The A.C.O.C. i a simple air oil heat exchanger, mounted on the lower L.H. side of the L.P. compressor case. Hot oil from the I.D.G. flows through the matrix, where it is cooled by L.P. compressor air, before returning to the I.D.G. There is a pressure relief valve (by–pass) between the oil inlet and outlet connections. If the oil is cold it will not flow easily through the matrix therefore thevalve will open and the oil by–passes the A.C.O.C.

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POWER PLANT OIL SYSTEM

Figure 109 FRA US-T TH NOV 99

Air cooled Oil Cooler Page: 219

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 FILLING THE ENGINE OIL SYSTEM WARNING You must not remove the oil tank filler cap for five minutes after the engine stops. This will let the pressure in the oil tank decrease. NOTE Only oils approved by Rolls–Royce can be used in the Trent engine. The following is a list of approved oils: S AeroShell Turbine Oil 500 (Royco Turbine Oil 500). S AeroShell Turbine Oil 555 (Royco Turbine Oil 555). S AeroShell Turbine Oil 560 (Royco Turbine Oil 560). S Mobil Jet Oil II. S Mobil Jet Oil 254. It is desirable to keep to one brand of oil but approved brands may be mixed if operationally essential.

For Training Purposes Only

GRAVITY FILLING S Open access panel on the right hand fan door. S Check oil level on sight glass. S If low, remove filler cap from oil tank. S Add the approved oil to the oil tank. S Replace filler cap. S Record amount of oil used. PRESSURE FILLING S Remove wire locking and blanking caps from pressure fill and overflow. S Fit the drain hose to the overflow coupling. S Fit the pressure filling hose to the pressure fill coupling. S Using the pressure filling equipment add the approved oil until a small quantity of oil is seen to come out of the overflow coupling. S Remove pressure filling equipment. S Replace blanking caps and wirelock. S Record amount of oil used.

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POWER PLANT OIL SYSTEM

Figure 110 FRA US-T TH NOV 99

Engine Oil Servicing Page: 221

Lufthansa Technical Training

POWER PLANT OIL SYSTEM

A 330 RR Trent 700

79-00 M.C.D. REMOVAL, INSPECTION AND REPLACEMENT Examine the M.C.D. as follows: S Gently wash in clean Kerosene to remove all oil, the Kerosene must be in a clean –container which is not made of metal. S Examine M.C.D. probe in a good light for contamination. Use a magnifying glass which will show contamination at least 5 times larger than its correct size. S The Maintenance Manual will explain the accept/reject standard. If there is contamination which can not be easily identified, it must be sent to a laboratory for analysis. It is also helpful to the engine shop if you send the contamination which has caused the engine rejection with the engine. You must also identify the M.C.D. position. Open access door on the right hand fan cowl door. Cut locking wire securing M.C.D. Turn and release M.C.D. with suitable wrench. Cut and discard seal ring.

For Training Purposes Only

INSTALLATION OF M.C.D. S Fit new seal ring. S Fit M.C.D. into its housing tighten with suitable wrench torque loading to between 60 and 120 lbf/in (0,68 and 1,35 MdaN). S Make safe with wire locking. S Close access door. NOTE: For leak check perform an Engine Idle Run.

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Figure 111 FRA US-T TH NOV 99

Master MCD Page: 223

A 330 RR Trent 700

79-00 PRESSURE OIL FILTER DESCRIPTION The oil pressure filter is screwed into the oil pump housing.It is required to remove the filter element with the filter housing. If the filter element is not removed at the same time it can fall and be damaged. When turning counter-clockwise a check valve closes to prevent oil drainage. NOTE: The filter is a cleanable element and can be used provided it is not damaged.

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POWER PLANT OIL SYSTEM

Figure 112 FRA US-T TH NOV 99

Oil Pressure Filter Page: 225

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OIL SYSTEM INDICATING

A330 RR Trent 700

79-30

79-30

OIL INDICATING SYSTEM

GENERAL DESCRIPTION The oil system has a number of comonents that are used for indication. These components monitor the oil temperature, pressure and quantity. The components send signals which permit the flight crew to monitor the status of the oil system and alert them to possible problems.

For Training Purposes Only

OIL QUANTITY TRANSMITTER The quantity transmitter sends a signal to the EEC that is in proportion to the oil level in the tank. The quantity is shown on the ENGINE SYSTEM PAGE

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OIL SYSTEM INDICATING

Figure 113 FRA US-T TH NOV 99

Oil Quantity Transmitter Page: 227

A330 RR Trent 700

79-30 OIL TEMPERATURE SENSORS There are two oil temperature sensors located in the scavenge return line. They send their signals to both EEC-channels where it is further transmitted to the ENGINE SYSTEM PAGE

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OIL SYSTEM INDICATING

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79-30

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OIL SYSTEM INDICATING

Figure 114 FRA US-T TH NOV 99

Oil Temperature Sensors Page: 229

A330 RR Trent 700

79-30 OIL PRESSURE TRANSMITTERS There are two oil pressure transmitters located on the lubrication unit. They send their signal to both EEC-channels to indicate the supply pressure on the ENGINE SYSTEM PAGE

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OIL SYSTEM INDICATING

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79-30

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OIL SYSTEM INDICATING

Figure 115 FRA US-T TH NOV 99

Oil Pressure Transmitters Page: 231

A330 RR Trent 700

79-30 LOW OIL PRESSURE SWITCH The low oil pressure switch closes if the oil pressure decreases to 35 PSI. The signal is transmitted to the FWC and to the flight deck

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OIL SYSTEM INDICATING

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79-30

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OIL SYSTEM INDICATING

Figure 116 FRA US-T TH NOV 99

Low Oil Pressure Switch Page: 233

A330 RR Trent 700

79-30 OIL PRESSURE FILTER DIFFERENTIAL PRESSURE SWITCH The differential pressure switch located on the lubrication unit sends a signal to the EEC if the differential pressure increases to 13 PSI

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OIL SYSTEM INDICATING

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79-30

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OIL SYSTEM INDICATING

Figure 117 FRA US-T TH NOV 99

Oil Pressure Filter Differential Pressure Switch Page: 235

A330 RR Trent 700

79-30 OIL SCAVENGE FILTER DIFFERENTIAL PRESSURE SWITCH The differential pressure switch located close to the oil tank sends a signal to the EEC if the differential pressure increases to 13 PSI

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OIL SYSTEM INDICATING

Figure 118 FRA US-T TH NOV 99

Oil Scavenge Filter Differential Pressure Switch Page: 237

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ENGINE FUEL AND CONTROL GENERAL

A 330 RR Trent 700

73-00

ATA 73

ENGINE FUEL AND CONTROL

73-00

FUEL SYSTEM PRESENTATION

INTRODUCTION The function of the system is to receive fuel from the aircraft tanks and deliver conditioned metered fuel into the combustion chamber for ignition. The fuel system is divided into S Fuel Control S Fuel Supply Fuel control is achieved electro–mechanically by the FADEC system or E.E.C., interfacing with the Fuel Metering unit which is integral with the fuel supply system. Note: FADEC control principles is covered in FADEC section. The fuel supply system is required to: S Uplift the fuel delivery pressure sufficient to cater for system pressure drop and fuel metering. S To heat the fuel in cold conditions. S To filter the fuel. S To meter the fuel delivery to satisfy engine thrust requirement. S To finely atomise the fuel and air mix during injection into the combustor. S Incorporate independent devices to shut–off fuel delivery to the combustor, in the event of severe shaft overspeed conditions and extremely unlikely shaft failures. S Incorporate a flight deck manually operated fuel shut–off valve.

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ENGINE FUEL AND CONTROL GENERAL

Figure 119 FRA US-T TH NOV 99

Fuel System Schematic Page: 239

Lufthansa Technical Training

ENGINE FUEL AND CONTROL GENERAL

A 330 RR Trent 700

73-00 FUEL COMPONENTS Low Pressure Pump (L.P.P.) The pump receives fuel from the aircraft system and ensures satisfactory pressure to the High Pressure Pump (H.P.P.). The pump has a single stage centrifugal impeller. Low Pressure Fuel Cooled Oil Cooler Fuel from the L.P.P. passes through the cooler to act as an oil cooling medium, and conversely for the oil to heat the fuel. Low Pressure Filter The low pressure filter is a 40 micron non–cleanable element housed in a casing separate but common with the Low Pressure Fuel Cooled Oil Cooler unit. The filter provides fuel filtration before the fuel enters the High Pressure system. A pressure differential switch is incorporated across the filter working at 5 psid and giving a flight deck warning. The by–pass operates at 25 psid.

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Low Pressure Fuel Pressure Switch This monitors the fuel pressure downstream of the element giving an indication to the flight deck if the fuel pressure falls below 70 psig.

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ENGINE FUEL AND CONTROL GENERAL

Figure 120 FRA US-T TH NOV 99

FCOC and Filter Page: 241

Lufthansa Technical Training

ENGINE FUEL AND CONTROL GENERAL

A 330 RR Trent 700

73-00 FUEL COMPONENTS High Pressure Pump (H.P.P.) The High Pressure Pump is a spur–gear type pump. It feeds fuel to the F.M.U. and provides servo pressure to the engine control unit actuators. Both the H.P. pump and the L.P. pump are housed in a common housing mounted and driven from the High Speed Gearbox. The H.P.P. is protected by a pressure relief valve which opens at 1600 psid. if a restriction occurs downstream of the pump. The valve returns H.P. fuel back to the H.P. pump inlet.

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Fuel Flow Transmitter This unit provides a signal of engine fuel flow to the flight deck. It is a displacement type unit which eliminates density variations due to temperature changes in the fuel.

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ENGINE FUEL AND CONTROL GENERAL

Figure 121 FRA US-T TH NOV 99

Fuel Pump Assembly Page: 243

A 330 RR Trent 700

73-00 Fuel Metering Unit (F.M.U.) This unit interfaces with the E.E.C. Unit. It fastens directly onto the L.P./H.P. fuel pump housing. The F.M.U. receives electrical signals from the E.E.C., to indirectly control fuel flow into the combustor. The signal controls the position of the Fuel Metering Valve. A spill valve incorporated within the F.M.U. maintains a constant pressure drop across the Fuel Metering Valve. Incorporated within the F.M.U. is a turbine overspeed valve and fuel shut–off valve. The shut–off valve can be operated by an electrical signal from the Pilots fuel control switch.

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ENGINE FUEL AND CONTROL GENERAL

Figure 122 FRA US-T TH NOV 99

Fuel Metering Unit Page: 245

Lufthansa Technical Training

ENGINE FUEL AND CONTROL GENERAL

A 330 RR Trent 700

73-00 FUEL COMPONENTS H.P. Fuel Filter The filter is a 250 micron element housed in a case attached to the inlet of the fuel manifold. The filter can be removed, cleaned and re–used. The function of the filter is to prevent blockage of the fuel spray nozzles

HP Fuel Filter

For Training Purposes Only

Fuel Manifold The primary fuel manifold is assembled in 2 halves and fits around the combustion outer case. The main fuel delivery line and filter connects to the manifold to the right of B.D.C. of the engine. Fuel is distributed to each of the 24 fuel spray nozzles through12 off equally spaced secondary manifolds. Each secondary manifold delivers fuel to 2 off fuel spray nozzles.

