270604153-B737-3-ATA-27-L3-e

Training Manual B 737-300/400/500 ATA 27 Flight Controls ATA 104 Level 3

Book Book No: No:

B737 B737-3 -3 27 L3 E

Lufthansa Technical Training GmbH Lufthansa Base

For Training Purposes Only Lufthansa

!

For training purpose and internal use only. Copyright by Lufthansa Technical Training GmbH. All rights reserved. No parts of this training manual may be sold or reproduced in any any form without permission of:

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

B737-3 27 L3 E

TABLE OF CONTENTS

ATA 27 FL F LIGHT CONTROL . . . . . . . . . . . . . . . . . . .

1

27-00 GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLIGHT CONTROL SYSTEM INTRODUCTION . . . . . . . . . . . . . . . . . . FLIGHT CONTROL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . PANEL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PANEL DESCRIPTION (CONT.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PANEL DESCRIPTION (CONT.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PANEL DESCRIPTION (CONT.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PANEL DESCRIPTION (CONT.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLIGHT CONTROL HYDRAULIC MODULE . . . . . . . . . . . . . . . . . . . . . . FLIGHT CONTROLS HYDRAULIC SHUTOFF VALVES . . . . . . . . . . . FLIGHT CONTROLS LOW PRESSURE INDICATION . . . . . . . . . . . . .

2 2 4 6 8 10 12 14 16 18 20

27-10 AILERON AND TAB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON CONTROL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LATERAL CONTROL SYSTEM OPERATION . . . . . . . . . . . . . . . . . . . . LEFT CONTROL COLUMN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RIGHT CONTROL COLUMN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON CONTROL WHEEL DRUM ASSEMBLY . . . . . . . . . . . . . . . . AILERON TRANSFER MECHANISM . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON POWER CONTROL UNIT . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON BUS DRUM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON ASSEMBLY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BALANCE PANELS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON CENTERING AND TRIM MECHANISM . . . . . . . . . . . . . . . . AILERON TRIM ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON AUTOPILOT ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . AILERON AUTOPILOT ACTUATOR (CONT.) . . . . . . . . . . . . . . . . . . . . 27-60 SPOILER AND SPEED BRAKE SYSTEM . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPOILER AND SPEED BRAKE SYSTEM OPERATION . . . . . . . . . . . SPEED BRAKE CONTROL LEVER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPEED BRAKE LEVER NO-BACK BRAKE . . . . . . . . . . . . . . . . . . . . . .

22 22 24 26 28 28 30 32 34 36 38 40 42 44 46 48

SPOILER CONTROL QUADRANT ASSEMBLY . . . . . . . . . . . . . . . . . . SPOILER MIXER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPOILER RATIO CHANGER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLIGHT SPOILER CONTROL QUADRANTS . . . . . . . . . . . . . . . . . . . . FLIGHT SPOILER ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPOILER SHUTOFF VALVE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GROUND SPOILER CONTROL VALVE . . . . . . . . . . . . . . . . . . . . . . . . . GROUND SPOILER IINTERLOCK VALVE . . . . . . . . . . . . . . . . . . . . . . . OUTBOARD GROUND SPOILER ACTUATOR . . . . . . . . . . . . . . . . . . . INBOARD GROUND SPOILER ACTUATOR . . . . . . . . . . . . . . . . . . . . . SPEED BRAKE LEVER ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . SPEED BRAKE ARMING SWITCH (S 276) . . . . . . . . . . . . . . . . . . . . . . RTO SWITCH (S 650) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOMATIC SPEED BRAKE LEVER ACTUATOR . . . . . . . . . . . . . . . AUTOMATIC SPEED BRAKE INDICATION . . . . . . . . . . . . . . . . . . . . . . AUTOMATIC GROUND SPEED BRAKE OPERATION . . . . . . . . . . . .

58 60 64 66 68 70 72 74 76 78 80 80 82 84 86 88

27-20 RUDDER & TAB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RUDDER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RUDDER SYSTEM PRESSURE REDUCER . . . . . . . . . . . . . . . . . . . . . RUDDER PEDAL ASSEMBLY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AFT RUDDER CONTROL COMPONENTS . . . . . . . . . . . . . . . . . . . . . . MAIN RUDDER POWER CONTROL UNIT . . . . . . . . . . . . . . . . . . . . . . RUDDER FEEL AND CENTERING MECHANISM . . . . . . . . . . . . . . . . RUDDER TRIM CONTROL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . RUDDER TRIM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STANDBY RUDDER ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STANDBY RUDDER SHUTOFF VALVE . . . . . . . . . . . . . . . . . . . . . . . . . RUDDER HYDRAULIC SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

90 90 92 94 98 100 102 104 106 108 110 112 114

50 50 52 54 56

27-30 ELEVATOR & TAB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELEVATOR SYSTEM INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . ELEVATOR CONTROL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELEVATOR FORWARD CONTROL QUADRANTS . . . . . . . . . . . . . . . . INPUT TORQUE TUBE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

116 116 118 120 122

B737-3 27 L3 E

TABLE OF CONTENTS ELEVATOR POWER CONTROL UNITS . . . . . . . . . . . . . . . . . . . . . . . . . OUTPUT TORQUE TUBE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELEVATOR TAB CONTROL MECHANISM . . . . . . . . . . . . . . . . . . . . . . ELEVATOR BALANCE PANELS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELEVATOR FEEL COMPUTER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ELEVATOR FEEL COMPUTER SCHEMATIC . . . . . . . . . . . . . . . . . . . . ELEVATOR FEEL AND CENTERING UNIT . . . . . . . . . . . . . . . . . . . . . . ELEVATOR HYDRAULIC SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HYDRAULIC ISOLATION VALVES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

122 124 126 1 28 130 132 134 136 1 38

27-32 STALL WARNING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . STALL WARNING INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OUTPUT SIGNALS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ADJUSTMENT /TEST . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPERATIONAL TEST (CONFIDENCE TEST) . . . . . . . . . . . . . . . . . . . . BITE TEST OF DSWC (CONFIG 1) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BITE TEST OF SMC (CONFIG 2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

14 0 140 144 146 148 148 150 152

27-40 HORIZONTAL STABILIZER . . . . . . . . . . . . . . . . . . . . . . . . . . . TRIM CONTROL SYSTEM INTRODUCTION . . . . . . . . . . . . . . . . . . . . STABILIZER TRIM CONTROL SCHEMATIC . . . . . . . . . . . . . . . . . . . . . STABILIZER FORWARD CONTROL MECHANISM . . . . . . . . . . . . . . . COLUMN SWITCHING MODULE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STABILIZER JACKSCREW AND GEARBOX ASSEMBLY . . . . . . . . . STABILIZER J JA ACKSCREW AND GEARBOX DIAGRAM . . . . . . . . . . . STABILIZER TRIM LIMIT SWITCHES . . . . . . . . . . . . . . . . . . . . . . . . . . . MAIN ELECTRIC ACTUATOR OPERATION . . . . . . . . . . . . . . . . . . . . .

1 54 154 1 56 158 158 160 164 166 168

27-50 FLAPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRAILING EDGE FLAPS INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . FLAP LEVER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLAP CONTROL UNIT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLAP POWER UNIT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLAP HYDRAULIC DRIVE SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLAP LOAD LIMITER . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLAP LOAD LIMITER (CONFIG. 1) . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 70 170 172 174 176 178 180 182

FLAP LOAD LIMITER (CONFIG 2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FLAP TRANSMISSION INSTALLATION . . . . . . . . . . . . . . . . . . . . . . . . . OUTBOARD MID FLAP DRIVE MECHANISM . . . . . . . . . . . . . . . . . . . . OUTBOARD FORE FLAP DRIVE MECHANISM . . . . . . . . . . . . . . . . . . OUTBOARD FLAP FAIRING DRIVE . . . . . . . . . . . . . . . . . . . . . . . . . . . . OUTBOARD AFT FLAP DRIVE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INBOARD MID FLAP DRIVE MECHANISM . . . . . . . . . . . . . . . . . . . . . . INBOARD AFT FLAP DRIVE MECHANISM . . . . . . . . . . . . . . . . . . . . . . INBOARD AFT FLAP CLUTCH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXHAUST GATES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRAILING EDGE BYPASS VALVE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRAILING EDGE BYPASS VALVE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ALTERNATE FLAP DRIVE UNIT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TRAILING EDGE FLAP POSITION INDICATING . . . . . . . . . . . . . . . . . FLAP ASYMMETRY CONTROL CIRCUIT . . . . . . . . . . . . . . . . . . . . . . . T. E. FLAP SYSTEM HYDRAULIC OPERATION . . . . . . . . . . . . . . . . . FLAP ALTERNATE DRIVE SYSTEM CIRCUIT . . . . . . . . . . . . . . . . . . .

184 186 190 192 1 94 196 198 200 204 206 208 210 212 214 216 218 220

27-80 LIFT AUGMENTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE FLAP AND SLATS INTRODUCTION . . . . . . . . . . . . LEADING EDGE FLAP SLAT SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE OPERATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE FLAP AND SLAT CONTROL VALVE . . . . . . . . . . . . LEADING EDGE STANDBY DRIVE SHUTOFF VALVE . . . . . . . . . . . . LEADING EDGE FLAP ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE ACTUATOR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE FLAP MECHANISM . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE SLAT MECHANISM . . . . . . . . . . . . . . . . . . . . . . . . . . . SLAT AUXILIARY TRACK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SLAT MAIN TRACK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE FLAPS AND SLAT OPERATION . . . . . . . . . . . . . . . . AUTOSLAT COMPUTER (CONFIG 1) . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOSLAT CONTROL VALVE (CONFIG 2) . . . . . . . . . . . . . . . . . . . . . AUTOSLAT COMPUTER (CONFIG 1) . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOSLAT COMPUTER SYSTEM (CONFIG 1) . . . . . . . . . . . . . . . . . . AUTOSLAT COMPUTER FRONT PANEL (CONFIG 1) . . . . . . . . . . . .

2 22 222 224 226 228 228 230 2 32 234 236 238 240 242 244 246 248 250 252

B737-3 27 L3 E

TABLE OF CONTENTS AUTOSLAT FAILURE WARNING (CONFIG 1) . . . . . . . . . . . . . . . . . . . AUTOSLAT SYSTEM (CONFIG. 2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOSLAT CONTROL VALVE (CONFIG 2) . . . . . . . . . . . . . . . . . . . . . STALL MANAGEMENT COMPUTER (SMC) (CONFIG 2) . . . . . . . . . . AUTOSLAT CHANNEL OF SMC (CONFIG 2) . . . . . . . . . . . . . . . . . . . . AUTOSLAT FAILURE WARNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE DEVICE INDICATION . . . . . . . . . . . . . . . . . . . . . . . . . LEADING EDGE FLAP POSITION SENSORS . . . . . . . . . . . . . . . . . . . SLAT POSITION SENSORS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THE STALL WARNING SWITCH (S 856) . . . . . . . . . . . . . . . . . . . . . . . . THE FLAP 10_ SWITCH (S 584) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LE FLAP / SLAT POS. INDIC. MODULE (CONFIG 1) . . . . . . . . . . . . . LE FLAP / SLAT POS. INDIC. MODULE (CONFIG 2) . . . . . . . . . . . . .

254 256 258 260 2 62 264 266 268 270 272 2 72 274 276

31-20 INDEPENDENT INSTRUMENTS . . . . . . . . . . . . . . . . . . . . . . TAKEOFF WARNING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TAKEOFF WARNING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TAKEOFF WARNING SCHEMATIC . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2 78 278 280 282

B737-3 27 L3 E

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

Flight Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Control General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Center Controlstand . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Center Control Stand . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Center Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Overhead Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Control Hydraulic Modules . . . . . . . . . . . . . . . . . . . . Flight Controls Hydraulic Shutoff Valves . . . . . . . . . . . . . . Flight Contros Low Pressure Indication . . . . . . . . . . . . . . Aileron Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . Roll Control System Schematic . . . . . . . . . . . . . . . . . . . . Lateral Control System Operation . . . . . . . . . . . . . . . . . . Right and Left Control Column . . . . . . . . . . . . . . . . . . . . . Aileron Control Whell Drum Assembly . . . . . . . . . . . . . . . Transfer Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Power Control Unit . . . . . . . . . . . . . . . . . . . . . . . . Aileron Power Control Units (System A and B) . . . . . . . Aileron Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balance Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Centering and Trim Mechanism . . . . . . . . . . . . . Aileron Trim Control System . . . . . . . . . . . . . . . . . . . . . . . Auto Pilot Actuator Location . . . . . . . . . . . . . . . . . . . . . . . Autopilot Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spoiler Panel Identification . . . . . . . . . . . . . . . . . . . . . . . . Spoiler & Speed Brake System . . . . . . . . . . . . . . . . . . . . Speed Brake Control Lever . . . . . . . . . . . . . . . . . . . . . . . . Speed Brake Lever Brake . . . . . . . . . . . . . . . . . . . . . . . . . Spoiler Control Quadrant and Spoiler Mixer . . . . . . . . . . Spoiler Mixer Control Schematic . . . . . . . . . . . . . . . . . . . Spoiler Deflection Schematic . . . . . . . . . . . . . . . . . . . . . . Spoiler Mixer Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Spoiler Control Quadrants . . . . . . . . . . . . . . . . . . . Flight Spoiler Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Spoiler Hydraulic System . . . . . . . . . . . . . . . . . . . .