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ENGINE FUEL AND CONTROL GENERAL

Figure 123 FRA US-T TH NOV 99

Fuel Spray Nozzle and Manifold Page: 247

A 330 RR Trent 700

73-00 FUEL SPRAY NOZZLES 24 off fuel spray nozzles (F.S.N’s) are used on the Trent engine. They are cast body fabrications of simplex air spray design. Fuel is delivered to the F.S.N. then through the body (feed arm) to the swirl chamber head for atomisation and air mix before entry into the combustor. A weight type distributor valve is fitted inside the feed arm used to control the individual fuel delivery pressure, to match all the F.S.N’s output during low flow conditions i.e. engine start, and decent. The fuel enters the swirl chamber and is partially atomised and centrifuged by the tangential entry ports, H.P. delivery air passes into the rear of the swirl chamber mixing with the fuel, the air/fuel is swirled by a series of vanes before exiting the swirl chamber.

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FRA US-T TH NOV 99

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ENGINE FUEL AND CONTROL GENERAL

Figure 124 FRA US-T TH NOV 99

Fuel Spray Nozzle Page: 249

A 330 RR Trent 700

73-00 FUEL SPRAY NOZZLE REMOVAL/INSTALLATION If a fuel spray nozzle has to be changed use the distributer weight from the nozzle removed. Fit new distributer seal to the distributer weight and put it into the position in the fuel nozzle. If the same nozzle is being replaced make sure the C-seal is visually satisfactory. NOTE: After nozzle change ensure the proper location by borescoping the combustion chamber.

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FRA US-T TH NOV 99

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ENGINE FUEL AND CONTROL GENERAL

Figure 125 FRA US-T TH NOV 99

Fuel Spray Nozzle Removal / Installation Page: 251

Lufthansa Technical Training

ENGINE FUEL AND CONTROL GENERAL

A 330 RR Trent 700

73-00 Fuel Temperature Thermocouples Two thermocouples are installed in the flow pipe between F.M.U. and fuel flow transmitter, supplying temperature input to the E.E.C. i.e. Lane ’A’ and Lane ’B’. The fuel temperature thermocouples are installled in the tube that is connected to the outlet of the FMU.

For Training Purposes Only

They are attached to a tube adapter with bolts and are electrically connected through engine electrical harnesses to the EEC. One thermocouple supplies a temperature input for Channel A of the EEC. The other supplies a temperature input for Channel B of the EEC. Each thermocouple has a stainless steel case which contains a temperature sensitive element.

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ENGINE FUEL AND CONTROL GENERAL

Figure 126 FRA US-T TH NOV 99

Fuel Temperature Thermocouples Page: 253

A 330 RR Trent 700

73-00 L.P. FUEL FILTER REMOVAL / INSTALLATION CAUTION You must prevent the movement of the drain plug adapter when you remove the drain plug. If you do not do this you can loosen the adapter and cause fuel leaks. S Position a container under the drain plug and remove drain plug. S Remove, cut and discard used sealing ring. S With the container still in position remove the 3 bolts holding end cap in position, allow fuel to drain. S Remove cut and discard sealing ring. S Remove and discard filter element. S Fit blank over the hole. Installation is as follows: S Fit new element,make sure filters bonded seal goes into its location in its housing. S Fit new sealing ring to cap assembly. S Fit cap to housing ensuring filter bonded seal goes into its location in the cap assembly. S Fit bolts and torque load to 200 to 220 lbf/in (2,25 to 2,48 MdaN). S Fit new seal to drain plug, fit plug, torque load to 110 to 120 lbf/in (1,24 to 1,35 MdaN) and wire lock plug. S Carry out test shown in the MM for leaks.

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FRA US-T TH NOV 99

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ENGINE FUEL AND CONTROL GENERAL

Figure 127 FRA US-T TH NOV 99

LP Fuel Filter Page: 255

A 330 RR Trent 700

73-00 H.P. FUEL FILTER REMOVAL/INSTALLATION CAUTION You must make sure the retaining bolt is correctly installed in the filter element. If the bolt becomes disconnected from the element it can cause blockage in the fuel supply. S Open C–ducts. S Position a containe runder the drain shown in the diagram below. S Remove drain plug. S Remove, cut and discard sealing ring. S Fit new sealing ring and refit drain plug torque load to 44 lbf/in (0,5 MdaN) and wire lock plug. S Remove bolts from inlet and outlet connectors. S Remove filter body, cut and discard seals. S Release bolt and remove filter element from filter body. S Fit blanks/covers to all openings. Installation is as follows: S Make sure retaining ring is fitted correctly to hold retaining bolt to the filter element. S Fit element into filter body tighten retaining bolt and torque load to 100 lbf/in (1,13 MdaN S Fitnew sealing rings to inlet and outlet connectors. S Position filter body between the two connectors. S Fit bolt and torque load to 100 lbf/in (1,13 MdaN). S Carry out test shown in MM for leaks.

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ENGINE FUEL AND CONTROL GENERAL

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ENGINE FUEL AND CONTROL GENERAL

Figure 128 FRA US-T TH NOV 99

HP Fuel Filter Page: 257

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

ATA 75

AIR

75-33

IP/HP COMPRESSOR AIRFLOW CONTROL

A 330 RR Trent 700

75-00

For Training Purposes Only

GENERAL The function of the IP and HP compressor airftow control system is to keep a smooth airftow through the IP and HP compressor. It also controls the votume of airftow through the IP and HP compressors. The system makes sure of the correct operation of the compressors durlng all ranges of operation. The IP and HP compressor airftow control system has one stage of Variable Inlet Guide Vanes (VIGVs) and two stages of IP compressor Variable Stator Vanes (VSVs). The VIGVs and VSVs control the angle at whlch the alrflow ls supplied to the first three stages of the IP compressor. The angle of the VIGVs and VSVs is changed to adapt to different conditions of compressor operation. This helps to prevent a statl /surge condition in the IP and HP compressors. The volume of airflow through the IP and HP compressors is controlled by four IP and three HP bleed vatves. At lower engine speeds the bleed valves bleed air from the IP and HP compressors to prevent a stall /surge condition. The bleed valves are closed at higher engine speeds to provide full airflow through the IP and HP compressors. The IP and HP compressor airflow control system includes: S a VSV control unit S two VSV actuators S a VIGV/VSV actuating mechanism S a bleed valve controller S four IP bleed valves S three HP bleed valves

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 129 FRA US-T TH NOV 99

VIGV / VSV Actuator Page: 259

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A 330 RR Trent 700

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VIGV / VSV SYSTEM DESCRIPTION The VIGV’s and VSV’s are adjusted during acceleration, deceleration and surge conditions. This makes sure of correct operation of the IP and HP compressors. The EEC uses N2 speed signals and IP compressor temperature signals to control the position of the VIGV’s and VSV’s. If these signals are not available the EEC uses signals based on a pressure ratio to control the VIGV’s and VSV’s. The VIGV / VSV system consists of the following units: S VIGV / VSV Control Unit S Two VIGV / VSV Actuators S VIGV / VSV Actuating Mechanism The E.E.C. is constantly monitoring the speed and the inlet pressure of the I.P. compressor when these conditions change during acceleration and deceleration the E.E.C. will send a signal to the V.I.G.V./V.S.V. control unit. The control unit responds by directing H.P. fuel to the actuators to either retract or extend the rams. The V.I.G.V/V.S.V. actuating mechanism changes the linear movement of each of the actuator rams to a movement that turns the VIGV / VSV. Linear Variable Differential Transducers (L.V.D.T.) send signals back to the E.E.C. of V.I.G.V./V.S.V. angle. The left actuator LVDT feedback is send to EEC channel A, the right actuator LVDT feedback is send to EEC channel B.

For Training Purposes Only

STARTING During an engine start the VIGV’s and VSV’s are held in the closed position until 8% N3. ENGINE ACCELERATION As the engine speed increases the VIGV’s and VSV’s start to move to their open position. ENGINE DECELERATION As the engine speed decreases the VIGV’s and VSV’s start to move to their closed position.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 130 FRA US-T TH NOV 99

VIGV / VSV System Schematic Page: 261

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

A 330 RR Trent 700

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V.I.G.V./V.S.V. OPERATION DESCRIPTION The V.I.G.V./V.S.V. control unit is fitted on the lower left hand side of the compressor intermediate case, the unit consists of the following: S A Constant Pressure Valve S A Torque Motor S A Pressure Drop Regulator S A Control Servo Valve The diagram below shows the actuators on the high speed stops therefore V.I.G.V’s/V.S.V’s are fully open, there is no signal coming from the E.E.C. to the torque motor it remains in a neutral position. In this position it can be seen that the control servo valve is covering the outlet ports to the actuators and there is a hydraulic lock across the piston. Fuel from the fuel pumps passes through a constant pressure valve which maintains the supply pressure to the torque motor and the return pressure from the torque motor at a constant pressure drop. The torque motor flapper valve controls the flow of servo pressure (extend and retract pressure) to the control servo valve. The torque motor flapper valve is controlled by electrical signals from the E.E.C. There are two signals from the E.E.C. to two coils, only one signal is used at any one time. Energising one of the coils can move the flapper in two directions, the direction is dependent upon the E.E.C. signal. When the flapper moves closer to one nozzle and away from the other nozzle. This causes an out of balance condition in the hydraulic circuit. The flapper valve near to a nozzle decreases the fuel flow from that nozzle to the L.P. return.This increases the servo pressure at one end of the control servo valve. Movement of the flapper valve away from the nozzle increases the flow of fuel from that nozzle to L.P. return. This decreases the servo pressure at one end of the servo control valve.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 131 FRA US-T TH NOV 99

VIGV / VSV Actuation Schematic Page: 263

A 330 RR Trent 700

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VIGC /VSV OPERATION CONTINUED It can be seen from the diagram that movement of the servo control valve piston in either direction will supply HP regulated pressure to either side of the actuator piston to: S retract the actuator and open the VIGV / VSV S extend the actuator ans close the VIGV / VSV. A constant pressure drop is maintained across the servo control valve by the pressure drop regulator When the VIGV / VSV reach the required position as determined by the EEC it will send a signal to the torque motor to move the flapper valve to its neutral position. This provides the same pressure to each end of the servo control valve. The springs then put the piston into a neutral position where both outlet ports to the actuator are closed and hydraulically lock both actuators in position. LVDT’s send signals to the EEC which gives an indication of the position of the piston. The EEC uses this indication to control piston movement. Both LVDT’s are energised but only one is used for the control of piston movement.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 132 FRA US-T TH NOV 99

VSV Control Unit Page: 265

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CMS VARIABLE STATOR VANE TEST This test carried out by EEC and done through the MCDU enables the maintenance to perform a active VSV test. Therefore perform all necessary safety precautions prior engine motoring. The EEC performs a full travel check from the VIGV / VSV Actuator and checks the given feedback signals from the system while the engine is nmotored.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 133 FRA US-T TH NOV 99

CMS VSV Test Page: 267

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

A 330 RR Trent 700

75-00

COMPRESSOR BLEED CONTROL SYSTEM GENERAL The compressor bleed control system ensures that adequate surge margin are maintained in the intermediate and high pressure compressors when the engine is operating at lower rpm ranges. To maintain a stable airflow during certain transient and steady state running conditions a percentage of air is vented from the IP and HP compressors. This is accomplished with 7 bleed valves.There are three bleed valves located on the HP3 stage, and four valves on the IP8 stage. All valves discharge to the fan discharge ducting. The EEC controls the bleed valves by means of a solenoid pack.

For Training Purposes Only

OPERATION The EEC uses IP compressor shaft speed ( N2 ) signals and IP compressor temperature signals to control the IP bleed valves. These signals control the open/closed position of the IP bleed valves. The EEC uses HP compressor shaft speed ( N3 ) and HP compressor temperature signals to control the HP bleed valves. These signals control the open/ closed position of the HP bleed valves. If these signals are not available the EEC uses signals based on a pressure ratio. The EEC can also use signals from the throttle resolver position angle to set each bleed valve.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 134 FRA US-T TH NOV 99

IP and HP Bleed Valves Page: 269

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

A 330 RR Trent 700

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COMPRESSOR BLEED VALVES IP BLEED VALVES The 4 IP Bleed valves are installed on and around the compressor intermediate case. These bleed valves are aligned with stage 8 of the IP Compressor. They operate to bleed a amount of stage 8 air into the LP Compressor airflow at low engine speeds and during an engine surge or stall. Each IP bleed valve has a body, piston and a spring. The piston and spring are installed in the body. Movement of the piston opens and closes the bleed valve. When the engine is not in operation the spring pressure holds the valve in the open position. This gives the correct airflow through the IP compressor for engine start. HP BLEED VALVES The 3 HP Bleed valves are installed near to the front of the combustion outer case. Two HP Bleed valves are installed at the top right and bottom right of the case. The other HP Bleed valve is installed at the bottom left of the case. The bleed valves are aligned with stage 3 of the HP compressor. They operate to bleed a amount of HP stage 3 air into the LP Compressor airflow at low engine speeds and during an engine surge or stall. Each HP bleed valve has a body, two springs, piston, stem and valve. The piston stem and valve are assembled together to make a valve assembly. This valve assembly is sealed against the valve body with two seal assemblies. Movement of the valve assembly opens ore closes the bleed valve. When the engine is not in operation the spring pressure holds the valve assembly in the open position. This gives the correct airflow through the HP compressor for engine start condition.

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IP BLEED VALVES

HP BLEED VALVES

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 135 FRA US-T TH NOV 99

IP and HP Bleed Valves Page: 271

A 330 RR Trent 700

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DESCRIPTION CONTINUED The four I.P. compressor stage 8 and the three H.P. compressor stage 3 bleed valves are controlled by five solenoid valves contained in one unit. Two solenoid valves operate the four I.P. bleed valves in pairs. One of the solenoid valves operates the I.P.bleed valves fitted at the top right and bottom left of the I.P. compressor intermediate case. The remaining three solenoid valves each operate one of the three H.P. bleed valves fitted to the combustion chamber outer case. H.P. compressor stage 3 air is supplied to each solenoid valve when the solenoid valve is energised by the E.E.C. it vents H.P. 3 air servo pressure, which is keeping the bleed valve open, to atmosphere. This allows bleed air from the compressor to close the bleed valve(s). When the solenoid is not energised H.P. 3 air servo pressure is supplied from the solenoid valve(s). This H.P. 3 servo pressure combined with a spring in the bleed valve(s) holds the valve in the open position and I.P. and H.P. compressor air flows into the by–pass casing.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 136 FRA US-T TH NOV 99

Bleed Valve System Page: 273

A 330 RR Trent 700

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SOLENOID PACK The I.P. and H.P. solenoid valves are attached together to form one unit (Pack). This pack is mounted on the rear flange of the I.P. compressor V.S.V. case, it is on the right hand side above the horizontal centre line of the engine. There is one pneumatic connector and two electrical connectors, these supply electrical power and air to the five solenoids. Each solenoid has two coils, one coil is connected to the E.E.C. lane ’A’. The other coil is connected to lane ’B’.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 137 FRA US-T TH NOV 99

Bleed Valve Solenoid Pack Page: 275

A 330 RR Trent 700

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H.P. 3 AND I.P. 8 BLEED VALVES S Bleed Valve open The schematic diagram below shows one of the H.P. 3 bleed valves in the open position. When the coils in the solenoid valve are not energised springs move the vent valve to close vent ’A’. This allows H.P. 3 air into chamber ’A’ moving the piston to the right. The piston moves the inlet valve against the spring. The piston has also closed vent ’B’ and opened the inlet valve. H.P. 3 air can now flow through the opened inlet valve into the H.P. 3 servo air tube to the bleed valve servo chamber. In this condition the valve is open. Although the diagram shows an H.P. bleed valve the solenoid control for opening the I.P. bleed valve is the same.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 138 FRA US-T TH NOV 99

Bleed Valve Open Page: 277

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H.P. 3 AND I.P. BLEED VALVES S Bleed Valve closed When the coils in the solenoid are energised the vent valve moves to the left. This action allows H.P. 3 air in chamber ’A’ to vent through vent ’A’. H.P. 3 air inlet pressure combined with spring pressure in chamber ’B’ moves the piston to the left. H.P. 3 servo air then vents through vent ’B’. This causes a reduction in pressure in the bleed valve servo chamber. It is the air that is being bleed from the compressors that closes the bleed valve. The E.E.C. uses I.P. compressor shaft speed and air inlet temperature signals to control the I.P. bleed valves. The E.E.C. uses H.P. compressor shaft speed and air inlet temperature signals to control the H.P. bleed valves. If these signals are not available the E.E.C. uses signals based on pressure ratio. The E.E.C. can also use signals from the throttle resolver angle (T.R.A.) to set each bleed valve. If electrical failure occurs the bleed valve’s are moved to the open position.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 139 FRA US-T TH NOV 99

Bleed Valve Closed Page: 279

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

75-20

A 330 RR Trent 700

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TURBINE CASE COOLING

INTRODUCTION GENERAL The system improvesengine performance by decreasing the I.P. turbine blade tip clearance in the stable cruise mode. At all other flight conditions the system allows the I.P. turbine blade tip clearance to increase. A secondary function is to keep the I.P. / L.P. turbine casing temperatures within satisfactory limits. This is achieved by directing a controlled flow of cooling to the outside of the I.P. turbine casing allowing the casing to contract decreassing the blade tip clearance, however by shutting off the air supply the casing will expand increasing blade tip clearance.

NOTE: When the valve is closed there is a small clearance, this allows a minimum supply of air to the I.P. / L.P. casing for satifactory cooling.

OPERATION The diagram below shows the main components in the TCC system: S TCC manifold S TCC actuator S TCC air manifold S TCC liner assembly The diagram shows the valve in the fuuly open position which is the mode in stable cruise conditions where it is necessary to decrease the tip clearance of the I.P. turbine blades. When specified flight conditions occur the EEC transmits a signal to the solenoid, this allows HP3 air to go through and enter the actuator and the piston extends to open the butterfly valve. A much larger quantity of L.P. compressor air will flow through the valve and around the manifold. The air flows through two rows of holes onto the I.P. turbine casing, this causes a decrease in casing temperature and a subsequent decrease in I.P. tip clearance. The air also flows through the L.P. turbine casing cooling liner assembly to decrease the temperature of the L.P. turbine casing. At all other flight conditions the EEC de-energises the solenoid and cuts off the supply of H.P. 3 air to the actuator. The actuator spring now causes the piston to retract and close the butterfly valve. The flow of cooling air in the manifold is now greatly reduced and thus the I.P. turbine tip clearance now increases. If there is an electrical failure the spring closes the butterfly valve. FRA US-T TH NOV 99

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Figure 140 FRA US-T TH NOV 99

Turbine Case Cooling Schematic Page: 281

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A 330 RR Trent 700

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TCC ACTUATOR AND VALVE GENERAL The actuator is located on a bracket to the right of the H.P. / I.P. turbine casing near the engines horizontal centre line. A flexible air tube connects to the actuator to provide the flow of H.P. 3 air. The link assembly is adjustable therefore its lenght can be adjusted to make sure the valve operates correctly.

TCC COOLING SOLENOID

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GENERAL The solenoid is located at the bottom of the intermediate casing. When the solenoid is commanded closed by either of the EEC lanes, air in the actuator is vented to atmosphere through the outlet hole.

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TCC ACTUATOR / VALVE

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TCC COOLING SOLENOID

Figure 141 FRA US-T TH NOV 99

TCC Actuator / Valve and Solenoid Page: 283

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TCC MANIFOLD AND COOLING LINER GENERAL The manifold assembly is in the shape of a box section. It is an assembly of three parts as shown in the diagram. Two rows of holes are drilled at equal distance around the inner surface of the manifold. The cooling air will flow through these holes directly onto the I.P. turbine casing. A rearward projection of the manifold inner surface forms the L.P. casing cooling liner. The liner assembly is in four parts, bolted to brackets on the L.P. turbine casing. There are seals between each section and the front edge is sealed to the manifold shroud. The contours of this shroud provide locations for the EGT thermocouples and overheat detector switches. The bolts that hold the thermocouples to the turbine casing also hold the manifold assembly in position.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 142 FRA US-T TH NOV 99

TCC Manifold and Cooling Liner Page: 285

A 330 RR Trent 700

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ENGINE COOLING AND SEALING SYSTEM GENERAL The engine is internally cooled with air supplied by the I.P. and H.P. compressors.This air is also used to seal bearing chambers to prevent internal leakage of oil. S Air which is supplied by the I.P. compressor is bled off at stages IP 5 and IP 8 S Air which is supplied by the H.P. compressor is bled off at stages H.P. 3 and H.P. 6. Parts of the engine which are at different pressures are isolated from each other by labyrinth seals. The temperature of the cooling air around the I.P. turbine disc is monitored by the turbine overheat detection system. The bearing compartment of the internal gearbox is usually kept cool by I.P. compressor air. But during hot day take–off conditions this is not sufficient to keep the gear box at a satisfactory temperature. More cool air is required and this is supplied by a different source, the bearing compartment cooling system. Stage 3 air is taken from the H.P. compressor and made cool by L.P. air in a heat exchanger. This air is supplied to the internal gearbox. Internal gearbox temperature is monitored by the E.E.C. and the system operates automatically.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 143 FRA US-T TH NOV 99

Cooling and Sealing Airflows Page: 287

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

A 330 RR Trent 700

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BEARING COMPARTMENT COOLING SYSTEM This system is contained in an Air Cooled Air Cooler (A.C.A.C.) which includes the following component parts: S Heat exchanger S A rotary valve assembly S A solenoid valve S An actuator The A.C.A.C. is mounted on the right hand side of the core engine between the H.P. bleed valves.