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 52 Figure 53 Figure 54 54 Figure 55 Figure 56 Figure 57 Figure 58 58 Figure 59 Figure 60 60 Figure 61 Figure 62 Figure 63 Figure 64 Figure 65 Figure 66 66 Figure 67 Figure 68 Figure 69 Figure 70

Ground Spoiler Control Valve . . . . . . . . . . . . . . . . . . . . . . Ground Spoiler Interlock Valve . . . . . . . . . . . . . . . . . . . . . Outboard Ground Spoiler Actuator . . . . . . . . . . . . . . . . . . Inboard Spoiler Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . Speed Brake Lever Actuator . . . . . . . . . . . . . . . . . . . . . . . RTO Switch Location . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Automatic Speed Brake System . . . . . . . . . . . . . . . . . . . . Automatic Speed Brake Indication . . . . . . . . . . . . . . . . . . Automatic Speed Brake Circuit . . . . . . . . . . . . . . . . . . . . . Rudder & Tab Introduction . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Pressure Reducer . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Pressure Reducer Circuit . . . . . . . . . . . . . . . . . . . Rudder Pedal Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Rudder Control Components . . . . . . . . . . . . . . . . . . . Main Rudder Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Feel and Centering Mechanism . . . . . . . . . . . . . Rudder Trim System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Electrical Schematic . . . . . . . . . . . . . . . . . . Standby Rudder Actuator . . . . . . . . . . . . . . . . . . . . . . . . . Standby Rudder Shutoff Valve . . . . . . . . . . . . . . . . . . . . . Rudder Hydraulic Schematic . . . . . . . . . . . . . . . . . . . . . . . Elevator and Tab System Introduction . . . . . . . . . . . . . . . Elevator Control System . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Forward Control Quadrants . . . . . . . . . . . . . . . . Elevator In Input Torque Tu Tube an and Po Power C Co ontrol Un Units . . Elevator Output Torque Tube . . . . . . . . . . . . . . . . . . . . . . Elevator Tab Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Balance Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Feel Computer Location . . . . . . . . . . . . . . . . . . . Elevator Feel Computer Schematic . . . . . . . . . . . . . . . . . Elevator Feel and Centering Unit . . . . . . . . . . . . . . . . . . . Elevator Hydraulic Schematic . . . . . . . . . . . . . . . . . . . . . . Hydraulic Isolation Valves . . . . . . . . . . . . . . . . . . . . . . . . . Stall Warning Introduction . . . . . . . . . . . . . . . . . . . . . . . . .

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

B737-3 27 L3 E

TABLE OF FIGURES Figure 71 71 Figure 72 72 Figure 73 Figure 74 Figure 75 Figure 76 Figure 77 77 Figure 78 Figure 79 79 Figu Figure re 80 Figu Figure re 81 Figure 82 82 Figure 83 83 Figure 84 Figure 85 Figure 86 Figure 87 Figure 88 Figure 89 Figure 90 90 Figure 91 91 Figure 92 92 Figure 93 Figure 94 Figure 95 95 Figure 96 96 Figure 97 Figure 98 Figure 99 99 Figure 10 100 Figure 10 101 Figure 102 Figure 103 Figure 10 104 Figure 10 105

Stall Warning Component Location . . . . . . . . . . . . . . . . . Stall Warning System Block Diagram . . . . . . . . . . . . . . . Output Signals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BITE Test (config. 1) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BITE Test (config. 2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Stabilizer Trim Control System Location . . . . . . . . . . . . . Stabilizer System Schematic . . . . . . . . . . . . . . . . . . . . . . . Stabilizer Forward Control Mechanism . . . . . . . . . . . . . . Stab Stabil iliz izer er Jack Jacksc scre rew w and and Gear Gearbo box x Asse Assemb mbly ly (con (confi fig. g. 1) Stab Stabil iliz izer er Jack Jacksc scre rew w and and Gear Gearbo box x Asse Assemb mbly ly (con (confi fig. g. 2) Stabilizer Jackscrew and Gearbox . . . . . . . . . . . . . . . . . . Stabilizer Trim Switches Location . . . . . . . . . . . . . . . . . . . Stabilizer Trim Control Circuit . . . . . . . . . . . . . . . . . . . . . . Trailing Edge Flap System . . . . . . . . . . . . . . . . . . . . . . . . . Flap Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Hydraulic Drive Schematic . . . . . . . . . . . . . . . . . . . . Flap Load Limiter Component Location . . . . . . . . . . . . . . Flap Load Limiter System Circuit (config. 1) . . . . . . . . . . Flap Load Limiter System Circuit (config. 2) . . . . . . . . . . Flap Transmission Schematic . . . . . . . . . . . . . . . . . . . . . . Flap Transmission Assembly . . . . . . . . . . . . . . . . . . . . . . Outboard Mid Flap Drive Mechanism . . . . . . . . . . . . . . . Outboard Fore Flap Drive Mechanism . . . . . . . . . . . . . . Outboard Flap Fairing Drive . . . . . . . . . . . . . . . . . . . . . . . Outboard Aft Flap Drive . . . . . . . . . . . . . . . . . . . . . . . . . . . Inboard Mid Flap Drive Mechanism . . . . . . . . . . . . . . . . . Inboard Aft Flap Drive Mechanism . . . . . . . . . . . . . . . . . Inboard Aft Flap Drive Mechanism . . . . . . . . . . . . . . . . . Inboard Aft Flap Clutch . . . . . . . . . . . . . . . . . . . . . . . . . . Exhaust Gates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Trailing Edge Bypass Valve Location . . . . . . . . . . . . . . . Trailing Edge Bypass Valve Circuit . . . . . . . . . . . . . . . .

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

Figure 106 Figu Figure re 107 107 Figure 10 1 08 Figure 109 Figure 11 110 Figure 11 111 Figure 11 112 Figure 113 Figure 114 Figure 115 Figure 116 Figure 117 Figure 118 Figure 11 119 Figure 120 Figure 12 121 Figure 12 1 22 Figure 12 123 Figure 124 Figure 125 Figure 12 1 26 Figure 12 127 Figure 12 1 28 Figure 12 129 Figure 130 Figure 13 131 Figure 132 Figure 133 Figure 134 Figure 135 Figure 13 1 36 Figure 137 Figure 138 Figure 139 Figure 14 140

Alternate Flap Drive Unit . . . . . . . . . . . . . . . . . . . . . . . . . Trail railin ing g Edge Edge Posi Positi tion on Indi Indica cati tion on Comp Compon onen entt Loca Locati tion on Flap Pos. Indicat. and Asym. Control Circuit . . . . . . . . Trailing Edge Flap System Hydraulic Operation . . . . . Flap Alternate Drive System Circuit . . . . . . . . . . . . . . . . Leading Edge Flaps and Slat Introduction . . . . . . . . . . . Leading Edge Devices Component Location . . . . . . . . Leading Edge De Device Basic System Schematic . . . . . Leading Edge Slat and Falp Component Location . . . Leading Edge Flap Actuator . . . . . . . . . . . . . . . . . . . . . . Leading Edge Slat Actuator . . . . . . . . . . . . . . . . . . . . . . . Leading Edge Flap Mechanism . . . . . . . . . . . . . . . . . . . Leading Edge Slat Mechanism . . . . . . . . . . . . . . . . . . . . Slat Auxiliary Track and Detent Arm . . . . . . . . . . . . . . . Slat Main Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Leading Edge Flap and Slat Operation . . . . . . . . . . . . . Auto Slat System Schematic (config. 1) . . . . . . . . . . . . Auto Slat Control Valve (config. 1) . . . . . . . . . . . . . . . . . Auto Slat Computer (config. 1) . . . . . . . . . . . . . . . . . . . . Auto Slat Schematic (config 1.) . . . . . . . . . . . . . . . . . . . Autoslat Computer Front Panel (config 1.) . . . . . . . . . . Autoslat Failure Warning (config. 1) . . . . . . . . . . . . . . . . Autoslat System Schematic (config. 2) . . . . . . . . . . . . . Auto Slat Control Valve (config. 2) . . . . . . . . . . . . . . . . . STALL MANAGEMENT COMPUTER (CONFIG. 2) . . Autoslat Channel of SMC (config. 2) . . . . . . . . . . . . . . . Autoslat Warning (config. 2) . . . . . . . . . . . . . . . . . . . . . . Leading Edge Device Indicating . . . . . . . . . . . . . . . . . . . Leasing Edge Position Sensors . . . . . . . . . . . . . . . . . . . Slat Position Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . Trailing Edge Control Unit Switch Location . . . . . . . . . . LE Flap/Slat Pos.. Ind. Module M229 (config 1) . . . . . LE Flap / Slat Pos. Ind. Module M229 (config 2) . . . . . Takeoff Warning System . . . . . . . . . . . . . . . . . . . . . . . . . Takeoff Warning Component Location . . . . . . . . . . . . . .

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

B737-3 27 L3 E

TABLE OF FIGURES Figure 141

Takeoff Warning Schematic . . . . . . . . . . . . . . . . . . . . . . .

283

g n i n i a r T l a c i n h c e T a s n a h t f u L

y l n O s e s o p r u P g n i n i a r T r o F

FLIGHT CONTROL

ATA 27

FLIGHT CO CONTROL

B737-300/-400/-500


Flight Controls General

27-00

27-00

GENERAL

FLIGHT CONTROL SYSTEM INTRODUCTION Purpose The flight controls provide maneuvering control about the lateral, longitudinal, and vertical axes. They also provide increased lift for takeoff and landing as well as increased aerodynamic drag both in flight and on the ground. System Description Flight controls are divided into three major groups: Primary Flight Controls Primary flight controls consisting of the ailerons, elevators, and rudder. Secondary Flight Controls Secondary flight controls consisting of the spoilers, trailing edge flaps, leading edge devices, and the stabilizer. Warning Systems Two warning systems are associated with the flight control system:  Stall warning provides a warning to the pilots when the airplane approaches a stall condition.  Takeoff warning provides an aural warning to the pilot when certain flight controls are not in the correct position for takeoff.


B737-300/-400/-500



B737-300/-400/-500 27-00

FLIGHT CONTROLS

PRIMARY SYSTEMS

SECONDARY SYSTEMS

AILERON

WARNING SYSTEMS

STALL WARNING SPOILER/SPEED BRAKE

LEADING EDGE FLAP AND SLAT

ELEVATOR y l n O s e s o p r u P g n i n i a r T r o F

TAKEOFF WARNING TRAILING EDGE FLAP

HORIZONTAL STABILIZER

RUDDER

Figu Figure re 1

Flig Flight ht Cont Contro roll Syst System em



27-00 FLIGHT CONTROL DESCRIPTION General Component Locations The ailerons with attached balance tabs are mounted outboard of the outboard flaps behind the rear spar of each wing. The elevators with attached balance tabs are mounted on the aft section of the horizontal stabilizer. A single conventional rudder without a tab is mounted on the aft side of the vertical stabilizer. Five spoiler panels are located on the upper surface of each wing. They are numbered from left to right, 0 thru 9. The flight spoilers, 2, 3, 6, and 7, are outboard of each engine. Ground spoilers 0, 0, 1, 8, and 9 are outboard of the flight spoilers and ground spoilers 4 and 5 are inboard of each engine. Lift devices consist of two pairs of triple slotted trailing edge flaps, three pairs of leading edge slats and two pairs of leading edge flaps. Trailing edge flaps are mounted on tracks attached to the lower surface of each wing, one set inboard of each engine and the other set outboard. Three leading edge slats are installed outboard of the engine and and two leading edge flaps are installed inboard of the engine on the forward surface of each wing. The adjustable horizontal stabilizer is located at the rear of the fuselage below the vertical stabilizer. each wing, one set inboard of each engine and the other set outboard. General Subsystem Features The flight control surfaces are constructed of advanced composite materials or metal, as required to incorporate the latest advances in technology. technology.


B737-300/-400/-500

Interfaces A stall warning system is provided to alert the pilots of an approaching stall condition. The warning is accomplished by applying vibrations to both pilot’s control columns. The takeoff warning system is installed to provide an aural warning to the pilot when takeoff is attempted with any of the following flight controls not in the proper position for takeoff: Speedbrakes, Horizontal Stabilizer, Trailing Edge Flaps or Leading Edge Flaps.

General Operation The ailerons, elevators and rudder control the airplane around the longitudinal, lateral and vertical axes respectively. Normal operation of these primary flight controls is hydraulic power supplied by Systems A and B. Either hydraulic system operating alone can provide effective control of the primary flight controls. Alternate operation with all hydraulic power lost is by manual control for the ailerons and elevators. Balance panels and balance tabs assist in moving the ailerons and elevators against the aerodynamic loads in flight. Alternate operation of the rudder is by standby hydraulic power to a separate actuator. There are no tabs or balance panels installed on the rudder. The rudder has no manual reversion capability. The spoilers are divided into two groups, flight spoilers and ground spoilers. Ground spoilers function only as ground speedbrakes. The flight spoilers function as speedbrakes, both in flight and on the ground, and operate with the ailerons for roll control at higher roll rates. The outboard flight spoilers, 2 and 7, are hydraulically powered by System B with no back-up. The inboard flight spoilers and all of the ground spoilers are powered by System A with no back-up. Both the inboard and outboard trailing edge flaps are operated by a single torque tube drive system. Normal flap operation is by a hydraulic motor supplied from System B. An electric motor can drive the same torque tube system when hydraulic power is not available. The leading edge flaps and slats are normally operated by hydraulic System B as a programmed function of trailing edge flap position. During normal B system operation the flaps are two-position devices, retract and extend, and the slats are three-position devices, retract, extend and full extend. Alternate operation of the leading edge flaps and slats is by the standby hydraulic system. During standby operation the leading edge flaps move only to the extend position and the slats move only to full extend. The standby system cannot be used to retract the leading edge devices. The moveable horizontal stabilizer is the pitch trim device to control the airplane around the lateral axis. The stabilizer is operated by manual trim wheels, a main electric motor or an autopilot electric motor.



B737-300/-400/-500 27-00

RUDDER

ELEVATOR TAB

AILERON

ELEVATOR AILERON BALANCE TAB

GROUND SPOILERS 9 FLIGHT SPOILERS

8

7

STABILIZER 6

OUTBOARD FLAP GROUND SPOILER

SLATS (SHOWN EXTENDED)

INBOARD FLAP 5

GROUND SPOILER FLIGHT SPOILERS 4 GROUND SPOILER


3

2 1

0

LEADING EDGE FLAPS (SHOWN EXTENDED) FROM LEFT TO RIGHT, 1 TO 4

SLATS from left to right, 1 to 6

Figu Figure re 2

Flig Flight ht Cont Contro roll Gene Genera rall



B737-300/-400/-500 27-00

PANEL DESCRIPTION 1

Flight Control Switch

5

OFF - Corresponding - Corresponding hydraulic system pressure ailerons, elevators and rudder shutoff.

Functions only when alternate flaps master switch is in ARM position. DOWN (Momentary) DOWN (Momentary) - Extends leading edge devices fully using standby hydraulic system. When held in DOWN, electrically extends trailing edge flaps.

STDBY RUD (either RUD (either switch) - Corresponding hydraulic system pressure to ailerons, elevators and rudder is shutoff. Turns on standby pump, opens standby rudder shutoff valve and pressurizes standby rudder power control unit.

UP - Electrically retracts trailing edge flaps. Note: Drive Note: Drive motor will cut off when either limit is reached.

ON (guarded ON (guarded position) - Normal operation.

2

Provided B-hydraulic is available and switch was not moved to DOWN before, TE flaps will retract electrically and LE devices will retract hydraulically upon positioning switch to UP. UP.

Flight Control Low Pressure Lights (amber) ON - Indicates low pressure of corresponding hydraulic system to ailerons, elevator and rudder. MASTER CAUTION light and FLT CONT annunciator illuminate. Deactivated when corresponding flight control switch is positioned to STDBY RUD and the standby rudder shutoff valve is open.

6

3 y l n O s e s o p r u P g n i n i a r T r o F

Flight Spoiler Switch (For maintenance purpose only) A - Controls inboard flight spoilers shutoff valve. B - Controls outboard flight spoilers shutoff valve.

4

Alternate Flap Master Switch ARM - Closes - Closes trailing edge flap bypass valve, turns on standby pump, arms alternate flaps position switch and arms standby hydraulic LOW PRESSURE light. OFF - (guarded position) - Normal operation operation

Elevator Feel Differential Pressure Light

(amber)

Armed only when the trailing edge flaps are up. ON - Indicates excessive differential pressure pressure in the elevator feel computer.

On airplanes with the rudder pressure reducer installed, the A system indicates a failure of the pressure reducer to switch back to full system pressure when commanded. NOTE: THE A SYSTEM LIGHT WILL REMAIN ILLUMINATED FOR APPROX. 5 SEC AFTER HYDRAULIC SYSTEM IS ACTIVATED. ACTIVATED.

Alternate Flap Position Switch

7

Auto Slat Fail Light (amber) ON - Indicates failure of both autosalat computers



B737-300/-400/-500 27-00

FLT CONTROL A

STANDBY HYD

B

A ON

1

1

OFF STDBY RUD

B ON

LOW QUANTITY

OFF

LOW PRESSURE

STDBY RUD

HYDRAULIC PUMP SWITCHES

FLIGHT CONTROL PANEL

ALTERNATE FLAPS ARM LOW LOW PRESSURE PRESSURE

UP

4 OFF OFF

SPOILER A

2

5 ON

3

3

OVERHEAD PANEL

ON

OFF

OFF


DOWN

B

FEEL DIFF PRESS

6

SPEED TRIM FAIL

YAW DAMPER

MACH TRIM FAIL

YAW DAMPER

AUTO SLAT FAIL

BRIGHT

7

CLOCK FIRE WARN

ON

MASTER CAUTION MAP

BELL CUTOUT PUSH TO RESET FLT CONT

OFF

IRS FUEL

FLIGHT CONTROL PANEL P5-8

ELEC APU OVHT/DET

LEFT LIGHT SHIELD FLT CONT FAULT LIGHT

Figu Figure re 3

Flig Flight ht Con Contr trol ol Panel anel



B737-300/-400/-500 27-00

PANEL DESCRIPTION (CONT (CONT.) .) 8

Speed Brake Lever

13

DOWN (detent) DOWN (detent) - All flight and ground spoiler panels are in faired position. ARMED - Automatic - Automatic speed brake system armed. All flight and ground spoiler panels extend upon touchdown ( speed brake lever moves to UP position). FLIGHT DETENT - All - All flight spoilers extended to their maximum position for flight use.

Cutout - Removes - Removes autopilot servo power to stabilizer drive.

14

9

All spoiler panels retract on the ground if either throttle is advanced for takeoff (speed brake lever moves to DOWN position). All spoiler panels extend if takeoff is rejected and the reverse thrust levers are positioned for reverse thrust (speed brake lever moves to UP position).

Stabilizer Trnim Handle Provided for manual operation of the stabilizer. Overrides any other stabilizer trim inputs. Handle should be folded inside stab trim wheel for normal operation. Rotates when stabilizer is in motion.


10

Stabilizer Trim Wheel

11

Stabilizer Trim Indicator Indicates units of airplane trim on the adjacent scale.

12

Stabilizer green Band Range Corresponds to allowable range of trim settings for takeoff. NOTE:

SPECIMEN ONLY. ACTUAL GREEN BAND RANGE MAY DIFFER FROM DRAWING.

Stabilizer Trim Main Electric Cutout Switch Cutout - Removes - Removes power from stabilizer main electric trim motor.

15

UP - All flight and ground spoilers are extended to their maximum position for ground use. NOTE:

Stabilizer Trim Autopilot Autopilot Cutout Switch

Flap Lever Selects position of flap control valve directing hydraulic pressure for flap drive unit. Position of leading edge devices is determined by selected trailing edge flap position. At flap lever position position 40, the flap load relief system is armed. This causes automatic flap redaction to flap position 30 or prevents flap extension to flap position 40 in the event of excessive airspeed. The flap lever remains in position 40.

16

Flap Gates Prevents inadvertent flap lever movement beyond:  Position 1 - to check flap position for one engine inoperative goaround.  Position 15 - to check flap position for normal go-around.



B737-300/-400/-500 27-00

8

9

DOWN

10

I N C R E A S E

T S U R H T

T H R U S T

E S A E R C N I

ARMED

16 FLAP UP 0 1

11

12

APL NOSE DOWN C T 0 G A 3 - K 0 % E 2 M O 5 0 1 0 A F C F

FLIGHT DETENT

1

2

2 5

10 UP

10

15

15

25

S T A N D

FLAP

15

0

C 0 F A 3 F M 0 O - 25 E % 0 K 1 G A T C

10

15 APL NOSE UP

30


STAB TRIM 40 PARKING BRAKE PULL

FLAP DOWN

MAIN AUTOELEC PILOT NORMAL

CUTOUT 1 A/T DISENGAGE

2 A/T DISENGAGE

Figu Figure re 4

Cent Center er Con Contr trol ols stand tand

14

13



27-00 PANEL DESCRIPTION (CONT (CONT.) .) 17

Rudder Trim Off Flag (amber) indicates loss of electrical power for rudder trim indicator.

18

Rudder Trim Indicator Indicates units of rudder trim

19

Rudder Trim Control Electrically trims the rudder in the desired direction. (Springloaded to neutral position)

20

Aileron Trim Switches Movement of both switches repositions the aileron neutral control position. (Springloaded to neutral position)

21

Stabilizer Trim Cutout Override Switch Override - Bypasses - Bypasses the control column actuated stabilizer trim cutout switches to restore power to the electric trim.


B737-300/-400/-500



B737-300/-400/-500 27-00

18 17


20

19

21

Figu Figure re 5

Cent Center er Cont Contro roll Sta Stand nd




Flap Position Indicator Indicates position of left and right outboard trailing edge flaps and provides trailing edge flaps asymmetry protection circuit.

23

Flap Load Relief Light (amber) Indicates activation of flap load relief system.

24

LE Flap Extended Light (green) ON - All leading edge flaps extended and all leading edge slats in intermediate position (Flap positions1, 2, and 5) or, all leading edge devices fully extended (Flap positions 10 through 40).

25

LE Flaps Transit Light (amber) ON - Any leading edge device in transit, or not in programmed position with respect to trailing edge flaps. Note: Light Note: Light is inhibited during autoslat operation in flight.

26

Speed Brake Armed Light (green) Light deactivated with speed brake lever in DOWN position. ON - Indicates valid speed brake system inputs.


B737-300/-400/-500

27

Speed Brake Do Not Arm Light (amber) Light deactivated with speed brake lever in DOWN position. On - Indicates invalid signals or test inputs to automatic speed brake system.

28

Speed Brake Test Switches

NOTE: IF INSTALLED MAINTENANCE TEST ONLY



B737-300/-400/-500 27-00

28

23

22

25 24

26 y l n O s e s o p r u P g n i n i a r T r o F

27

Figure 6

Center ter Panel nel




Leading Edge Light Devices Annunciator Panel

Indicates position of individual leading edge flaps and slats. Lights out - Corresponding - Corresponding leading edge device retracted.

30

Leading Edge Devices Transit Light (amber)

On - Corresponding leading edge device in transit.

31

Leading Edge Devices Extended Light (green)

On - Corresponding leading edge slat in intermediate position.

32

Leading Edge Devices Full Extended Light (green)

On - Corresponding leading edge device fully extended.

33

Leading Edge Light Devices Annunciator Panel Test Switch

Press - Tests all annunciator panel lights. NOTE: LIGHT IS INHIBITED DURING AUTO SLAT OPERATION OPERATION IN FLIGHT.


B737-300/-400/-500



B737-300/-400/-500 27-00

30

31 32 y l n O s e s o p r u P g n i n i a r T r o F

29

33

Figu Figure re 7

Aft Aft Ove Overh rhea ead d Pane Panell


FLIGHT CONTROL GENERAL

27-00 FLIGHT CONTROL HYDRAULIC MODULE Purpose Two flight controls hydraulic modules are provided to control hydraulic power for the aileron, rudder, elevator and flight spoiler control systems. Location The System A flight controls module is located in the lower left outboard corner of the main wheel well forward wall. The System B module is in the same position on the right side. Physical Description/Features Each hydraulic module is a manifold assembly containing a spoiler shutoff valve, flight controls shutoff valve, low pressure warning switch and a compensator cartridge. The compensator cartridge maintains return fluid from the aileron, rudder and elevator power control units after hydraulic system shutdown. This fluid compensates for volume changes in the hydraulic fluid due to temperature change or minor fluid loss. Control The flight controls shutoff valves are 28 volt dc motor operated shutoff valves controlled by the respective (A and B) flight control switches on the forward overhead panel. Spoiler shutoff valves are 28 volt dc motor operated valves controlled by the respective (A and B) spoiler stitches on that panel.


B737-300/-400/-500

Operation All of these valves are normally open to provide pressure to the respective flight controls. Each valve can be actuated electrically or manually to remove system pressure from specific flight controls. Monitor Each valve’s manual override lever also functions as a mechanical position indicator. This is the only position indication for the spoiler shutoff valves. A low pressure switch, downstream of the flight control shutoff valves, is connected to an amber low pressure light beneath the respective control switch. The respective light illuminates for low pressure at the switch. Maintenance Practices The A and B flight controls hydraulic modules are interchangeable.



B737-300/-400/-500 27-00

FLT CONTROL A STDBY RUD OFF

STANDBY HYD

B

1

STDBY RUD OFF

2

LOW QUANTITY LOW PRESSURE

B ON

A ON

ALTERNATE FLAPS OFF LOW LOW PRESSURE PRESSURE

UP OFF

OFF

SPOILER A

B

3

4

ON

ARM

DOWN

OFF ON

YAW DAMPER YAW DAMPER

SPEED TRIM FAIL MACH TRIM FAIL AUTO SLAT FAIL

OFF

P

R

R

AILERON

ON

INBOARD

AILERON

”A” SYSTEM

POWER

POWER

PRESSURE

UNIT ”A”

UNIT ”B”

1

3

TO P5 PILOTS OVHD PANEL

P FEEL DIFF PRESS

”B” SYSTEM PRESSURE

2

4

TO OUTBOARD

SPOILERS

SPOILERS

M


M

PS

PS

M

M

TO RUDDER AND ELEVATOR POWER UNITS

TO ”A” SYSTEM

MODULAR UNIT

RETURN

Figur Figure e8

MODULAR UNIT

TO ”B” SYSTEM RETURN

Fligh Flightt Cont Contro roll Hydra Hydrauli ulic c Modul Modules es



27-00 FLIGHT CONTR OLS HYDRAULIC SHUTOFF VAL VALVES VES Location The flight controls shutoff valves are located on the right side of the respective (A and B) flight controls hydraulic modules. Physical Description/Features The flight controls shutoff valve is a spool valve mounted in a cavity in the hydraulic module and attached by four bolts. An electric motor is attached to the valve by four bolts. It is splined to a cam which converts rotary motor action to linear spool travel. Power Each flight controls shutoff valve is controlled electrically by 28 volts dc thru either flight controls valves circuit breaker. Either dc bus number one or number two can power both valves. Control The valves are electrically controlled independently by either the flight control A or flight control B switch on the forward overhead panel. Each valve can also be manually controlled by the manual override lever on the valve. Operation Both flight controls shutoff valve switches are normally guarded ON. 28 volts dc is applied to the open windings and the valves open allowing hydraulic pressure to the aileron, rudder, and elevator systems. When the switch is moved to OFF, OFF, the valve closes and pressure is removed.