For Training Purposes Only

HOT DAY CONDITIONS During hot day take off and climb conditions it is necessary to supply more cool air to the internal gearbox. When such conditions occur the E.E.C. will send a signal to de–energise the solenoid. This stops H.P.3 air supply to the piston head of the actuator by closing a ball valve and opening a vent from the actuator. H.P.3 air can now move the piston which in turn opens the rotary valve. This allows L.P. compressor air to flow through the heat exchanger where the H.P.3 air is made cool. The cool H.P.3 air flowing through the rotary valve passes to the internal gearbox. When the rotary valve is fully open cams move to operate micro switches. This transmits a signal to the E.E.C. to tell it the rotary valve is open.

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Figure 144 FRA US-T TH NOV 99

ACAC Open Position Page: 289

A 330 RR Trent 700

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BEARING COMPARTMENT COOLING SYSTEM Cold day conditions When conditions monitored by the E.E.C. are such that additional cooling air is not required the solenoid is automatically energised. This closes the actuator vent and also lets the H.P.3 air supply pass through the ball valve to the head of the actuator piston. This being the larger surface area than the ram side there is a greater force which moves the piston to close the rotary valve. This action will stop the flow of H.P.3 air to the internal gearbox and L.P. compressor air through the heat exchanger. The cams move away from the micro switches and give the E.E.C. an indication that the rotary valve is closed. In the event of an electrical failure the solenoid will be in a de–energised condition. Therefore more cool air will be supplied to the internal gearbox during all conditions. To keep the solenoid cool throughout all operating conditions there is a constant cooling flow around the solenoid.

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Figure 145 FRA US-T TH NOV 99

ACAC Closed Position Page: 291

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COMPONENT DESCRIPTION AND LOCATION A.C.A.C. This diagram shows the location of the A.C.A.C. and its component parts. The H.P. 3 air ’in’ tube is installed between the combustion outer case and the inlet to the heat exchanger. The H.P.3 air tube is connected to a 4 way connector block. From this block 3 different tubes go to three intermediate case vanes to supply air to the internal gearbox. The L.P. compressor air ’in’ duct aligns with an opening in the inner structure of the ’C’ duct. The L.P. air ’out’ duct takes the used air to the lower bifurcation and is dumped overboard.

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

Figure 146 FRA US-T TH NOV 99

Air Cooled Air Cooler Page: 293

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AIR COMPRESSOR AIRFLOW CONTROL SYSTEM

A 330 RR Trent 700

75-00

TURBINE OVERHEAT DETECTION SYSTEM GENERAL The Turbine overheat detection system is used to monitor the IP-turbine for overheat. The IP turbine could soften causing blade root weakening and uncontained multiple blade release. This system provides an ECAM warning message in case of IP turbine overheat and forces the flight crew to shut down the engine. DESCRIPTION The system consists of two thermocouple assemblies, each thermocouple assembly is a dual unit. One thermocouple of each unit is wired to the EEC channel A, and the other thermocouples of each unit is wired to the EEC channel B. In the event that both front and rear thermocouples sense an overheat, the EEC outputs a signal to the aircraft via an ARINC 429 output bus.

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OPERATIOIN There are two alert levels based on time: S Front thermocouple – 60 sec alert - 677_ C – 5 sec alert - 802_ C S Rear thermocouple – 60 sec alert - 662_ C – 5 sec alert - 802_ C If both thermocouples in an assembly become unserviceable then the warning system will operate.

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Figure 147 FRA US-T TH NOV 99

Turbine Overheat Detection Probes Page: 295

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ACCESSORY COOLING

GENERAL I NTRODUCT I ON The powerplant is divided into three primary fire–resistant zones isolated from each other by fireproof bulkheads and seals. Calibrated airf lows are supplied to the zones to keep the temperature around the powerplant to an acceptable level.These airflows also provide a ventilation function to prevent the accumulation of hazardous vapours. An equally important fire resistant zone is the electronic unit protection box and it protects the following items: S Engine electronic controller (E.E.C.) S Power control unit (P.C.U.) S Overspeed protection unit (O.P.U.)

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Figure 148 FRA US-T TH NOV 99

Fireproof Bulkheads Page: 297

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AIR ACCESSORY COOLING

A 330 RR Trent 700

75-00 SYSTEM DESCRIPTION ZONE 1 Zone 1 is the annular space between the low pressure (L.P.) compressor case and the fan cowl doors. The zone runs longitudinally from a fireproof bulkhead at the rear of the nose cowl and the firewall. This zone houses most of the fuel and oil accessories and is ventilated by ram air ducted through the air intake cowl and is exhausted through an opening in the lower part of the right hand hinge cowl. If the zone pressure exceeds a pre–determined value then a pressure relief door will open. This door is located in the lower part of the left hand hinge cowl.

also two pressure relief doors, in the event of air pressure increasing above a pre–determined value.

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ZONE 2 Zone 2 is the annular space between the intermediate pressure (I .P.) compressor and the gas generator fairings. There are six of these fairings which are removable. This zone contains the actuators for the variable inlet guidevanes (V.I.G.V’s) and variable stator vanes (V.S.V’s) and related fuel supply tubes. It also includes oil supply and scavenge tubes. Air enters the zone through two holes at the top rear of the zone. It flows around the zone to decrease the temperature of the components, and to prevent the collection of fumes in the area. The air exhausts into the by–pass casing through four holes in the front of the gas generator fairings. The holes are big enough to ensure a satisfactory flow of air through the zone. ZONE 3 Zone 3 is the annular space between the gas generator and the inner surface of the thrust reverser ’C’ ducts. The inner surface of the zone includes the combustion and turbine cases. Separation of zone 1 and 3 air is catered for by seals on the ’C’ duct mating surfaces. Zone 3 is known as the hot zone which contains some of the hydraulic components which operate the thrust reverser pivot doors. It also contains oil supply and scavenge tubes, fuel manifold and drain tubes. The airflow comes from the L.P. compressor entering the zone through ducts in the inner surface of the ’C’ ducts. It flows through the zone to decrease the temperature of the components and prevent the collection of fumes in the area. In the bottom of the ’C’ duct longitudinal beam is an exit for the air, there is

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Figure 149 FRA US-T TH NOV 99

Accessory cooling Zones Page: 299

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AIR ACCESSORY COOLING

A 330 RR Trent 700

75-00 NACELLE TEMPERATURE INDICATION GENERAL Attached to the bottom of the engine at 06:00 position there is a thermocouple probe which monitors the nacelle air temperature. If there is a sudden incrase of this temperature there will be an electrical input to the EEC channel A. The EEC then transmit an indication to the cockpit. DESCRIPTION S The temperature indication system includes a thermocouple probe and an electrical harness which connects it to the EEC. The probe body has a flange to install the unit, and an receptacle for connection of the harness. To the rear the probe extends from the body as a longer sensor tube. This unit is a assembly and cannot disassembled. S At the bottom of the core engine on the front flange of the combustion outer case there is a bracket. The probe is attached to this bracket by two nuts and bolts, and the sensor extends rearwards across the zone.This part is held in position by two clamps. The electrical harness is connected at the receptacle and it goes forward to the EEC.

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INTERFACE The analog signal from the nacelle temperature thermocouple is received by the EEC, trimmed down by a coefficient of 0,788 and digitalized by the EEC and sent as EEC ARINC output to the DMC for display on the SD, and to the FWC for warning activation. CONTROL AND INDICATION The nacelle temperature is indicated – in green colour in normal operation – in green pulsing colour if the temperature exceeds 260_ C ( advisory ) The indication above is replaced by the starter valve indication on engine start selection.

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NACELLE ADVISORY LIMIT

WHITE

GREY

GREEN

WHITE

SCALE MINIMUM VALUE

Figure 150 FRA US-T TH NOV 99

CYAN

SCALE MAXIMUM VALUE

Zone 3 Thermocouple Page: 301

A 330 RR Trent 700

75-00 C-DUCT SEALS The following diagram shows the location of the c-duct seals which seperate the zone 1 from the zone 3.

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AIR ACCESSORY COOLING

FRA US-T TH NOV 99

Page: 302

A 330 RR Trent 700

75-00

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AIR ACCESSORY COOLING

Figure 151 FRA US-T TH NOV 99

Fire Seals Page: 303

A 330 RR Trent 700

75-00 ELECTRONIC UNIT PROTECTION BOX This is a rectangular box which is positioned in zone 1. It contains the E.E.C. the power control unit (P.C.U.) andthe overspeed protection unit (O.P.U.) and protects these units from heat and flame. The box has a curve which aligns with the contour of the engine case and the hinged cowl.The box is made of titanium. When the lid is opened more than 75 _ it will automatically lock.The lid is held closed by 26 quick release fasteners. The routing of all electrical cables to the components is through fire/fume seals. The box is kept externally cool by the zone 1 airflow. Intake air also enters a tube which takes the air into the box as far as the O.P.U. This cooling air passes through the box and out through a tube in the lower wall. The tube goes down and forward across the L.P. compressor case and lets the air bleed back into the air intake through different outlets.

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AIR ACCESSORY COOLING

FRA US-T TH NOV 99

Page: 304

A 330 RR Trent 700

75-00

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AIR ACCESSORY COOLING

Figure 152 FRA US-T TH NOV 99

EEC Unit Cooling Page: 305

Lufthansa Technical Training

ENGINE ENGINE CONTROLS

A330 RR TRENT 700

76-00

ATA 76

ENGINE CONTROLS

76-00

GENERAL

The power control system consists of: – one throttle control system for thrust setting – two ENG / MASTER switches which control the HP and the LP fuel shutoff valves

THROTTLE CONTROL The throttle control system is fully electrical . It includes seperate throttlel control lever assemblies, one for each engine. Each throttle control lever drives one throttle control unit which indicates the Throttle Resolver Angle ( TRA ). The Throttle Control Unit is located under the pedestal.