B737-300/-400/-500

Maintenance Practices The flight controls shutoff valves and spoiler shutoff valves are interchangeable. The valve and motor can be replaced as a unit or the motor can be replaced separately.



27-00

SPOILER SHUTOFF VALVE


B737-300/-400/-500

FLIGHT CONTROL SHUTOFF VALVE

Figur Figure e9

Flig Flight ht Contr Control ols s Hydra Hydraul ulic ic Shuto Shutoff ff Val Valve ves s



27-00 FLIGHT CONTROLS LOW PRESSURE INDICATION Purpose The flight controls low pressure indicating system provides an amber light indication of low pressure to the aileron, rudder and elevator systems. Location Low pressure switches are screwed into the respective flight controls hydraulic modules for Systems A and B. The amber indicator lights are located directly beneath the respective (A and B) flight controls switches on the forward overhead panel, P5-3. Physical Description/Features Separate indicating systems for hydraulic systems A and B consist of a pressure switch downstream of the flight controls shutoff valve connected to an amber light on the overhead panel through the ON and OFF positions of the respective flight controls switch. Power Electric power for the indicating systems is dc supplied by Master Dim. Control The low pressure light is connected to the pressure switch only when the flight control switch is in the ON or OFF position. When the switch is in the STBY RUD position, the light is operated by a valve position relay controlled by the position of the standby rudder shutoff valve.


B737-300/-400/-500

Operation When system pressure is available downstream of the flight controls shutoff valve, the pressure switch is held open. When pressure falls below a preset value, the switch closes and provides a ground to the light through the flight controls switch on the overhead panel. Monitor Illumination of the low pressure light is accompanied by Master Caution and the FLT CONT annunciator.




B737-300/-400/-500 27-00

Figur Figure e 10

Flig Flight ht Contro Contros s Low Low Pres Pressu sure re Indi Indica cati tion on


FLIGHT CONTROLS AILERON AND TAB

27-10

27-10

AILERON AND TAB

INTRODUCTION Purpose The aileron and aileron trim control system provides airplane lateral control about the longitudinal axis. System Description The aileron system consists of one aileron with balance tab on each wing. The ailerons are positioned by cables that are driven by two hydraulic power control units located on the forward wall of the main wheel well. Control inputs to these power control units are through a cable system actuated by rotation of either control wheel, an electric aileron trim system or the autopilot.


B737-300/-400/-500


FLIGHT CONTROLS AILERON AND TAB

B737-300/-400/-500 27-10 CAPTAIN’S CONTROL WHEEL FIRST OFFICER’S CONTROL WHEEL

CONTROL WHEEL DRUM

AILERON TRIM CONTROL SWITCHES CABLE AB

CABLE ACBA

AILERON CONTROL QUADRANT

CABLE ACBB

CABLE ABSB CABLE AA CABLE ABSA

AILERON CONTROL LINKAGE

AILERON POWER CONTROL UNIT (MAIN WHEEL WELL)


ARTIFICIAL FEEL, CENTERING AND TRIM MECHANISM

CABLE ABSA

CABLE ABSB

AILERON WING QUADRANT

AILERON TRIM ACTUATOR

AILERON BALANCE TAB AILERON

Figu Figure re 11

Aile Ailero ron n Con Contr trol ol Syst System em


FLIGHT CONTROL AILERON AND TAB

27-10 AILERON CONTROL SYSTEM General Component Locations The two control wheels are interconnected by cables attached to the base of each control column in the lower nose compartment. The left and right aileron cables are driven by control drums at the base of each column. The left and right aileron cables run aft through the outboard section of the floor beam to quadrant assemblies above the main wheel well. Quadrant assembly torque tubes project into the wheel well. The aileron control quadrant assembly, assembly, centering spring and trim mechanism, electric trim actuator, aileron power units and aileron control bus drums are located on the left forward wall of the main wheel well. The spoiler control quadrant assembly, aileron spring cartridge and spoiler mixer are located on the right forward wall of the main wheel well. General Subsystem Features The base of the first officer’s control column is equipped with a transfer mechanism. The transfer mechanism allows normal control wheel motion to be transmitted through the left aileron cables only. If a malfunction occurs which jams the aileron control system, lateral control is accomplished by positioning the flight spoilers through the right aileron cables controlled from the first officer’s control wheel.


B737-300/-400/-500

System Interfaces The flight spoilers assist the aileron system in maintaining lateral control. Normal inputs to the aileron system, above a preset amount of control wheel movement, cause a proportionate movement of the flight spoiler panels. General Operation The ailerons may be actuated either hydraulically or mechanically. mechanically. Normal operation is with both Systems A and B hydraulic pressure. Rotation of either control wheel drives the left aileron cables and rotates the aileron quadrant assembly. assembly. This causes input rods to actuate individual A and B hydraulic power control units. Movement of the power control units operates the cable system that positions the ailerons and the spring cartridge which drives the spoiler control quadrant assembly and inputs to the spoiler mixer. The spoiler mixer linkage moves the cables that operate the flight spoilers.


FLIGHT CONTROL AILERON AND TAB

B737-300/-400/-500 27-10

SPEED BRAKE HANDLE

TRANSFER MECHANISM SPOILER CONTROL DRUM

FLIGHT SPOILER

SPOILER CONTROL QUADRANT

AILERON CONTROL WHEEL DRUM

AILERON BUS SYSTEM CABLES


SPOILER MIXER


AILERON SPRING CARTRIDGE

FLIGHT SPOILER

AILERON

AILERON POWER CONTROL UNITS


ARTIFICIAL FEEL CENTERING AND TRIM MECHANISM

AILERON BALANCE TAB

CABLE SLACK TAKEUP AILERON WING QUADRANT

AILERON

Figur Figure e 12

Roll Roll Contro Controll Syste System m Sche Schema mati tic c

AILERON BALANCE TAB



FLIGHT CONTROL Aileron & Tab

B737-300/-400/-500 27-10

LATERAL CONTROL SYSTEM OPERATION Control Sequence Rotation of either control wheel causes rotation of both aileron control bus drums. The captain’s bus drum drives the aileron control drum to tension the left body cables. The cables actuate the left quadrant which rotates the aileron control quadrant assembly and drives the input rods to both hydraulic power control units. The input rods displace the external cranks on the power control units. These cranks position internal control valves to port hydraulic pressure to the actuators. Rotation of the left quadrant assembly also rotates the cam against the springloaded roller inside the trim and centering mechanism. This provides the artificial feel forces at the control wheel. The power control units respond to hydraulic pressure and stroke on their piston rods. This action rotates the output bus drums and positions the ailerons through cables: up on the wing in the direction of roll and down on the opposite wing. The external crank and control valve are returned to null by actuator response. At the same time the spring rod is driven by a crank and rotates the right quadrant assembly in the direction opposite to that which the left had moved. The right cables are tensioned and a spoiler input rod drives the spoiler mixer/ratio changer linkage. At higher rates of roll, the flight spoilers on the up aileron wing are signalled to move up, and those on the opposite wing are signalled to move down. Rotation of the first officer’s bus drum offsets the crank on the right column within the gap space (12" L and R) of the lost motion device. Rotation of the first officer’s control drum recenters the lost motion device. If no hydraulic pressure is available the power control unit external crank would contact the stops on the housing and allow the pilot to manually drive the unit on the piston. The result would be the same as if hydraulic power was operating the system.

Backup Operation A failure in the aileron control system that jams the left cable system prevents an input to the power control units. In this case the first officer could maintain lateral control by positioning the flight spoilers. He would rotate his control wheel until the crank on his column contacted the lugs on the right control drum. The spring in the transfer mechanism would be overridden by continued input to tension the right body cables. Rotation of he right quadrant assembly would input to the spoiler mixer/ratio changer to move the spoilers. A failure in the spoiler control system that jams the right cables would still allow the captain to input to the power control units by overcoming the spring in the transfer mechanism.




B737-300/-400/-500 27-10

Figur Figure e 13

Late Latera rall Con Contro troll Sys Syste tem m Ope Opera rati tion on



27-10 LEFT CONTROL COLUMN Purpose The left control column is installed to provide the captain a means of controlling airplane roll and pitch. Location The left control column projects thru the floor of the flight compartment. The cable drums, force transducer and force limiter are in the lower nose compartment. Physical Description/Features The captain’s control wheel is mounted at the top of the left control column. It is interchangeable with the first officer’s control wheel as long as the stabilizer trim switches and autopilot release are maintained on the outboard horn. A shaft through the control column is connected to the aileron control wheel drum assembly by a blade type universal joint. Bus cables connect the left and right aileron control bus drums. A CWS roll force transducer provides the connection between the bus drum and aileron control drum. The left body cables run between the aileron control drum and the aileron control quadrant assembly. assembly. The force transducer will be covered in detail in chapter 22. Mounted at the bottom of the drum assembly is an aileron force limiter. The force limiter will be covered in detail in chapter 22.


B737-300/-400/-500

Control Lateral control is accomplished by rotating the control wheel either left or right which rotates a shaft in the column through an angle gear. The shaft rotates the bus drum through a blade type universal joint. This motion is transmitted to the control drum through the force transducer and the cables are actuated to input to the power control units.

RIGHT CONTROL COLUMN Purpose The right control column is installed to provide the first officer a means of controlling airplane roll and pitch. Location The right control column projects through the floor of the flight compartment. A transfer mechanism which includes the cable drums and the lost motion device are in the lower nose compartment. Physical Description/features The first officer’s control wheel is mounted at the top of the right column. It is interchangeable with the captain’s control wheel if the previously discussed changes are made. A shaft through the control column is connected to the aileron transfer mechanism by a blade type universal joint. Bus cables connect the left and right aileron control bus drums. A lost motion device provides the connection between the bus drum and spoiler control drum. Control Lateral control is accomplished by rotating the control wheel either left or right which rotates a shaft in the column through an angle gear. The shaft rotates the bus drum through a blade type universal joint. The lost motion device prevents this motion being transmitted to the spoiler control drum during normal hydraulic or mechanical operation. Input to the aileron system is transmitted through the bus control cables to the left bus drum and left body cable system to the power control units. The lost motion device is returned to neutral by the follow-up action from the movement of the power control units driving the right body cables through the spring cartridge. If a malfunction occurs which jams the aileron control cables the first officer can maintain lateral control with the flight spoilers by engaging the lost motion device after 12 left or right control wheel movement.



B737-300/-400/-500 27-10

FIRST OFFICERS CONTROL COLUMN

AILERON CONTROL BUS DRUM

CAPTAINS AILERON CONTROL WHEEL

WHEEL

CABLE AB

CABLE ACBA

TO FIRST OFFICER’S CONTROL COLUMN

CAPTAINS CONTROL COLUMN

BUS DRUM CABLE SHIELDS CWS ROLL FORCE TRANSDUCER

CABLE AA LOST MOTION DEVICE AILERON CONTROL DRUM

CABLES TO POWER CONTROL UNIT


CABLE ACBB

AILERON FORCE LIMITER

AILERON CONTROL BUS DRUM

2

CABLE AB CONTROL COLUMN - ROLL AILERON CONTROL DRUM

2

AIRPLANES WITH MECHANICAL AILERON FORCE LIMITER

Figu Figure re 14

CABLE AA

Righ Rightt and and Left Left Con Contr trol ol Col Colum umn n

FWD



27-10 AILERON CONTROL WHEEL DRUM ASSEMBLY Purpose The aileron control wheel drum assembly provides the drive connection between both control wheels and the aileron (left) input cables. Location The drum assembly is located at the base of the captain’s column, within the lower nose compartment. Physical Description/Features The drum assembly consists of a bus drum, control drum, crank and force transducer. The bus drum and crank are mounted on the shaft and directly driven by shaft rotation. The control drum is bearing mounted and not directly driven by shaft rotation. The force transducer provides the connection between the drum assembly shaft driven crank and the aileron control drum. Control The CWS force transducer is an electrical device that supplies control wheel steering signals to the autopilot roll control channels which are proportional to pilot control wheel force when the autopilot is engaged. The rod end of the force transducer is attached to the crank and the housing is attached to an arm that is part of the control drum casting. Rotation of either control wheel either compresses or extends the transducer through the crank and generates the electric signal. The CWS force transducer will be covered in detail in Chapter 22, Autoflight.


B737-300/-400/-500

Operation During normal, non-autopilot operation, rotation of either control wheel causes simultaneous rotation of both the bus drum and crank. Motion of the crank is transmitted to the control drum through the force transducer and drives the aileron input cables. Mechanical stops are installed to ensure control input should the force transducer attach points fail.



B737-300/-400/-500 27-10

WHEEL

TO FIRST OFFICER’S CONTROL COLUMN BUS DRUM

CWS ROLL FORCE TRANSDUCER


CABLES TO POWER CONTROL UNIT

AILERON FORCE LIMITER 2 CONTROL COLUMN - ROLL

2

AIRPLANES WITH MECHANICAL AILERON FORCE LIMITER

Figur Figure e 15

Aile Ailero ron n Cont Control rol Whel Whelll Drum Drum Ass Assem embly bly



27-10 AILERON TRANSFER MECHANISM Purpose The transfer mechanism is installed to provide separation between the captain’s (left) and first officer’s (right) cable systems so that, if either side is jammed, lateral control can be maintained maintained by the operational system. Location The transfer mechanism is located at the base of the first officer’s column. Physical Description/Features The transfer mechanism is composed of a lost motion device, device, consisting of a crank and lugs, and a torsion spring that is preloaded inside a spring container. The aileron control bus drum is attached to the upper half of the spring container and the spoiler control drum is attached to the lower half.