For Training Purposes Only

ENGINE MASTER SWITCH The engine master switch controls: – the energization of the pressure raising and shutoff valve – the energization of the Lp fuel shutoff valve, which can also be controlled by the associated ENG FIRE pushbutton switch – the reset of the EEC channel A and B

FRA US-T TH NOV 99

Page: 306

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ENGINE ENGINE CONTROLS

A330 RR TRENT 700

76-00 THROTTLE CONTROL LEVER AUTOTHRUST INSTICTIVE DISCONNECT PUSHBUTTON

5

1

2

4

3 UPPER MECHANICAL RODS

0 0

0 IDLE STOP 1 MAX CLIMB ( MCL ) 2 MAX CONTINUOUS / FLEX T.O. / DERATED T.O.

For Training Purposes Only

3 MAX T.O. STOP ( MTO ) 4 REVERSE IDLE 5 MAX REVERSE STOP ARTIFICIAL FEEL UNIT

LOWER MECHANICAL RODS

RESOLVER

Figure 153 FRA US-T TH NOV 99

Throttle Control System Page: 307

Lufthansa Technical Training

ENGINE ENGINE CONTROLS

A330 RR TRENT 700

76-00 HP FUEL SHUTOFF VALVE The high pressure fuel shutoff valve is electrically controlled by a torque motor in the FMU It opens with fuel pressure coming from the fuel metering valve provided the torque motor is enrgized There is a second torque motor, overspeed torque motor, which can be activated by the OPU or LPTOS in case of engine overspeed. Note that the command from the master switch takes priority over the EEC.

LP FUEL SHUTOFF VALVE

For Training Purposes Only

The low pressure fuel shut off valve is normally controlled by the engine master switch. The low pressure fuel shutoff valve opens when the master switch slave relay is deenergized ( Master switch in ON ) and provided the ENG FIRE switch is not released out. The low pressure fuel shutoff valve closes when it receives a shut signal through the master switch slave relay by setting the master control switch to OFF. The low pressure fuel shutoff valve also closes when it receives a shutoff signal from the ENG FIRE pushbutton in the released out position.

FRA US-T TH NOV 99

Page: 308

A330 RR TRENT 700

76-00

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ENGINE ENGINE CONTROLS

Figure 154 FRA US-T TH NOV 99

HP and LP Fuel Shutoff Valve Page: 309

Lufthansa Technical Training

ENGINE ENGINE CONTROLS

A330 RR TRENT 700

76-00 ARTIFICIAL FEEL UNIT The throttle control artificial feel unit is a friction system which provides a load feedback to the throttle control levers.

TRA

DEFINITION MAX REVERSE STOP HIGH END OF MAX REV STOP

THROTTLE CONTROL UNIT The throttle control unit transforms a mechanical movement into electrical signals. Dual thrust lever resolvers send thrust lever position to each EEC channel.These analog signals are used by the EEC for thrust commands.

TLA

MANUAL MODE A/THR NOT ACTIVE

LOW END OF REVERSE IDLE FLAT REVERSE IDLE DETENT POINT HIGH END OF REVERSE IDLE FLAT IDLE STOP HIGH END OF IDLE FLAT

AUTOMATIC MODE A/THR ACTIVE

LOW EN D OF MCL FLAT MAX CLIMB DETENT POINT HIGH END OF MCL FLAT

LOW END OF MCT FLAT MAX CONT DETENT POINT HIGH END OF MCT FLAT

For Training Purposes Only

MANUAL MODE A/THR NOT ACTIVE

LOW END OF TAKE OFF FLAT TAKE OFF DETENT POINT

FRA US-T TH NOV 99

Page: 310

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ENGINE ENGINE CONTROLS

A330 RR TRENT 700

76-00

CRANK

CASING

INPUT

ROD

ROCKER

LEVER ELECTRICAL CONNECTORS

THROTTLE RESOLVER

COMPRESSION

CAM GEAR

SCREW

ADJUSTABLE FRICTION

FRICTION

ASSEMBLY

ADJUSTMENT

For Training Purposes Only

SCEW

PTS 0

POS IDLE STOP

1

MAX CLIMB

2

MAX CONTS

3

MAX T.O.

4

REV IDLE

5

MAX REV

5

3

4

POTIS

2

POTIS

ELECTRICAL CONNECTORS

RESOLVER

1

Figure 155 FRA US-T TH NOV 99

PINION

Throttle Control / Artificial Feel Unit Page: 311

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

A 330 RR Trent 700

30-20

ATA 30

ICE AND RAIN PROTECTION

30-20

ENGINE AIR INTAKE ICE PROTECTION

GENERAL Ice may form on the leading edge of the nose cowl when the engine is operating in conditions of low temperature and high humidity. 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: S The Nose Cowl leading edge (Thermal) S The P20/T20 Probe (Thermal) S The Spinner (Dynamic) NOSE COWL ANTI–ICING This is achieved by ducting hot air from the 3rd stage of the H.P. compressor to a spray ring fitted in a ’D’ shaped chamber formed in the front of the nose cowl. Fitted in the ducting is a pressure regulating and shut off valve (P.R.S.O.V.). Used hot air is vented overboard through an exhaust grill located in the nose cowl outer skin lower left hand side.

For Training Purposes Only

SPINNER ANTI–ICING The spinner is protected from ice build up by a solid rubber nose tip which vibrates naturally to break up and dislodge the ice immediately it starts to form.

FRA US-F PT MAY 96

Page: 312

A 330 RR Trent 700

30-20

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

Figure 156 FRA US-F PT MAY 96

Anti Icing System Page: 313

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

A 330 RR Trent 700

30-20 ANTI–ICING SYSTEM OPERATION DESCRIPTION Air is taken from the third stage of the engine H.P. compressor. On selection of the anti–icing switch light to ’ON’ the P.R.S.O.V. opens and air passes through ducting to the spray ring. The discharge of air onto the lip skin is through 4 staggered rows of holes in the spray ring. The supply duct inside the nose cowl is completely enclosed by the exhaust duct. This design ensures that if the supply duct – in the nose cowl – or spray ring was to burst the airflow from the burst would discharge overboard with no adverse structural affect. Where the supply duct enters the nose cowl there is a venturi which acts as a flow restrictor during system operation. High and low pressure switches are provided downstream of the P.R.S.O.V. to indicate system malfunction. Selection of anti–icing is by pushing the ’ENGINE ANTI–ICE’ switch light located on the flight deck overhead panel. When the system is selected a white ’ON’ light is shown on the switch light and a green ’ENG–A–ICE’ message is displayed on the upper E.C.A.M. screen If the low pressure switch operation does not agree wIth the switch light selected position the ’FAULT’ warning and ’MASTER CAUT’ comes on, the failure is also shown on the upper E.C.A.M. screen as an amber warning. There is also an aural warning. The ’FAULT’ warning light comes o during valve transit from on to off and vice versa. If pressure downstream of the P.R.S.O.V. becomes excessive the high pressure switch will function, this operation gives indication to the flight deck and produces a maintenance message. The complete system has been designed to permit aircraft despatch locked fully closed or fully open.

FRA US-F PT MAY 96

Page: 314

A 330 RR Trent 700

30-20

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

Figure 157 FRA US-F PT MAY 96

Anti Icing Diagram Page: 315

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

A 330 RR Trent 700

30-20 OPERATION The anti–icing valve P.R.S.O.V. is of the ’butterfly’ type and has two functions. It acts as the system ON/OFF valve and when selected ON controls the maximum outlet pressure to 75 PSI. The schematic drawing of the valve shown below comprises: S Butterfly valve S A double headed diaphragm type piston S A pilot regulator valve – to control system pressure S High and low pressure switches S A filter S ON/OFF solenoid The valve is shown in the open position, this would be the position when anti– icing is selected ON by the flight crew. The solenoid is de–energised and a spring has opened the vent from the upper chamber. Pressure upstream of the butterfly valve passes through the filter to the upper and middle chambers. It can be seen the middle chamber pressure combined with the spring has a greater force than the upper chamber, therefore the val”e is open. During operation if the pressure downstream of the butterfly valve exceeds the design system maximum pressure, that pressure acts on the piston of the pilot regulating valve. The piston moves to the right overcoming the spring force (this has a value set to control downstream maximum pressure) and increases the size of the variable restrictor orifice. This has the effect of reducing the pressure in the middle chamber, the force in the upper chamber is now greater than the combined middle chamber and spring force and the piston moves down closing the butterfly valve which reduces downstream pressure within the design limits. Before start up with no airflow in the system the butterfly valve remains open by the action of the spring. As soon as a start is initiated the solenoid is energised and closes the vent. Air pressure in the upper chamber now has a force greater than the combined spring and middle chamber force, therefore the valve closes and will remain closed until selected ’ON’ by the flight crew, when the solenoid will be de–energised to open and vent the upper chamber. In the event of a power failure to the solenoid the butterfly valve will open.

FRA US-F PT MAY 96

Page: 316

A 330 RR Trent 700

30-20

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

Figure 158 FRA US-F PT MAY 96

Anti Ice Valve Schematic Page: 317

A 330 RR Trent 700

30-20 P.R.S.O.V. MANUAL OVERRIDE To prevent aircraft delays due to non–operation of the anti–icing valve, the valve can be locked in the fully open position or fully closed position. This is done by cutting the locking wire securing the locking pin, removing it and manually wrenching the hexagon selector to the closed or open position. Refit the locking pin and replace the wire locking.

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FRA US-F PT MAY 96

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ICE AND RAIN PROTECTION ENGINE AIR INTAKE ICE PROTECTION

A 330 RR Trent 700

30-20

For Training Purposes Only

LOCKING SCREW

Figure 159 FRA US-F PT MAY 96

Anti Ice Valve Page: 319

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ENGINE EXHAUST

A330 RR Trent 700

78-00

ATA 78

EXHAUST

78-00

THRUST REVERSER

INTRODUCTION

For Training Purposes Only

The thrust reverser assists the wheel brakes in decelerating the aircraft quickly and safely on landing. The thrust reverser is an integral part of the fan stream duct and comprises of two ’C’ shaped ducts hinged at the top to the aircraft pylon. The ’C’ ducts can be opened to provide access to the core engine. Actuation of the thrust reverser movements is initiated when the pilot has his throttle lever in the idle detent, and the thrust reverser lever which is mounted on the pilots throttle lever (piggy back) is moved to reverse thrust (deploy select) or vice versa (stow select). The thrust reverser pivot doors are operated hydraulically utilising the aircraft hydraulic system pressure. The E.E.C. controls the actuation of the thrust reverser doors via two control valves, the Isolation Valve (I.v.) and the Directional Control Valve (D.C.V.). Ground/air sensors ensure that the thrust reverser can only be operated on the ground. There is a stow switch and a Rotary Variable Transducer (R.V.T.) for each door to indicate to the E.E.C. the position of each pivot door. These signals are used for thrust reverser control and thrust scheduling purposes. The thrust reverser can be ·locked out’ in forward thrust position to allow aircraft despatch with an inoperative thrust reverser. THRUST REVERSER SYSTEM The picture below is showing a cross section of the ·C’ ducts showing the pivot doors in the forward and reverse thrust position. When the pivot doors are in reverse thrust position a moveable plate has moved to expose a deflector plate (kicker plate). These plates control the efflux of fan airflow while in reverser thrust. When the pivot door is in the stowed position the moveable plate aligns flush with the inner duct.