B737-300/-400/-500

Operation In normal operation, motion is transmitted between the first officer’s column shaft and bus drum through the preloaded torsion spring. If a malfunction occurs, jamming either the aileron or spoiler control system, the other system can be operated independently. In this case, the captain or the first officer, depending on the jammed system, will have to overcome the spring preload in the transfer mechanism and operate the other system to maintain lateral control. If the aileron cables are jammed, the first officer will exert a force to overcome the spring preload and operate the spoilers. The first officer must turn his control wheel 12" before the lost motion device begins driving the spoiler control drum. If the spoiler cables are jammed, the captain will exert a force to overcome the spring preload and operate the ailerons through the left cable system.



B737-300/-400/-500 27-10

AILERON CONTROL BUS DRUM TRANSFER MECHANISM

SPOILER CONTROL DRUM

TORSION SPRING SPRING CONTAINERS


LOST MOTION DEVICE PICK-UP

PICK-UP LUG

INBD

AFT

Figu Figure re 16

Trans ransfe ferr Mech Mechan anis ism m



27-10 AILERON POWER CONTROL UNIT Purpose Two identical power control units are installed to provide hydraulic power from System A and System B to operate the ailerons. Location The power control units are mounted on the left forward wall of the main gear wheel well. System A supplies hydraulic pressure to the lower unit, System B to the upper unit. Physical Description/Features The power control unit incorporates a main actuator, a bypass valve, a filter and a main control valve operated by a dual input crank. Control is normally by the primary slide with the secondary slide available in case of failure of the primary. An input rod provides actuation of the control valve from the control system. Control Rotation of either control wheel operates the left cables to rotate the aileron control quadrant assembly torque tube. Two input rods are operated by the torque tube to actuate the input cranks on both power control units.


B737-300/-400/-500

Operation The power control unit piston rod ends are fixed to structure and the housing is connected to a shaft that drives output bus drums and the spring cartridge. when the hydraulic pressure is ported inside, the housing strokes on the piston and positions the ailerons. Movement of the power control unit returns the input crank to neutral which closes the main control valve and stops aileron movement. The bypass valve closes when input pressure drops below 645 75 psi. Input pressure is blocked from reaching the control valve. Both actuator chambers are interconnected to provide a runaround for fluid during aileron movement by external sources. Either power control unit is capable of hydraulically powering both ailerons. Mechanical stops on the housing allow the external cranks to drive the power control unit manually on the piston in response to aileron control system inputs when no hydraulic power is available.

Maintenance Practices The power control units are interchangeable with each other and with the elevator power control units.




B737-300/-400/-500 27-10

Figu Figure re 17

Aile Ailero ron n Powe Powerr Cont Contro roll Unit Unit



27-10 AILERON BUS DRUM Purpose Two aileron bus drums transmit the motion from the hydraulic power control units to the ailerons by means of cables. Location The bus drums are located on the forward wall of the main gear wheel well. Physical Description/Features Each aileron bus drum is mounted on an individual crankshaft. The power control unit housings are connected to the respective output cranks that drive their shafts. Cables from the upper bus drum run out to the right wing aileron and from the lower drum to the left wing aileron. Three shear bolts between each bus drum and its output shaft protect the system against one jammed aileron preventing response of the other one. Control Both bus drums are joined by a fork and lug that permits simultaneous operation of both ailerons from one power control unit. Should either power control unit jam, however, all aileron response will be stopped because of this fork and lug arrangement. Operation Movement of the power control units rotate their respective bus drums through the output cranks. This rotation positions one aileron up and the other down.


B737-300/-400/-500



B737-300/-400/-500 27-10

AILERON BUS DRUMS


Figur Figure e 18

Ailer Aileron on Powe Powerr Contr Control ol Unit Units s (Sys (Syste tem m A and and B)



27-10 AILERON ASSEMBL ASSEMBLY Y Purpose One aileron is installed on each wing to provide airplane lateral control. Physical Description/Features The aileron assembly consists of the aileron, balance tab, and a balance panel. The aileron front spar is connected to wing structure by four hinge fittings. The nose of the aileron is connected to a balance panel in aileron bay number three. The tab is connected to the aileron rear spar by four hinge fittings. Control Cables connect the respective output bus drum in the wheel well with the aileron cable quadrant forward of each aileron. Aileron movement is controlled by a pushrod between the aileron and the cable quadrant. Balance tab movement is controlled by dual tab control rods that pass through the aileron and connect the tab to a support fitting mounted on the wing rear spar. Operation Response to a roll command results in quadrant rotation that moves the aileron up on the wing in the direction of roll. The opposite aileron moves down. Balance tab movement is opposite to the aileron down when the aileron moves up and vice versa.


B737-300/-400/-500

Maintenance Practices The aileron and tab are independently balanced. The balance tab is balanced by weights attached to the lower surface of the balance panel. The correct number of weights is stamped on a data plate attached to each tab. Both the balance panel and the balance tab can be replaced without rebalancing the aileron. Rigging of the ailerons is accomplished by aligning a target on the aileron with another on the adjacent trailing edge rib. Maximum allowable travel is measured in inches between these targets.




B737-300/-400/-500 27-10

Figu Figure re 19

Aile Ailero ron n As Assemb sembly ly



27-10 BALANCE PANELS Purpose An aileron balance panel is installed on each aileron assembly to reduce the force required to position the aileron in flight. Location The balance panel is located in the number three aileron bay. Physical Description/Features The aft end of the balance panel is attached to the aileron nose by a continuous hinge. The forward end is attached to wing structure through an idler hinge to provide articulation of movement. Seals are connected to the hinge assemblies and along both sides of the balance panel creating two separate chambers. The upper chamber is vented to the airstream over the upper wing and the lower chamber to the airstream over the lower wing. Operation When there is no lateral input, pressure forces are developed across the balance panel that maintain the aileron in neutral. When the aileron is deflected, the change in differential pressure drives the balance panel in the opposite direction of control surface movement. Differential pressures are also developed at the tail of the aileron by balance tab movement. Movement of the balance panel and tab provide an assist to the power source deflecting the aileron. These forces are always applied but are particularly useful during manual control.


B737-300/-400/-500



B737-300/-400/-500 27-10

NEUTRAL


DEFLECTED

Figure ure 20

Balance P Pa anel nel


Flight Controls Aileron & Tab

27-10 AILERON CENTERING AND TRIM MECHANISM Purpose The aileron centering spring and trim mechanism provides aileron control system centering, trim and artificial feel. Location This mechanism is located near the base of the aileron quadrant assembly between the two power control unit input rod cranks. The quadrant assembly is attached to the left forward wall of the main gear wheel well. Physical Description/Features The aileron centering spring mechanism consists of a cam, roller arm support and two springs. The centering cam is bolted to the control quadrant shaft. A cam roller is mounted on the roller arm which is mounted on the support. Two springs are connected between the roller arm and the support to provide the force necessary to hold the roller against the cam, thus providing aileron control system centering and artificial feel. An aileron trim electric actuator is connected between a fixed bracket and the roller arm support.


B737-300/-400/-500

Operation Centering springs hold the roller against the center of the cam until a control wheel input is made. The cam rotates with the quadrant shaft when an input is made driving the roller up the inclined cam. This stretches the springs and provides artificial feel at the control wheel. Trim input drives the roller arm up the inclined cam. The springs hold the cam centered, thus backdriving the quadrant assembly. This will cause an input to the power control units and, through the cables, to the control wheel.



B737-300/-400/-500 27-10 PRESSURE DECK


CONTROL QUADRANT SHAFT

AILERON BUS DRUMS


CENTERING SPRING (2 LOCATIONS)

SEE


ACTUATOR ARM ROLLER ARM

INPUT ROD

TRIM MECHANISM


CENTERING SPRING (2 LOCATIONS)

ACTUATOR ATTACH POINT

TRIM ROLLER

A

A

SUPPORT AILERON CONTROL QUADRANT SHAFT

Figur Figure e 21

Ailer Aileron on Cente Centeri ring ng and and Tri Trim m Mech Mechan anis ism m

CAM



27-10 AILERON TRIM ACTUATOR ACT UATOR Purpose The aileron trim actuator electrically operates the ailerons and control wheels to adjust the ailerons to a neutral or trimmed (wings level) condition. Location The aileron trim actuator is located near the left forward wall of the main wheel well, adjacent to the centering spring mechanism. Physical Description/Features The aileron trim actuator is a 115 volt ac single phase, reversible motor. The motor is equipped with limit switches, mechanical stops at the stroke ends, and a brake to limit overcast. The actuator arm is connected to the roller arm support in the centering spring mechanism. It extends or retracts to position the ailerons. Power Electric power for actuator operation is 115 volts ac from Transfer Bus Number One through a circuit breaker on P-6. Control Two aileron trim switches on the aft end of the control stand control actuator operation. The switches must be operated simultaneously. simultaneously.


B737-300/-400/-500

Operation Aileron trim is only available when at least one hydraulic system is operable. Operation of the aileron trim switches drives the actuator arm to alter the position of the centering spring mechanism. This causes an input to the power control units that changes aileron neutral. Ten Ten units either direction of aileron trim is available to move the aileron 15 " up or down. Monitor An aileron trim indicator is located on the top of each control column.



B737-300/-400/-500 27-10

FWD

CAM ROLLER ARM

AILERON PCU INPUT TORQUE TUBE CAM HUB CENTERING CAM ROLLER ARM AND CENTERING SPRING SUPPORT AILERON TRIM ACTUATOR, M1124

L WING DN

L WING DN 115V AC XFR BUS 1


C786 AILERON TRIM CONTROL

P6-2 CIRCUIT BREAKER PANEL

L WING DN

R WING DN

R WING DN M1126 AILERON/ RUDDER TRIM CONTROL (P8-43)

AILERON

LEFT WING DOWN

R WING DN M1124 AILERON TRIM ACTUATOR

RIGHT WING DOWN

Figur Figure e 22

Aile Ailero ron n Trim rim Contr Control ol Syst System em



27-10 AILERON AUTOPILOT AC TUATOR Purpose Two independent hydraulic autopilot actuators provide autopilot input to the roll control system. Location The autopilot actuators are located on a support bracket mounted on the left forward wall of the main gear wheel well. Power The left autopilot actuator is powered by hydraulic System B, the right by hydraulic System A. Either actuator, as selected, may be used for single channel operation or both may be selected for dual channel operation. Control The output cranks of the autopilot actuators are linked together and connected to the upper power control unit crank on the aileron quadrant assembly shaft. Output from either autopilot actuator drives the quadrant assembly resulting in an input to both power control units. Maintenance Practices The autopilot actuators are interchangeable with each other and with the elevator autopilot actuators. The autopilot actuator output crank arm contains shear pins (not shown) that allow a jammed actuator to be overridden by pilot force.


B737-300/-400/-500



B737-300/-400/-500 27-10

TRANSFER VALVE

PRESSURE SWITCH ROD ASSEMBLY

A/P ACTUATOR B

SEE

A/P ACTUATOR A LEVER ASSEMBLY LEVER ASSEMBLY

PRESSURE SWITCH AND ANGLE ADAPTER

ROD ASSEMBLY


ACTUATOR POSITION SENSOR (LVDT) A

SHEAR RIVETS (4 PLACES)

PCU INPUT LEVER

Figu Figure re 23

Auto Auto Pilo Pilott Actu Actuat ator or Loc Locat atio ion n

A



27-10 AILERON AUTOPILOT ACTUATOR (CONT.) (CONT.) Physical Description/Features Each autopilot actuator unit consists of a filter, two solenoid valves, a pressure regulator and relief valve, an electrohydraulic transfer valve, a main actuator piston, two detent pistons and a linear transducer. Power The left autopilot actuator is powered by hydraulic System B, the right by hydraulic System A. Either actuator, as selected, may be used for single channel operation or both may be selected for dual channel operation. Control The output cranks of the autopilot actuators are linked together and connected to the upper power control unit crank on the aileron quadrant assembly shaft. Output from either autopilot actuator drives the quadrant assembly resulting in an input to both power control units.


B737-300/-400/-500

Operation Hydraulic pressure is filtered and delivered to solenoid valve number 1. When the respective roll channel is engaged, solenoid valve 1 opens and delivers pressure to solenoid valve number 2 and the transfer valve. Electric signals from the autopilot operate the transfer valve to control the main piston. After a delay, to allow the main piston to move to the current position of the aileron control system, solenoid number 2 opens and delivers pressure, regulated by the pressure regulator, to the detent piston. These detent pistons clamp the output crank so that movement of the main piston is transmitted to the output crank. The pilot can override this detent pressure and take over manual control. Monitor The pressure switch is part of the auto pilot engage interlocks. The autopilot cannot remain engaged with low detent regulator pressure. The linear transducer sends a follow-up signal to the autopilot.