FRA US/T TH NOV 99

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ENGINE EXHAUST

Figure 160 FRA US/T TH NOV 99

Thrust Reverser Assembly Page: 321

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ENGINE EXHAUST

A330 RR Trent 700

78-00 THRUST REVERSER INDICATION The thrust reverser status indications for the stow and deploy are received by the EEC from switches, and transmitted via the EEC ARINC output bus to the DMC for display on the EWD and to the FWC for warning activation. Thrust Reverser Unlocked When reverse thrust is set the door locks are released in sequence: S tertiary locks S primary locks S secondary locks Release of the locks plus the subsequent movement of the pivoting doors operate the sotw switches. This signal to the cockpit is seen on the EWD when an amber REV indication is indicated.

For Training Purposes Only

Thrust Reverser Deployed When the pivoting doors are at 90 % of their fully deployed position, the RVDT‘s give a signal to the cockpit. This is seen on the EWD when the amber REV indication changes to green indication. Thrust Reverser Stowed When forward thrust is set the actuators retract and this turns the pivoting doors to the stowed position. This RVDT moving is seen on the EWD when the green REV indication changes to amber indication. When the pivoting doors are in their fully stowed position, the primary, secon dary and tertiary locks engage. This operates the stow switches and the amber REV indication disapperars on the EWD. Thus when all pivoting doors are correctly stowed no visual indications are displayed to the flight crew.

FRA US/T TH NOV 99

Page: 322

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ENGINE EXHAUST

A330 RR Trent 700

78-00 ISOLATION INHIBITION SYSTEM ISOLATION CONTROL

THRUST REVERSER

UNIT

REV

EEC CH A

FWC 1

REV

EEC CH B

DMC 1

FWC 2

DMC 2 DMC 3

For Training Purposes Only

REV

EPR

1.282

REV GREEN INDICATION WHEN REVERSER IS FULL DEPLOYED AMBER INDICATION WHEN AT LEAST ONE PIVOTING DOOR IS UNLOCKED AMBER INDICATION FLASHES FOR 9SEC IN FLIGHT AND UNSTOWED CONDITION

Figure 161 FRA US/T TH NOV 99

Thrust Reverser Control Indication Page: 323

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ENGINE EXHAUST

A330 RR Trent 700

78-00 SYSTEM OPERATION SELECTING REVERSE THRUST (DEPLOY) When the thrust reverser is selected a signal is sent to the E.E.C. The E.E.C. will have monitored the condition of the system, if satisfied it will send a signal to the Isolation Solenoid Valve (I.S.V.). On receipt of this signal the I.S.V. will open a valve to allow hydraulic pressure to pass to the Directional Control Valve (D.C.V.) where pressure will be directed to the stow side of the actuators to overstow, this enables the locks to be released when selected. When the throttle is at idle and the aircraft is on the ground and the E.E.C. has monitored the release of the electrically controlled locks (third locks) the deploy solenoid is energised and hydraulic pressure is directed to release the primary door locks. The primary locks are released in sequence and when the last lock is released hydraulic pressure is directed to the actuators via the D.C.V. This pressure releases the secondary locks – inside the actuators – to extend the rams and comence opening the pivot doors. Stow switches will function and send a signal to the E.E.C. to give an indication in the flight deck that the doors are not stowed. Rotary Variable Transformers (R.V.T.) detect movement of the doors towards full deployment and will send position signals to the E.E.C. If the doors do not deploy correctly the E.E.C. will signal the D.C.V. to auto restow. When the pivot doors reach the fully deployed position the R.V.T. via the E.E.C. will give visual indication in the flight deck that all four pivot doors have fully deployed. If the reverser will not deploy or auto restow the E.E.C. will not let the engine power go above reverse idle.

The stow switches send electrical signals to the E.E.C. to give an indication in the flight deck that the pivot doors are stowed.

CANCELLING REVERSE THRUST (STOW) When reverse thrust is cancelled a signal is sent to the E.E.C. The E.E.C. will have monitored the condition of the locking and actuating systems. If satisfactory it will send a signal to the I.S.V. opening a valve allowing hydraulic pressure to pass to the D.C.V. and be directed to retract the actuators and co~ence closing of the pivot doors. When the doors reach the stowed position the primary, secondary and third locks engage. During the engagement of the locks the doors will have been in an overstow condition and will remain there until the E.E.C. removes the electrical signal to the I.S.V. shutting off hydraulic pressure to the actuators. The pivot doors will now return to their normal stow position. FRA US/T TH NOV 99

Page: 324

A330 RR Trent 700

78-00

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ENGINE EXHAUST

Figure 162 FRA US/T TH NOV 99

Thrust Reverser Actuation Diagram Page: 325

A330 RR Trent 700

78-00 HYDRAULIC CONTROL OPERATION The hydraulic control system consists of two Line Replacement Units (L.R.U.’s). The I.S.V., fitted in the aircraft pylon, which isolates the thrust reverser hydraulic system from the aircraft supply when the thrust reverser is not in use i.e. aircraft in normal flight. The D.C.V., fitted on the ’C’ duct front frame, which directs pressure to the actuators to deploy or stow the pivot doors. The diagram below shows the system in the forward thrust position (Stow). I.S.V. de–energises shutting off the hydraulic supply to the D.C.V. The D.C.V. de–energises to the stow position. The actuator piston head and rod ends connected to return. Pivot door locks mechanically locked and vented to return. There is no pressure to the pressure switch therefore no flight deck indication.

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FRA US/T TH NOV 99

Page: 326

A330 RR Trent 700

78-00

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Figure 163 FRA US/T TH NOV 99

Hydraulic Control Fwd Thrust Position Page: 327

A330 RR Trent 700

78-00 HYDRAULIC OPERATION DEPLOY SELECTED When a deploy selection has been made in the flight deck and all the deploy conditions are met as previously explained, the I.S.V. and D.C.V. solenoids are energised. Hydraulic pressure is passed to the D.C.V. and is directed to the actuator piston rod end to overstow and to the pivot door locks to release the primary locks, the third locks are also released electrically. The pressure switch functions when hydraulic pressure has passed through the isolation valve and provides a signal to both channels of the E.E.C. The purpose of the switch is to assist in fault diagnosis. There will only be an indication in the flight deck when the isolation valves commanded position and the pressure switch position do not agree. At this point in the sequencing the actuator piston head end is still connected to return.

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FRA US/T TH NOV 99

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A330 RR Trent 700

78-00

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Figure 164 FRA US/T TH NOV 99

Hydraulic Control Commanding Deploy Page: 329

A330 RR Trent 700

78-00 DEPLOY SEOUENCE CONTINUED When the last of the four pivot door locks is released pressure is directed to the D.C.V. control valve chamber. As the pressure builds up it moves the D.C.V. control valve to the deploy position closing off the actuator piston head end to return and connecting it to pressure. This pressure releases the secondary locks inside the actuators and all four pivot doors move to the deployed position. Note: It can be seen from the diagram below that both head end and rod end are open up to pressure. However, the head end of the piston has the greater surface area which extends the actuators. The displaced oil from the rod end joins the pressure to the actuator head end.

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FRA US/T TH NOV 99

Page: 330

A330 RR Trent 700

78-00

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Figure 165 FRA US/T TH NOV 99

Hydraulic Control Deploy Position Page: 331

A330 RR Trent 700

78-00 PRIMARY LOCKS When the I.S.V. and D.C.V. solenoids are energised, hydraulic pressure is directed to the ”lock in port” of one of the primary locks. This pressure moves the release plunger which moves the latch lever and roller away from the lock hooks. Movement of the release piston movesa valve to allow pressure to the ”lock out port” to release the next lock in the sequence, when pressure reaches the fourth lock pressure from its ”lock out port” is sent to the D.C.V. and directed to the four actuators to deploy the pivot doors. As the pivot doors move the door pin mechanically moves the hook out of the way. The hook spring will keep the hook in the released position and also keep the latch lever and roller in the unlocked position even in the absence of pressure on the release piston. When a stow selection is made the D.C.V. solenoid is de–energised and removes the pressure from the ”lock in port” the piston spring moves the release plunger out of the way. The action of the pivot doors closing, mechanically moves the hook into the locked position, the latch spring moves the latch lever and roller into a position to keep the hook locked.

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FRA US/T TH NOV 99

Page: 332

A330 RR Trent 700

78-00

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Figure 166 FRA US/T TH NOV 99

Primary Lock Mechanism Page: 333

A330 RR Trent 700

78-00 ACTUATORS AND SECONDARY LOCKS OPERATION The actuators contain the secondary locks which consist of a tine lock attached to the piston head, a lock release sleeve, a spring and a manual lock release. The diagram below shows the secondary lock engaged by the spring keeping the release sleeve in a position preventing the tine lock releasing. When the thrust reverser is selected to deploy the D.C.V. sequences hydraulic pressure to piston rod end of the actuator, first this moves the actuator to overstow which will assist in releasing the tine lock. The next part of the sequencing is tha the D.C.V. directs hydraulic pressure to the piston head end of the actuator and moves the lock release sleeve this allows disengagement of the tine lock and hydraulic pressure extends the actuator deploying the pivot doors. At 95% of deployment snubbing takes place to slow the actuator down to prevent internal damage. When the thrust reverser is selected to stow the D.C.V. directs pressure to the piston rod end of the actuator, just before the stow position is reached the tine lock mechanically moves the lock release sleeve against its spring allowing the tine lock to latch into its locked position. The spring now moves the release sleeve back to prevent the tine locks moving from their locked position. A manual release is provided for manually opening the doors during maintenance.

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Page: 334

A330 RR Trent 700

78-00

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Figure 167 FRA US/T TH NOV 99

TR Actuator Commanding Extending Page: 335

A330 RR Trent 700

78-00 THRUST REVERSER SYSTEM PURPOSE OF DIAGRAM To show the interaction between the thrust reverser major components during deploy and stow.

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Page: 336

A330 RR Trent 700

78-00

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Figure 168 FRA US/T TH NOV 99

TR Actuation System Schematic Page: 337

A330 RR Trent 700

78-00 MAINTENANCE PRACTICES THRUST REVERSER DEACTIVATION/ACTIVATION This task explains the procedure for deactivating and activating the thrust reverser by inhibiting the isolation valve by the use of the manual inhibit lever. Refer back to the description and operation of this chapter to remind you of the system. The procedure is quite straight forward and is as follows:– S Open right hand fan cowl door. S Get access to the isolation control valve unit. S Remove the quick release pin. S Move lever to inhibited position. NOTE: By referring back to the diagram of the isolation valve in the description and operation section of this chapter it can be seen that the isolation valve is now mechanically held in the deactivation position and that an indication of the position of the lever is sent to both channels of the E.E.C. Activation is carried out as follows:– S Open right hand fan cowl door. S Get access to the isolation control valve unit. S Remove quick release pin. S Move lever to normal position. S Put quick release pin back into the isolation control unit to keep the inhibit lever in the normal (activated) position.

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Page: 338

A330 RR Trent 700

78-00

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Figure 169 FRA US/T TH NOV 99

TR Activation/ De–Activation Page: 339

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ENGINE EXHAUST

A330 RR Trent 700

78-00 MAINTENANCE PRACTICES THRUST REVERSER DEACTIVATION FOR FLIGHT This procedure allows you to fly with an unserviceable thrust reverser, local regulatory authorities may have to be consulted as to when this procedure can be used. WARNING: You must deactivate the thrust reverser before you do work on or around it. If you do not, the thrust reverser can operate accidentally and cause an injury and/or damage. S Open fan cowl doors. S Remove inhibition bolt attachment covers (5) from pivot doors. S Remove the 4 inhibition bolts from the keep position by removing the clamps and lock plates. S Fit bolts through the holes in each pivot door and into the inhibition bolt holes in the front frame. NOTE: The longest bolts are fitted into the lower pivot doors. S Tighten bolt until the doors are the same level as the structure around them. S Make the bolts safe, fitting the lock plates and screws. S Torque load the screws to lOO lbf/in (1,13 MdaN). S Fit the inhibition bolt attachment covers (5) to the bracket (19) with screws. S Close the clamps and lock them with the screws. The inhibition bolt heads are painted red for ease of identification. NOTE: This procedure must never be used to lock out one defective pivot door only.

FRA US/T TH NOV 99

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A330 RR Trent 700

78-00

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Figure 170 FRA US/T TH NOV 99

TR In–Flight Lock Out Page: 341

A330 RR Trent 700

78-00 MANUAL OPENING OF THE PIVOT DOORS S Deactivate the thrust reverser. S Remove the tertiary lock access panel. S Use a standard O.3125 in. spanner to release the lock mechanisms. S Hold the primary and tertiary locks in the unlock position. S Open the pivot door using hand pressure on the front and rear edges of the door. Release the primary and tertiary locks. S Install the collar HU87114 on the actuator.

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Page: 342

A330 RR Trent 700

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Figure 171 FRA US/T TH NOV 99

Primary, Secondary and Tertiary Locks Release Mechanism Page: 343

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ENGINE EXHAUST

A330 RR Trent 700

78-00 RVT REMOVAL/ INSTALLATION Removal Procedure S Deactivate the thrust reverser. S Remove the RVT access panel. S Disconnect the electrical connectors and fit blanking caps. S Remove the four screws and washers and remove the RVT.

For Training Purposes Only

Installation Procedure CAUTION: If you have installed a replacement RVT mounting plate, you must adjust it before you install the RVT S Install the RVT on the dowel and the mounting plate. S Lift the shaft lock then install the screws and washers and torque load to 35 lbf/in (0,4 mdaN). S Make sure the electrical connectors are clean before connecting. S Fit the access panel. S Activate the thrust reverser.

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Figure 172 FRA US/T TH NOV 99

Pivoting Door RVDT Page: 345

A330 RR Trent 700

78-00 THRUST REVERSER STOW SWITCH There are 4 stow switches installed, one for each pivoting door. They are mounted on the front frame of each reverser cowl door. The stow switches are the microswitch type. Each switch has got two identical electrical circuits that give the same signal at the same time. The switches are connected to a junction box. The junction box is connected to the EEC. When the pivoting doors are in the stowed position the switch lever is pushed in by the stop of the door. As the pivoting door turns to deployed position the lever is released.

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Figure 173 FRA US/T TH NOV 99

Pvoting Door Stow Switch Page: 347

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A330 RR Trent 700

78-00 TERTIARY LOCK REMOVAL/ REPLACEMENT Note: The procedure is the same for all four pivot doors. S Deactivate the thrust reverser. S The pivot door may be in the closed or open position. CAUTION: When you remove the shoot bolts from the solenoid assembly you must keep the shoot bolt with the thrust reverser.

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Removal Procedure – Pivot Door Closed S Remove the front access panel. S Disconnect the electrical connectors and fit blanking caps. S Remove the four nuts and disengage the solenoid assembly from its four studs on the fitting. S Pull the solenoid out until the shoot bolt is disengaged from its housing in the fitting. S Disengage the shoot bolt from the solenoid plunger. Removal Procedure – Pivot Door Open S Remove the front access panel S Disconnect the electrical connectors and fit blanking caps. S Turn the unlock shaft counterclockwise with a square wrench to unload the solenoid. Hold the unlock shaft in this position and remove the four nuts. S Disengage the solenoid assembly from its four studs and release the unlock shaft. S Pull the solenoid out until the shoot bolt is disengaged from its housing in the fitting. S Disengage the shoot bolt from the solenoid plunger. POWER CONDITIONING MODULE The Power Conditioning Module changes 115VAC supply from the aircraft electrical system to a 104 VDC supply for the solenoids of the tertiary lock system.

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Figure 174 FRA US/T TH NOV 99

Tertiary Lock/ Power Conditioning Module Page: 349

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A330 RR Trent 700

78-00 TERTIARY LOCK REMOVAL/INSTALLATION Installation Procedure – Pivot Door Closed CAUTION: You must make sure the shoot bolt is engaged with the plunger before you install the solenoid assembly. S Engage the shoot bolt with the plunger and position the solenoid assembly over the four studs and engage the shoot bolt in its housing in the fitting. S Engage the solenoid assembly on its four studs and install and torque load the nuts to l00lbf/in (1,13 mdaN). S Make sure electrical connectors are clean before connecting.

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Installation Procedure – Pivot Door Open S Engage the shoot bolt with the plunger and position the solenoid assembly over the four studs and engage the shoot bolt in its housing in the fitting. S Turn the unlock shaft counterclockwise to unload the solenoid. Hold the unlock shaft in this position and engage the solenoid assembly on its four studs install the nuts and torque load to l00 lbf/in (1,13 mdaN). S Release the unlock shaft. S Make sure electrical connectors are clean before connecting. S Install the front access panel and torque load the screws to 30 lbf/in (0,35 mdaN). S Close the pivot door. S Activate the thrust reverser

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Figure 175 FRA US/T TH NOV 99

Tertiary Lock Page: 351

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78-00 THRUST REVERSER CMS TEST During this test, hydraulic pressure is supplied to the Thrust Reverser System. The Thrust Reverser is deployed and stowed byy moving the throttle in the reverse and forward positions. The inhibition circuits and EIVMU relays are checked.

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The test can be aborted by selecting the return line select key. The test not completed screen appears when the operator does not follow the instructions in time.

THRUST REVERSER SAFETY SWITCH

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Figure 176 FRA US/T TH NOV 99

CMS Thrust Reverser Test 1 Page: 353

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Figure 177 FRA US/T TH NOV 99

CMS Thrust Reverser Test 2 Page: 354

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Figure 178 FRA US/T TH NOV 99

CMS Thrust Reverser Test 3 Page: 355

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ENGINE IGNITION SYSTEM

A330 RR Trent 700

74-00

ATA 74

IGNITION SYSTEM

74–00

IGNITION SYSTEM

INTRODUCTION The ignition distribution system is in two parts, each part has an ignition unit, an ignition lead and an igniter plug. The ignition units are mounted on bracket assemblies on the lower left hand side of the L.P. compressor case. The leads span the by–pass casing inside the lower bifurcation to the igniter plugs. The igniter plugs are adjacent numbers l0 and 16 fuel spray nozzles.

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IGNITION UNITS Each unit has a case assembly and an ignition exciter. An input of 115 volts 400 Hz A.C. power is supplied by the aircraft electrical system. This supply is transmitted to a relay in the Power Control Unit (P.C.U.) which is controlled by the E.E.C. The exciter is a capacitive discharge circuit. The exciter changes the input voltage to an output voltage of 2.7 to 2.9 K volts. Energy is stored in the ignition unit at 8 to 11.3 joules. This energy is released by the exciter at the rate of 60 to 135 sparks per minute.

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1 Emergency Bus

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2 Normal Bus

Figure 179 FRA US-T TH NOV 99

Ignition System Components Page: 357

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A330 RR Trent 700

74-00 IGNITER PLUGS The igniter plug is a surface discharge type. It has a body and a ground electrode, it also has a centre wire with a centre electrode at the tip. The centre wire is sealed with glass and has insulation along its length. The space between the centre electrode and the ground electrode is filled with a semi–conductive material. The igniter plug has a contact button which touches the contact button in the ignition lead.

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OPERATION When the ignition system is energised an electrical current flows through the centre wire and the centre electrode of the igniter plug. The current flows through the semi conductor to the ground electrode. This current produces a magnetic field which ionises gas near the igniter tip. This gives a low resistance path for the energy from the ignition unit and a pulse of energy occurs. The energy pulse gives a high energy spark from the centre electrode to the outer electrode. The electrical current flows through the igniter plug body and to the outer conductor of the ignition lead and on to the ignition unit case.

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Figure 180 FRA US-T TH NOV 99

Igniter Plug Page: 359

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ENGINE IGNITION SYSTEM

A330 RR Trent 700

74-00 IGNITER PLUG REMOVAL/INSTALLATION Before installation of the igniter plug use the protrusion gauge HU38434 as follows: Move the locking collar, locating plate and body up against the indicator and install the protrusion gauge in the applicable igniter plug hole. Make sure indicator notch points to rear of engine. Carefully move the indicator until probe foot touches igniter sleeve. Use the probe foot to feel if the igniter sleeve is in the correct position and has not become loose. If the igniter sleeve is not in the correct position or has become loose – reject the engine. CAUTION: Remove protrusion gauge carefully to prevent damage to combustion chamber coating. Installation of the igniter plug is straight forward but remember the following points: S Apply dry film lubricant to the threads of the igniter plug (OMat 4/20). S If the same igniter plug is being fitted in the same position, fit the same adjusting washers. S Use correct socket wrench UT1152 and torque load to between 200 and 300 lbf/in (2,26 to 3,39 MdaN). S Clean ignition lead contact button with OMat 5/43 emery paper, remove dust with a clean lint free cloth. S Connect the ignition lead and torque load to between 140 and 150 lbf/in (1,58 and 1,69 MdaN). Safety the ignition lead with OMat 238 lockwire. S If a different igniter plug is being installed then a depth of immersion check is required as follows:

Install protrusion gauge carefully to prevent damage to the combustion liner. S Move the locking collar, the locating plate and the body up against the indicator. S Install protrusion gauge into the applicable igniter plug location hole. Ensure indicator notch points to the rear of the engine. S Ensure the arrow on the locating plate points to the front of the engine and the locating plate aligns with one of the three bolts. S Turn indicator until the notch points to the front of the engine. S Carefully pull the indicator up until the probe foot touches the combustion liner. S Lightly tighten the locking collar. Measure dimension ’X’ again (Step 3) take the initial dimension ’X’ from the new dimension ’X’. The result is the thickness of adjustment washer(s) necessary to get the correct igniter immersion. S Fit the minimum number of adjustment washers to give a thickness of between 0.265 and 0.280 in. (6,72 and 7,12 mm). S c.Remove the protrusion gauge:– S Loosen the locking collar and push the indicator down, turn the indicator until the notch points to the rear of the engine and remove the protrusion gauge. S Install the igniter plug with the adjustment washers in the applicable igniter plug hole (see removal/ installation procedure). Note: You must always use the 0.13in (3,20mm) washer in the adjustment.