B737-300/-400/-500 27-10 ARM SOLENOID

DETENT SOLENOID AUTOPILOT TRANSFER FILTER VALVE

FILTER

DETENT CONTROL ENGAGE ORIFICE PRESSURE PORT

A/P JET SIGNAL PIPE INPUT CONTROL SPOOL

AIL HYD PRESSURE SWITCH


RATE LIMITING ORIFICES

RETURN PORT

PRESSURE REGULATOR & RELIEF VALVE

ACTUATOR POSITION SENSOR (LVDT)

DETENT SPRINGS

Figu Figure re 24

EXTERNAL OUTPUT CRANK

DETENT PISTONS INTERNAL OUTPUT CRANK

Auto Autopi pilo lott Actu Actuat ator or


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS

27-60

SPOILER AN AND SPEED BRAKE SYSTEM

INTRODUCTION Purpose Spoiler panels are installed to supplement the ailerons for lateral control and to provide increased drag and reduced lift when used as speedbrakes. System Description Five hydraulically powered spoiler panels are installed on each wing. They are numbered from left to right for identification, - 0 thru thru 4 on the the left left wing wing and and - 5 thru thru 9 on the the right right w wing ing.. The fight spoilers are - 2, 3, 6, and and 7. 7. - The remaini remaining ng six are ground ground spoile spoilers. rs. Hydraulic Source The Ground Spoilers are supplied by the - Hydraulic Hydraulic System System A (can (can be used used on Ground Ground only). Flight Spoiler are supplied by: - Inboard Inboard Flight Flight Spoilers Spoilers (3 and and 6) Hydraulic Hydraulic System System A - Outboard Outboard Flight Flight Spoilers Spoilers (2 (2 and 7) Hydr. Hydr. System System B


B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

GROUND

GROUND

SPOILERS

SPOILERS

0

1

5

4

2

3

6

7

FLIGHT

FLIGHT

SPOILERS

SPOILERS


Figur Figure e 25

Spoi Spoile lerr Pane Panell Iden Identi tific ficati ation on

8

9



27-60

SPOILER AND SPEED BRAK E SYSTEM OPERATION Control Sequence The spoiler panels are divided into two groups: Flight Spoilers - Num Number bers s 2 and 7, 7, out outboa board. rd. - Num Number bers s 3 and and 6, inbo inboard ard.. Ground Spoilers - Numbers 0, 1, 4, 5, 8 and 9. Flight spoilers operate in two modes: - Lateral Lateral Control Control - aileron aileron assist assist or backup backup - Speed Brakes Brakes - both both in flight flight and on on the ground ground Ground spoilers operate only as ground speed brakes Spoiler panels are hydraulically operated: - Outboard Outboard Flight Flight Spoilers Spoilers - System System B - Inboard Inboard Flight Flight Spoilers Spoilers - System System A - Ground Ground S Spoi poiler lers s - Syste System mA Subsystem Sequence Flight spoilers assist the ailerons in maintaining airplane lateral control. An input from either control wheel results in an input to hydraulic power control units via the left cable system. The power control units move one aileron up and the other down. At rates of roll in excess of 10" control wheel the flight spoilers are signal led to move in the same direction as the aileron on that wing.


B737-300/-400/-500

If the left cable system is jammed, the first officer would input through the right cables and maintain lateral control with the flight spoilers. Flight spoilers are used as speed brakes in flight. Control is by manual operation of the speed brake lever through cables to the spoiler mixer and ratio changer. Flight spoilers on both wings rise when operated as speed brakes. Normal Sequence Ground speed brake operation is normally accomplished automatically by an electric actuator driving the speed brake lever. The pilot arms this system in flight and, after landing, all of the spoiler panels rise in response to lever movement.

Backup The speed brake lever is manually operated after landing when speed brakes are required and the automatic system is not operable.



B737-300/-400/-500 27-60


Figu Figure re 26

Spoi Spoile lerr & Spee Speed d Brak Brake e Syst System em


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS SPEED BRAKE CONTROL LEVER Purpose The speed brake lever provides the pilot with control over speedbrake operation. Location The speed brake lever is located on the left side of the control stand in the flight compartment. The forward control assembly is located in the lower nose compartment. Physical Description/Features The speedbake lever is connected to a forward control assembly by a set of control rods. The forward assembly consists of a cable quadrant connected to the lever linkage through a no-back brake. An electric actuator is mounted between structure and the forward control assembly. Control The speedbrakes are controlled manually by rotation of the lever which operates the cables attached to the forward cable quadrant. The speedbrakes are also controlled automatically by the electric actuator driving both the cable quadrant and the lever. A speed brake arming switch is actuated by a cam on the quadrant to arm the electric actuator for automatic operation.


Operation The speedbrake lever is held in a detent in the control stand by a compression spring when positioned to the full down position. The lever must be lifted clear of the detent and rotated to operate the speed brakes. A cam is incorporated in the thrust reverser drum assembly which will lift the speed brake lever out of the down detent whenever thrust reverse is selected. A refused takeoff switch is actuated when this happens to arm the electric actuator. Ground speedbrakes will automatically deploy if the airplane is on the ground with wheelspeed above 60 knots. Monitor speed brake takeoff warning switch inside the control stand sends a signal to the takeoff aural warning system whenever the speedbrake is moved out of the down detent.

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

SPEED BRAKE SEE

A

CONTROL LEVER

SPEED BRAKE LEVER TAKEOFF WARN. SWITCH CONTROL CABIN FLOOR SPEED BRAKE LEVER BRAKE

MOUNTING SCREWS


SPEED BRAKE LEVER ACTUATOR

SPEED BRAKE FORWARD DRUM

A

SPEED BRAKE ARMING SWITCH, S276

Figu Figure re 27

Spee Speed d Brak Brake e Con Contr trol ol Leve Leverr

FW FWD

INBD


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS SPEED BRAKE LEVER NO-BACK BRAKE Purpose The speed brake lever no-back brake prevents the lever from being repositioned by vibration or cable feedback. Location The no-back brake is part of the speed brake forward control assembly, mounted on the ceiling of the lower nose Compartment. Physical Description/Features The no-back brake consists of an input arm, a shoe assembly, a drive arm, a brake quadrant, a shaft and a support bracket. The lower control rod from the speed brake lever is attached to the input lever. The speed brake lever electric actuator is attached to the support bracket which is bearing mounted on the shaft. The outboard end of the shaft is splined for attachment of the speed brake forward drum. Control Four pins in the shoe assembly are spring-loaded against the quadrant. The shoe assembly is mounted on a pivot pin through the drive arm. The drive arm is fixed to the shaft and rotates with the shaft and drum. The brake quadrant aid support bracket are held from rotating by the electric actuator. A rotational force applied to the forward drum is transmitted to the shoe assembly through the pivot pin. This force on the shoe assembly results in a locking action by two diagonally opposed locking pins.


Operation Motion of the input arm caused by rotation of the speed brake lever causes striker bolts to unload the pins in the shoe assembly, unlocking the brake. This allows the drive arm to rotate the shaft. When lever rotation is stopped, the compression spring returns the shoe pins to the locked position. Operation of the electric actuator rotates the support bracket and quadrant assembly. The pins remain locked to the quadrant and the motion is transmitted through the shoe assembly to the drive arm causing the shaft to rotate. The speedbrake lever is also driven through the input arm. A force at the speed brake lever will unlock the shoe assembly and override the electric actuator.

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

CONTROL ROD

DRIVE ARM


INPUT ARM

SPEED BRAKE LEVER BRAKE SCHEMATIC SHOE ASSEMBLY


SHOE PIN (4) SHOE PIVOT PIN

SHAFT ALLOY STEEL FRICTION QUADRANT

SUPPORT BRACKET

SPEED BRAKE LEVER BRAKE

Figu Figure re 28

Spee Speed d Bra Brake ke Leve Leverr Bra Brake ke


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS SPOILER CONTROL QUADRANT ASSEMBLY Purpose The spoiler control quadrant assembly provides the lateral control system input to the flight spoilers. Location The spoiler control quadrant is located in the pressurized area above the right main gear wheel well. The shaft projects into the wheel well, near the forward wall. Physical Description/Features The assembly consists of a spoiler control quadrant attached to the first officer’s (right cables) and a shaft containing two cranks. Control A spring cartridge is connected between the upper aileron bus drum shaft and the input crank at the base of the spoiler control quadrant shaft. A ratio changer input rod is connected between the output crank near the top of the quadrant shaft and the spoiler mixer and ratio changer. The spoiler system is isolated from the aileron system by four shear rivets at the attach point between the spring cartridge and the control quadrant shaft input crank.


Operation When the aileron power control units drive the bus drums to position the ailerons, rotation of the upper bus drum shaft rotates the input crank on the spoiler quadrant shaft through the spring cartridge. This drives the ratio changer input crank to operate the spoiler mixer linkage. The linkage rotates cable quadrants mounted on the ratio changer to signal the flight spoilers to move up on one wing and down on the other.

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60 CABLES TO TRANSFER MECHANISM WHEEL WELL FORWARD BULKHEAD SPOILER CONTROL QUADRANT RATIO CHANGER SPOILER INPUT ROD RATIO CHANGER

FWD

GROUND SPOILER CONTROL VALVE

SPOILER MIXER

SEE A

WHEEL WELL CEILING

SHEAR RIVETS (4 PLACES)


SPOILER OUTPUT QUADRANTS AILERON SPRING CARTRIDGE AILERON BUS DRUMS A

Figur Figure e 29

Spoi Spoile lerr Contr Control ol Quad Quadra rant nt and and Spoil Spoiler er Mixe Mixerr



27-60

SPOILER MIXER Purpose The spoiler mixer combines lateral input from the aileron system with speed brake Iever position to allow the flight spoilers to augment lateral control when simultaneously being used as speedbrakes. Location The spoiler mixer is mounted on the spoiler ratio changer located on the forward bulkhead of the right wheel well. Physical Description/Features The spoiler mixer housing is fastened to the ratio changer by four bolts and four splined shafts which mate with the ratio changer. These shafts are speedbrake input, aileron input, left spoiler output and right spoiler output. An additional shaft on the mixer provides the speed brake signal to the ground spoiler control valve. The housing of the spoiler mixer contains an aileron cam and related levers and links. Input/Output Two separate inputs operate the spoiler mixer linkage. Aileron system input from the right quadrant assembly controls the spoiler mixer cam through the ratio changer. Input from the speedbrake lever is by cables from the forward quadrant to the speed brake quadrant mounted on the ratio changer. A no-back device is added to the speedbrake input quadrant to ensure that spoiler authority for lateral control is maintained if speed brake cable failure occurs.


B737-300/-400/-500

Operation Counterclockwise rotation of the speedbrake input quadrant operates the linkage in the mixer to cause an UP signal at both output quadrants. Clockwise rotation causes a DOWN signal at both quadrants. In either case the linkage also positions the ground spoiler control valve. The no-back device prevents quadrant movement in either direction unless actuated by an input from the speedbrake lever. Inputs from the aileron system through the ratio changer rotate the mixer cam. Clockwise rotation of the cam drives the lower part of the linkage to provide an UP signal to the left output quadrant and a DOWN signal to the right output quadrant. Counterclockwise cam rotation reverses the signals to the output quadrants. Maintenance Practices Sealed bearings are used in all linkages, so the housing contains no oil. The spoiler mixer may be removed from the airplane without disturbing the rigging of the cables or linkage.


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS CPT. CONTROL WHEEL

B737-300/-400/-500 27-60

NORMAL

F/O CONTROL WHEEL

SPEED BRAKE LEVER

ALTERNATE

AILERON

spring cartridge

PCU’S

RATIO CHANGER

MIXER

SPOILER ASYMMETRICAL

>31 SPEED BRAKE LEVER

SYMMETRICAL


OR


AIR/GRD SIGNAL

GROUND SPOILER INTERLOCK VALVE

MIN.1,5”

FLIGHT SPOILER 2, 3, 6 AND 7

Figur Figure e 30

RIGHT MAIN LANDING GEAR

GROUND SPOILER 0, 1, 4, 5, 8 AND 9

Spoi Spoile lerr Mix Mixer er Contr Control ol Sche Schema mati tic c



THIS PAGE INTENTIONALLY LEFT BLANK


B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

SPEED BRAKE LEVER DOWN

SPEED BRAKE LEVER INTERMEDIATE

SPEED BRAKE LEVER UP y l n O s e s o p r u P g n i n i a r T r o F

Figu Figure re 31

Spoi Spoile lerr Defle Deflecti ction on Sche Schema matic tic



27-60

SPOILER RATIO CHANGER Purpose The ratio changer varies the magnitude of the output to the spoiler mixer for a given magnitude of input from the aileron system, depending on speed brake lever setting. The output decreases as speed brakes are raised. Location The spoiler ratio changer is mounted on the forward bulkhead of the right wheel well. Physical Description/Features The ratio changer case supports spoiler output quadrants and a speedbrake quadrant. A ratio changer input rod from the spoiler control quadrant connects to a bell crank mounted on the ratio changer case. Two rollers engaged in a slot in the bellcrank are connected to a lever and a link. The lever is connected to a crank that drives the spoiler mixer cam. The link is connected through a lever and two part link to the speedbrake quadrant. Operation Rotation of the speed brake quadrant moves the rollers in the bellcrank. Moving the speed brake quadrant to raise the spoilers will cause the rollers to move toward the bell crank pivot. This will cause a decrease in the magnitude of the ratio changer output for a given input.


B737-300/-400/-500

RATIO CHANGER INPUT ROD RIGGING PIN HOLE

RIGHT SPOILER OUTPUT QUADRANT WSA

SPEED BRAKE INPUT QUADRANT LEFT SPOILER OUTPUT QUADRANT

WSB

WSA

RIGGING PIN HOLE WSB



B737-300/-400/-500 27-60

PIVOT POINT STRUCTURE PIVOT RATIO CHANGER INPUT ROD BELL CRANK

SHAFT AILERON SPRING CARTRIDGE

CABLES SBA = SPEED BRAKES UP SBB = SPEED BRAKES DOWN WSA = SPOILERS UP WSB = SPOILERS DOWN

ROLLERS TWO PART LINK

SPOILER CONTROL QUADRANT

AILERON BUS DRUM LEVER

SBA

SPEED BRAKE INPUT QUADRANT

RIGGING PIN HOLE

TO GROUND SPOILER CONTROL VALVE

SBB

UP

WSA

UP WSA

RIGGING PIN HOLE


LEFT SPOILER OUTPUT QUADRANT

RIGHT SPOILER OUTPUT QUADRANT

WSB

WSB UP

UP RIGGING PIN HOLE

Figu Figure re 32

AILERON CAM

Spoi Spoile lerr Mi Mixe xerr Sch Schem emat atic ic


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS FLIGHT SPOILER CONTROL QUADRANTS Purpose The spoiler control quadrants operate the flight spoiler hydraulic actuators in response to cable inputs from the spoiler mixer linkage. Location One control quadrant for each flight spoiler is bracket mounted on the hydraulic actuator support fitting on the rear wing spar. Physical Description/Features Two cables from the ratio changer output quadrants are attached to each spoiler actuator quadrant. Each quadrant pivots on two sealed bearings to drive a crank connected to it by two shear rivets. The crank is connected to the ad justable input arm by an input link. Operation Commands from the spoiler mixer and ratio changer are transmitted to the actuator quadrant by cables. Rotation of the quadrant operates the input arm to port hydraulic pressure inside the actuator and position the spoiler panel. Maintenance Practices Spoiler pickup, that amount of pilots control wheel rotation when the spoilers start moving up, is set by adjusting a hex head adjusting bolt on the input arm. It is set at 11 control wheel rotation. Clearance between spoiler and trailing edge flap is adjusted at the rod end.


B737-300/-400/-500 27-60



B737-300/-400/-500 27-60 FLIGHT SPOILER

SPOILER TO FLAP CLEARANCE +0.10 0.03 INCH -0.00 1

RUB STRIP 1

MEASURE THE CLEARANCE BETWEEN THE RUB STRIP AND THE MIDFLAP AT THE NEAREST POINT

CHECKNUT LOCKING KEY ACTUATOR LINK

PISTON ROD ACTUATOR

CHECKNUT


ADJUSTMENT SCREW INPUT ARM ACTUATOR PISTON ROD

ACTUATOR INPUT ARM

Figur Figure e 33

Flig Flight ht Spoi Spoile lerr Cont Contro roll Qua Quadr dran ants ts


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS FLIGHT SPOILER ACTUATOR Purpose A flight spoiler hydraulic actuator is installed to position each of the flight spoilers. Location Each actuator is trunnion mounted to a support fitting on the rear wing spar beneath each spoiler panel. Physical Description/Features Each of the identical actuators includes a cylinder, actuator rod and piston, control valve, filter, relief check valve, blow-down check valve, extension check and thermal relief valve, snubber check valve and overtravel pistons. Hydraulic pressure and return lines are connected to the actuator through trunnion fittings. The piston rod end of the actuator is attached to the spoiler. Control With hydraulic pressure turned off, the over travel pistons are relaxed preventing the input lever from transmitting motion to the control valve thru the internal crank. Springs reseat the extension check and thermal relief valve, trapping hydraulic pressure on the actuator, to prevent spoiler float. Excess pressure on the down side of the piston will open the thermal relief valve and relieve this pressure. When hydraulic pressure is applied, the overtravel pistons are pressurized to clamp the internal crank. This allows rotation of the external input lever to position the main control valve. The extension check and thermal relief valve is also pressurized open.


Operation With the spoilers retracted and the piston in the full down position, an up signal to the control valve allows hydraulic fluid to flow to the base of the cylinder and through the snubber check valve to the piston. The piston is extended to position the spoiler panel up. The snubber check valve prevents a hydraulic lock when the piston is in the full down position. The blow-down check valve in the pressure line allows the spoilers to blow down at critical speeds.

B737-300/-400/-500 27-60



FILTER

B737-300/-400/-500 27-60

BLOW-DOWN CHECK VALVE

THERMOSTAT VALVE

RELIEF CHECK VALVE EXTENSION CHECK AND THERMAL RELIEF VALVE

>>

CONTROL VALVE

SPOILER UP OVERTRAVEL PISTONS


PRESSURE PORT

CYLINDER

RETURN PORT

SNUBBER CHECK VALVE

INPUT LEVER (EXTERNAL) RETRACT

EXTEND

Figu Figure re 34

Flig Flight ht Spo Spoil iler er Actu Actuat ator or

ACTUATOR ROD AND PISTON



27-60

SPOILER SHUTOFF VALVE Purpose Spoiler shutoff values are installed in the flight controls hydraulic module to control hydraulic power to the flight spoiler actuators. Location Spoiler shutoff valves are located on the left side of each flight controls hydraulic module at the left (A system) and right (B system) outboard sides of the main wheel well forward wall. Physical Description/Features The spoiler shutoff valve is identical to the flight controls shutoff valve. It is a spool valve mounted in a cavity in the flight controls hydraulic module, attached by four bolts with an electric motor attached to the valve by four bolts. The motor is splined to a cam which converts rotary motor action to linear spool travel. The valve is equipped with a manual override lever and position indicator. Power The spoiler shutoff valves are powered electrically by 28 volts dc from bus 2 through a single circuit breaker on P6. P6. Control The valves are independently controlled electrically by either the spoiler A or spoiler B switch on the forward overhead panel. Either valve can also be controlled manually by the manual override lever.


B737-300/-400/-500

Operation Both spoiler shutoff valve switches are normally guarded ON. The 28 volts dc is applied to the open windings of the valve, opening it to pass hydraulic pressure to the flight spoiler actuators system A to 3 and 6 and system B to 2 and 7. Moving the switch OFF applies power to the close windings to drive the valve closed and remove hydraulic pressure from the respective actuators. Monitor There are no indicator lights associated with the flight spoiler shutoff valves. Valve position can be monitored by the position lever. Valve ON (OPEN) is position 1 and valve OFF (CLOSED) is position 2. Maintenance Practices Spoiler shutoff valves and flight controls shutoff valves are interchangeable. The valve and motor can be replaced as a unit or the motor can be replaced separately.



FLT CONTROL A

B737-300/-400/-500 27-60

STANDBY B

HYD

A ON

B ON

OFF

OFF

STDBY RUD

LOW QUANTITY LOW PRESSURE

STDBY RUD ALTERNATE FLAPS ARM LOW PRESSURE

UP

LOW PRESSURE

OFF SPOILER

ON

A

B

2

1

OFF

OFF

DOWN

FLIGHT SPOILER ACTUATOR (TYPICAL) ON

OFF

FEEL DIFF PRESS SPEED TRIM FAIL

YAW DAMPER

MACH TRIM FAIL

YAW DAMPER

AUTO SLAT FAIL

1 2

OFF ON

HYDRAULIC SYSTEM”B”

HYDRAULIC SYSTEM”A” PRESS RET


PRESS

RET

SPOILER SHUTOFF VALVE

FLIGHT CONTROLS SHUTOFF VALVE

5

6

5

6

3

7

3

7

4

Figu Figure re 35

2

Flight Flight Spoi Spoile lerr Hydra Hydraul ulic ic Syst System em

4

2


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS GROUND SPOILER CONTROL VALVE VALVE Purpose A ground spoiler control valve is installed to control hydraulic pressure to the ground spoiler actuators. Location The valve is bolted to a bracket on the outboard side of the right main gear wheel well forward wall. Physical Description/Features The valve consists of a piston and sleeve assembly enclosed in a valve body that has five external ports connected to drilled passages in the sleeve. A lever which pivots about a lug on the valve body is connected to the piston. The opposite end of the lever is connected by a pushrod to a crank on the spoiler mixer. Control Moving the speedbrake control lever in the flight compartment rotates the speedbrake input quadrant on the spoiler mixer. This operates the spoiler mixer linkage to transmit motion through the crank and pushrod to the control valve.


Operation When the speed brake control lever is in the DOWN position, the valve ports hydraulic pressure directly to the down ports of the ground spoiler actuators. As the control lever begins to move aft, the valve moves through a neutral position where the pressure port is blocked and all other ports are connected to return. As the speed brake [ever is moved further aft, the control valve supplies pressure to a ground spoiler bypass valve in the line to the up ports of the ground spoiler actuators. The actuator down ports are connected to return through the control valve.

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

GROUND SPOILER ACTUATOR LINKAGE TO GROUND SPOILERS NO. 8 AND 9

TO GROUND SPOILERS NO. 0, 1 AND 4 GROUND SPOILER NO. 5

GROUND SPOILER ACTUATOR

RIGHT LANDING GEAR

UP TO SYSTEM A RETURN

DOWN

FROM SYSTEM A PRESSURE GROUND SPOILER CONTROL VALVE

FLEXIBLE DRIVE ROD SPOILER MIXER


RATIO CHANGER GROUND SPOILER INTERLOCK VALVE UPPER TORSION LINK

SBB SPEED BRAKE INPUT QUADRANT

FWD

Figu Figure re 36

CONTROL CABLES

SBA TO FORWARD DRUM MECHANISM

Groun Ground d Spoi Spoile lerr Contr Control ol Valve alve


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS GROUND SPOILER IINTERLOCK VALVE VALVE Purpose The ground spoiler interlock valve is provided to limit the use of ground spoilers to ground operation. Location The valve is secured to a bracket on the aft side of the rear wing spar above the right main gear trunnion. Physical Description/Features The valve consists of a three port housing with passages for hydraulic fluid and a valve slide. The slide is actuated by a bellcrank that is connected by a teleflex cable to the upper torsion link. Control The ground spoiler interlock valve is controlled through the teleflex cable by extension and compression of the right main gear shock strut. Operation With the landing gear retracted the shock strut inner cylinder is fully extended. The position of the upper torsion link closes the bypass valve and blocks pressure from reaching the ground spoiler actuator up ports. With the landing gear extended, the valve continues to block hydraulic pressure until the right main gear shock strut is compressed by contacting the ground. As the inner cylinder is compressed, the upper torsion link moves the teleflex cable to open the bypass valve and connect the ground spoiler control valve to the up port of the actuators.


Maintenance Practices Adjustment and test of the ground spoiler bypass valve must be accomplished with the airplane jacked so that the right main gear shock strut is fully extended

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60 GROUND SPOILER ACTUATOR

SPOILER NO. 0

SPOILER NO. 1

SPOILER NO. 4

GROUND SPOILER CONTROL VALVE SEE

SPOILER NO. 5

SPOILER NO. 8

SPOILER NO. 9

LANDING GEAR SAFETY SENSOR ACTUATORS

1 3 2 5 1

A


4 3 2

SYSTEM A PRESSURE SEE

FWD

B

SYSTEM A RETURN INTERLOCK VALVE CABLE 5

4

5

4

3


1

SPOILERS DOWN

2

3

1

2

SPOILERS UP


1

3

1

2

INTERLOCK VALVE YOKE

3

2

OLEO COMPRESSED

OLEO EXTENDED

HYDRAULIC LINE

HYDRAULIC TUBES


A

B

B

Figur Figure e 37

Grou Ground nd Spo Spoil iler er Inter Interlo lock ck Valve alve



FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS OUTBOARD GROUND SPOILER ACTUATOR Purpose One ground spoiler actuator is installed to position each outboard ground spoiler panel. Location The actuator housing is connected to a support on the rear wing spar and the piston rod end is connected to the spoiler panel. Physical Description/Features Each actuator contains an internal mechanical lock which locks the actuator in the retract position. A spring maintains this lock when hydraulic pressure is removed. The bearing mounted actuator housings and swivel joints in the hydraulic lines allow the actuators to rotate as the spoilers are positioned. Operation With the piston retracted, locking keys are springloaded to the locked position. Directing extend hydraulic pressure to the actuator compresses the locking piston against the spring and allows the keys to move inward. This unlocks the piston and allows it to extend. The spoiler panel extends to 60" from the down position. Outboard ground spoiler panel travel is limited by the stroke of the hydraulic actuator. There are no mechanical stops.


B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

ACTUATOR PISTON ROD END SEE A

SPOILER TO FLAP CLEARANCE +0.10 0.03 INCH -0.00

GROUND SPOILER

1

FLAP

RUB STRIP

LOCKING KEY RETRACT PORT TAPERED LOCKING PISTON


PISTON AND ROD ASSEMBLY

LOCKING KEY EXTEND PORT LOCKING SPRING

1

MEAS MEASUR URE E THE THE C CLE LEAR ARAN ANCE CE BET BETWE WEEN EN THE THE RUB RUB STR STRIP IP AND THE MIDFLAP AT THE CENTER OF SPOILER PANEL

Figur Figure e 38

Outbo Outboar ard d Grou Ground nd Spoi Spoile lerr Actua Actuator tor


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS INBOARD GROUND SPOILER ACTUATOR Purpose Two identical actuators and actuator linkages are used to position each inboard ground spoiler. Location Each actuator housing is connected to a support on the rear wing spar and the piston is attached to the linkage. Physical Description/Features Each actuator contains an internal mechanical lock, similar to that described for the outboard actuators, that locks the actuator in the extend position (spoiler down). The actuator linkage consists of an idler crank and push rod. Operation When up hydraulic pressure is directed to the actuator the mechanical lock is released, allowing the piston rod to retract. This rotates the idler crank and drives the push rod to raise the spoiler panel. The spoiler panel rises to 60 " above the down position. Maximum extension is limited by the stroke of the hydraulic actuator. There are no mechanical stops.