IGNITER PLUG DEPTH OF IMMERSION CHECK a.Set protrusion gauge HU38434 to the igniter plug depth (Step 1) as follows : S Remove any adjustment washers. S Put the igniter plug against protrusion gauge (Step 1) and measure dimension ’x’ with a depth gauge. b.Install protrusion gauge HU38434 to engine (Step 2) as follows: CAUTION: FRA US-T TH NOV 99

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Figure 181 FRA US-T TH NOV 99

Igniter Plug Depth of Immersion Check Page: 361

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74-00 CMS IGNITION TEST The engine ignition test is performed through the CMS in the cockpit. In order to perform the active ignition test consult the AMM for safety precautiones. The operational test is selected by the maintenance through the MCDU and carried out by the EEC. NOTE: Audibly verify if both ignitor plugs spark on the test sequence!

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Figure 182 FRA US-T TH NOV 99

CMS Ignition Test Page: 363

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ENGINE STARTING SYSTEM

A330 RR Trent 700

80-00

ATA 80

STARTING

80-00

STARTING

STARTING SYSTEM INTRODUCTION The engine starting system provides the power which turns the H.P. rotor to a speed at which an engine start can occur. The system comprises: S An air starter motor S A start control valve S Air ducting S Dual ignition system S Start control panels in the flight deck Air is used to turn a turbine in the starter motor which provides the torgue at the starter output shaft. The starter motor being fitted to the front face of the external gearbox turns the gears, and drives a drive shaft which spans the by–pass casing to the gas generator which will turn the H.P. rotor. To start the engine it is necessary to: S Rotate the engine to induce an airflow through the H.P. section. S Provide the correct quantity of fuel to the combustion chamber. S Ignite the resultant air/fuel mixture. Air to operate the air starter motor comes from: S A ground air supply. S The auxiliary power unit (A.P.U.). S The other running engine.

FRA US-T TH NOV 99

Ignition is provided by two ignition plugs which can be operated together or independently. The operation of one igniter plug is called single ignition, the operation of both igniters is called dual ignition. Single ignition is used for ground starts. Dual ignition is used for manual–starts, in–flight starts, auto– relight, and for continuous ignition. The E.E.C. controls the opening and closing of the start control valve and the electrical supply to the ignition units.

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Figure 183 FRA US-T TH NOV 99

Starting System Overview Page: 365

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80-00 STARTER DUCT The starter air duct flanges are connected together by ’v’ band coupling clamps. Air leakage is prevented by the ’E’ type seals that are located between the mating flanges. There are two flexible joints which let the engine move, in relation to the aircraft pylon, without damage to the ducts. These flexible joints also help align the pylon duct with the aircraft duct.

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Figure 184 FRA US-T TH NOV 99

Starter Motor and Air Duct Page: 367

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80-00 AIR STARTER MOTOR The air starter motors primary components consist of: S A turbine rotor S A reduction gear configuration S A clutch mechanism S An output drive shaft These components are contained in a case which includes a containment ring, an air intake and an exhaust. The containment ring is made to contain a failure of the turbine rotor. The air starter motor also has an oil filler plug, an oil level sight glass and drain plug. The drain plug has a magnetic chip detector (M.C.D.). Air entering the air inlet will turn the turbine rotor at high speed with low torque, the reduction gears reduce the speed to the output drive shaft. This reduction in speed produces high torque to the output drive shaft. The torque is transmitted through the ratchet gear on the ring gear carrier to the clutch mechanism. The clutch mechanism has pawls which engage with the ratchet gear to turn the output shaft. When the engine has reached self sustaining R.P.M. the output drive shaft is turning faster than the ring gear carrier and centrifugal force disengages the clutch pawls from the ratchet gear.

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Figure 185 FRA US-T TH NOV 99

Air Starter Motor Page: 369

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80-00 START CONTROL VALVE The start control valve controls the flow of air to the air starter motor. The solenoid contains a double coil assembly which is controlled by electrical signals from the E.E.C. One coil is connected to lane ’A’ of the E.E.C., the other to lane ’B’. The valve also contains a butterfly valve operated through linkage by two air operated pistons. An extension of the butterfly shaft has a visual control valve position indicator. The control valve position indicator operates two micro switches one is connected to lane ’A’ the other to lane ’B’ to give indication to the E.E.C. of the valve position. Also the extension is a square socket to permit manual operation of the butterfly valve.

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Figure 186 FRA US-T TH NOV 99

Air Start Control Valve Page: 371

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80-00 START CONTROL VALVE CLOSED The valve will remain closed when electrical power is removed from the solenoid, the ball valve is closed by a spring and the air pressure on piston ’B’ is released through vent ’C’. The air pressure acting on piston ’A’ and actuator spring force acting on piston ’B’ keeps the butterfly valve closed. The actuating spring will also close the butterfly valve if there is a decrease in air pressure upstream of the butterfly valve during starting. Any loss of the electrical supply to the solenoid, the butterfly valve will remain closed or go to closed if it was in the open position.

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Figure 187 FRA US-T TH NOV 99

Air Start Valve: Closed Position Page: 373

A330 RR Trent 700

80-00 START CONTROL VALVE OPEN When the valve is commanded by the E.E.C. to open the solenoid is energised, this moves the ball valve to close off vent ’C’ and also allow upstream air pressure to get to piston ’B’. The force acting on piston ’B’ is greater than the combined force of the spring and piston ’A’ therefore, piston ’B’ will move.This movement is transmitted through linkage to open the butterfly valve.

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Figure 188 FRA US-T TH NOV 99

Air Start Valve: Open Position Page: 375

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80-00 ENGINE START CONTROL AND INDICATION FLIGHT DECK CONTROLS The start controls consist of the following: – S Two master switches (levers) S A rotary selector S Manual start push buttons The ’MAN START’ button incorporates a blue ’ON’ legend and is normally in the released position with ’ON’ legend off. Pressing the switch will open the start control valve and illuminate the ’ON’ legend.The amber’FAULT’ warning light will illuminate when a disagreement occurs between the start control valve position and that commanded by the E.E.C. in the ’AUTO’ mode. MASTER SWITCH This switch is for the pilot to use as a master ’ON/OFF’ switch for the engine. S ·’OFF’ POSITION – P.R.S.O.V. commanded closed. F.A.D.E.C. system will behave as it would in the de–powered state. S ·’ON’ POSITION – The E.E.C. will be able to control the P.R.S.O.V. position, and the F.A.D.E.C. system will perform engine control functions.

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ROTARY SELECTOR/MODE SELECTOR The rotary selector valve can be used in conjunction with the master switch and manual start push button to perform: S engine dry cranking S wet cranking S Pilot control starting sequence S automatic starting and continuous ignition

FLIGHT DECK INDICATION During the starting sequence, the ignition and starting parameters are displayed. During the start sequence the nacelle temperature indications are replaced by the following: S Ignition (A, B or AB). S Start control valve position. S Air pressure to the starter. ENGINE INTERFACE VIBRATION MONITORING UNIT (E.I.V.M.U.) This unit is located in the aircraft avionics bay. It receives discrete electrical signals from the flight deck.These signals are digitised and transmitted to the E.E.C. The unit also sends discrete signals to close the air conditioning pack flow valves and to accelerate the auxiliary power unit (A.P.U.) if required. ENGINE ELECTRONIC CONTROLLER (E.E.C.) Generates starter control valve opening and closing signals from information received from the rotary selector, master switch, manual start push button and N3 signal. The E.E.C. also generates warning and caution messages for display in the flight deck through the Electronic Centralised Aircraft Monitoring System (E.C.A.M.)

MANUAL START PUSH BUTTON Selection of the manual start push button enables the pilot to perform alternative engine starting i.e. manual start.

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Figure 189 FRA US-T TH NOV 99

Engine Start Control and Indications Page: 377

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A330 RR Trent 700

80-00 START PROCEDURES PRE–START S Thrust Lever – Idle S Master Switch – Off S Rotary Selector – Norm S Manual Start Push Button – Off S Aircraft Booster Pumps – On AUTO START S Rotary Switch – Ign Start S Master Switch – On S After Successful Start S Rotary Switch – Norm

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DRY CRANKING (ROTATION) S Rotary Switch – Crank S Manual Start Button – On S Engine Accelerates to Maximum Motoring Speed During dry crank the starter motor operates but the P.R.S.O.V. and both ignition systems remain inoperative.

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N3

25%

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N3 IS AT 50%

Figure 190 FRA US-T TH NOV 99

Auto Start Procedure Page: 379

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ENGINE STARTING SYSTEM

A330 RR Trent 700

80-00 MANUAL START S Rotary Switch – Ign Start S Manual Start Push Button – On S When N3 reaches a pre–determined speed pilot put master switch to On S After successful start S Rotary Switch – Norm S Manual Start Push Button – Off Note: During auto start the E.E.C. monitors engine speed and E.G.T. if hung or hot starts are detected, the P.R.S.O.V., start control valve, and ignition are automatically shut off. This control is not available during manual start and conventional pilot monitoring is required. Start abort in both modes can be made by placing master switch to the ’OFF’ position.

AUTOMATIC RE–LIGHT The F.A.D.E.C. system detects ’flame out’ conditions by low combustion chamber pressure and a change in H.P. shaft speed. The E.E.C. will select dual ignition while conditions last and for lO seconds afterwards. Automatic start following ’flame out’ will never be automatically initiated by the E.E.C., but will be commanded by the pilot. The pilot has to: S switch OFF then ON the master switch S for 10 sec are A+B ignition available when N3 > 50% or S master switch OFF / ON until 30 sec S N3 < 10% S P.R.S.O.V (fuel) ON and start relight is initiated with both ignitors ON

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WET CRANKING (MOTORING) S Rotary Selector – Crank S Manual Start Switch – On S Master Switch – On During wet crank the selection of the master switch opens the P.R.S.O.V. but because crank is selected ignition is inoperative. IN–FLIGHT RE–LIGHTING Both automatic and alternative starting is available for in–flight re–lighting. The selection of switches is the same as on the ground. When alternative start is selected the E.E.C. will always command starter assistance. When automatic start is selected the E.E.C. will determine, based on flight envelope and engine parameters, whether a starter assisted start or a windmilling start is required. The E.E.C. will receive a signal from the E.I.V.M.U. as to flight/ground status. The E.I.V.M.U. receives its signal from the Landing Gear Control Interface Unit (L.G.C. I .U.). CONTINUOUS IGNITION Move rotary switch ’NORM’ to ’IGN START’.

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N3

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N3 AT ABOUT 50%

N3

Figure 191 FRA US-T TH NOV 99

Manual Start Procedure Page: 381