B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

IDLER CRANK

1 GROUND SPOILER ACTUATOR

FLAP

TAPERED LOCKING PISTON


GROUND SPOILER

PUSHROD

LOCKING SPRING

EXTEND PORT

PISTON AND ROD ASSEMBLY

LOCKING KEY RETRACT PORT

1

RUB STRIP

SPOILERS NO. 4 AND 5 SPOILER TO FLAP CLEARANCE

Figu Figure re 39

Inbo Inboar ard d Spoi Spoile lerr Act Actua uato torr


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS SPEED BRAKE LEVER ACTU AT ATOR OR Purpose The speed brake lever electric actuator is installed to automatically actuate the ground speed brakes to aid braking during the landing roll. Location The speed brake lever actuator is connected between structure in the lower nose compartment and a support bracket on the no-back brake mechanism. Physical Description/Features The actuator consists of a 28 volt dc electric motor, gearing, a displacement rod, nonjamming mechanical stops and adjustable limit switches. The limit switches are bench adjusted to limit the stroke at the extended and retracted position. The gearing prevents actuator motion due to external loads.

SPEED BRAKE ARMING SWITCH (S 276) The speed brake arming switch (S 276) - is mounted mounted on the left ceiling ceiling area, area, in the lower lower nose compartme compartment. nt. - is operated operated by the movement movement of the the speed brake brake lever (armed) (armed) - switches switches electrical electrical power power to the automatic automatic speed brake brake system. - is a signal signal to terminat terminate e (disarm) (disarm) the the autobrake operation, when it’s active, if the speed brake lever was restowed to the down detent position.


B737-300/-400/-500 27-60



B737-300/-400/-500 27-60 SPEED BRAKE INDICATOR DETENT PLATE SEE B B

DOWN

ARMED

SPEED BRAKE CONTROL LEVER SEE A

CONTROL CABIN FLOOR

FLIGHT

A

DETENT

SPEED BRAKE FORWARD DRUM ROLLER

UP

ACTUATOR

A

ARMED DETENT



MOUNTING SCREWS

CAM ROLLER LEVER ARM

CAM

CAM

FLIGHT DETENT

FORWARD DRUM QUADRANT

SPEED BRAKE ARMING SWITCH, S276

A-A

Figu Figure re 40

Spee Speed d Brak Brake e Leve Leverr Ac Actu tuat ator or

A

FW FWD

INBD

B



FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS RTO SWITCH (S 650) The RTO (rejected takeoff) - Switch (S 650) - is located located on the left left side of of the control control stand. stand. - will be switched switched by movemen movementt of the thrust thrust revers levers levers (48 ) into the reverse position. - energizes the RTO - Relay (electronic (electronic compartment) to supply the automatic speed brake control module.

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

SPEED BRAKE LEVER

LOCATION E+E COMPARTMENT E 3-2

RTO - RELAY

ENGINE NO. 1 THRUST DRUM CAM FOLLOWER

RTO SWITCH SEE

ENGINE NO. 2 THRUST DRUM CAM FOLLOWER

A

SPEED BRAKE INDICATOR PLATE SPEED BRAKE DOWN DETENT

SWITCH ACTUATING ARM

SPEED BRAKE LEVER LIFTING ARM LIFTING TAB


RTO ACTUATION SHAFT (SPLINED) THRUST DRUM CABLE GUARDS

FWD

A

Figu Figure re 41

RTO RTO Swi Switc tch h Loca Locati tion on




B737-300/-400/-500 27-60

AUTOMATIC SPEED BRAKE LEVER ACTUATOR Power Electrical power to activate the speed brake lever actuator is 28 volts dc from Bus number 2 through a circuit breaker on P-6. Control When the speed brake lever is in the down detent no electrical power is provided to the lever actuator electrical circuit. Moving the lever to the ARMED position actuates S-276, by a cam attached to the forward quadrant, and supplies electric power to the system. After landing with both throttles retarded, power is applied to the raise side of the actuator motor when the wheels spin up above 60 knots. A ground signal from air /ground sensing will provide the same same result in the absence of the wheelspeed wheelspeed signal. Electrical power can also be provided to the control circuit when the thrust reverse levers are moved to reverse by activating RTO switch S-650. During the takeoff roll power is applied to the actuator motor only when wheelspeed is above 60 knots. Input/Output The primary input to the electrical circuit in the auto speed brake module that will power the speed brake lever actuator to raise the lever after arming is from the anti skid system. A wheelspeed above 60 knots signal is required from one wheel of each of these pairs; left inboard or right outboard and left outboard or right inboard. The alternate input to raise the lever is from the E11, landing gear logic shelf. Both the air sensing relay and the ground sensing relay must be energized to the ground mode. This alternate signal is removed 4 seconds after ground is sensed. The input to remove a raise signal and energize the lower side of the actuator motor is from switches that are activated by advancing either throttle or depressing test pushbutton 3. Output of the auto speed brake module is to the raise or lower contacts of the lever actuator motor.

Operation The pilot positions the speed brake lever to ARM prior to landing for automatic speed brake operation. This resets K6 and K7, the air/ground relays inside the auto speed brake module. After landing, wheelspeed signals from the antiskid system or air/ground sensing causes the actuator to drive the lever to the UP position. All spoiler panels respond to speed brake lever position. The air/ground signal times out 4 seconds after ground is sensed. The speed brakes are lowered and the actuator is reset by the pilot advancing either throttle after rollout. During the takeoff roll the speed brake lever is in the down detent. If the pilot selects reverse thrust, a cam pushes the lever out of the down detent and activates S-650. The actuator will raise the speed brakes if wheelspeed is above 60 knots. Monitor four blue lights on the front of the auto speed brake module are connected to the four wheel speed relays (K1 through K4). The lights illuminate whenever the respective relay is energized.




B737-300/-400/-500 27-60

Figu Figure re 42

Autom Automat atic ic Spee Speed d Brak Brake e Syst System em


FLIGHT CONTROLS SPOILER, DRAG DEVICES AND VARIABLE AEROD YNAMIC FAIRINGS AUTOMATIC SPEED BRAKE INDICATION Purpose The speed brake indicating system monitors the automatic speed brake control system for electrical failures. Location Two indicator lights are located on the center instrument panel, P2. The lights are controlled by circuits inside the auto speed brake module installed on the E3 rack in the Electronic Equipment Compartment. Three indicating system test switches are installed on the center instrument panel, P2. Physical Description/Features


A green SPEED BRAKE ARMED light illuminates to indicate no faults in the automatic ground speed brake electrical system. - No speed speed brake test button button pushed pushed - and - At least one (inboard (inboard or outboard outboard)) antiskid antiskid channel ON and operativ operative e - and - Speed brake brake lever lever actuator actuator in in lower positio position n - and - Speed brake brake electrica electricall circuitry circuitry ready ready to operate. operate. The amber SPEED BRAKE DO NOT ARM light illuminates when an electrical failure is detected. - Any speed speed brake test button button pushed pushed - or - Total antisk antiskid id OFF OFF or inopera inoperative tive - or - Speed brake brake lever lever actuator actuator not not in lower lower position position - or - Speed brake brake electri electrical cal circuit circuit malfunc malfunction. tion.

B737-300/-400/-500 27-60



B737-300/-400/-500 27-60

SPEED BRAKE y l n O s e s o p r u P g n i n i a r T r o F

(g)

ARMED

SPEED BRAKE DO NOT ARM (a)

Figur Figure e 43

Autom Automati atic c Spee Speed d Bra Brake ke Indi Indica catio tion n



27-60

AUTOMATIC GROUND SPEED BRA KE OPERATION Operation Electrical power for the control circuit inside the auto speed brake module is 28 volts dc from bus 2 through the auto speed brake circuit breaker. The lights are powered from the same circuit breaker through R87, auxiliary master dim relay. Control The indicating system is activated when the speed brake lever is moved to the ARMED position. Inputs to the indicating system are from the lower contact of the speed brake lever actuator, antiskid system failure monitoring and the auto speed brake control circuitry. Output is to the green or amber light as appropriate. The lights are interlocked so that only one light can be illuminated at a given time.


B737-300/-400/-500

Normal Sequence The green SPEED BRAKE ARMED light illuminates whenever the speed brake lever is moved to armed, at least one anti skid channel is on and operative, and none of the monitored electrical failures exist. The amber SPEED BRAKE DO NOT ARM light illuminates whenever the speed brake lever is moved to armed and at least one of the following conditions are present: - Speed brake lever actuator is is not fully retracted to the the speed brake lower position. - Both anti anti skid channe channels ls either either off or or inoperativ inoperative e - Electrical power in the distribution (BUS) system. Most likely at the normally open contacts of K-1 and K-3. - Ground available available in the the distribution (BUS) system. Most likely at the normally open contacts of K-2 and K-4.

Pilot response to an amber light is to return the speed brake lever to the down detent and manually operate the speed brakes after landing. Operation of the indicating system can be tested using the three pushbuttons on the center panel. The speed brake lever must be moved to armed and a green light illuminated to conduct the test. Proper light response to depressing one of the pushbuttons is green light extinguishes and amber light illuminate’. Release the pushbutton and the lights return to normal. Test switch 1 simulates a short to power failure of a wheel speed relay (K1 or K3). Test switch 2 simulates a short to ground failure of a wheelspeed relay (K2 or K4). Test switch 3 simulates a throttle advanced and applies power on the BUS.




B737-300/-400/-500 27-60

LIGHT CIRCUIT

Figur Figure e 44

Automa Automati tic c Spee Speed d Brak Brake e Circ Circui uitt


FLIGHT CONTROLS RUDDER & TAB

27-20

27-20

RUDDER & TAB

INTRODUCTION Purpose The rudder provides yaw control of the airplane around the vertical axis. System Description A single conventional rudder without tab is powered by a main power control unit supplied by hydraulic systems A and B. A separate power control unit supplied by the standby system provides backup power. Any one of the three hydraulic systems will provide effective rudder control. The power·control units are actuated by cables operated from either the captains or first officers rudder pedals. Rudder trim is accomplished by operating a trim control switch on the aisle stand that inputs to the power control units via an electric actuator mounted on the feel and centering mechanism.


B737-300/-400/-500



B737-300/-400/-500 27-20

TORQUE TUBE

RUDDER FEE L AND CENTERING UNIT

RUDDER AFT CONTROL QUADRANT

SEE A


CONTROL STAND

RUDDER TRIM KNOB

A

Figu Figure re 45

Rudd Rudder er & Tab Tab Int Intro rodu duct ctio ion n



27-20 RUDDER SYSTEM General Component Locations The rudder is mounted on the aft side of the vertical stabilizer. The power control units and aft control components are installed inside the vertical stabilizer, forward of the rudder. The rudder system is operated from the flight compartment by either the rudder pedals or the rudder trim switch. General Subsystem Features The rudder is operated by hydraulic power only. There is no manual reversion capability. capability. Normal operation is with both system A and system B applied to the main power control unit. The standby system can be applied through the standby power unit when it is selected. No more than two hydraulic systems can power the rudder at a giver time. Wind gust protection for the rudder is hydraulic dampening. Maximum rudder travel is limited by actuator piston stroke. System Interfaces The rudder is also controlled by the yaw damper system to prevent dutch roll. This system operates through the system B hydraulic control section of the main power control unit. The yaw damper system operates independently of the rudder control system and does not result in feedback at the rudder pedals. The yaw damper system will be covered in the autoflight block of instruction.


B737-300/-400/-500

General Operation Control of the rudder is by either set of rudder pedals. They operate forward quadrant assemblies which move cables to rotate a torque tube that drives input control rods to the power control units and to the rudder feel and centering mechanism. The pressurized power control unit will then actuate, deflecting the rudder. Rudder trim is accomplished by rotating a switch on the aisle stand. This operates an electric actuator that drives the rudder feel and centering mechanism to rotate the torque tube and actuate the input rods. The pressurized power control unit deflects the rudder.



B737-300/-400/-500 27-20

RUDDER STANDBY RUDDER ACTUATOR

STANDBY RUDDER ACTUATOR

SEE B

RUDDER POWER UNIT


STANDBY INPUT ROD

EXTERNAL SUMMING LEVER

FEEL AND CENTERING MECHANISM TRIM AND FEEL ROD

RUDDER TRIM ACTUATOR RIG PIN

INPUT LINK INPUT CRANK RUDDER POWER UNIT

RA

INPUT ROD

RB

TORQUE TUBE

RUDDER CONTROL CABLES

CONTROL ROD LOWER CRANK (BELL CRANK) FEEL AND CENTERING MECHANISM CRANK


FEEL AND CENTERING MECHANISM

RUDDER TRIM ACTUATOR RB RA

Figu Figure re 46


B

Rudd Rudder er Cont Contro roll Syst System em

FWD



27-20 RUDDER SYSTEM PRESSURE REDUCER General To lessen the effects of large rudder deflections, a hydraulic rudder pressure reducer in the A hydraulic system supply line to the main rudder PCU is installed. This will reduce available rudder authority by approx. 20 percent during those phases of flight when large rudder deflections are not required. This reduced authority will enhance safety by reducing the airplane’s reaction to full rudder inputs and allow the ailerons and spoilers to be more effective in counteracting an excessive rudder input. Together with the rudder pressure reducer, a new yaw damper coupler is installed, too. Failure of the new yaw damper coupler results in the yaw damper coupler to disengage. Physical Description The rudder pressure reducer is connected to the A system hydraulic line upstream of the main rudder PCU. Hydraulic pressure to the rudder is reduced to 1000 psi when the airplane climbs above 1.000 ft AGL (CM1 RA). Hydraulic pressure returns to normal when - the airplan airplane e descends descends through through 7 700 00 ft AGL, AGL, or - lf B hydrau hydraulic lic system system depres depressurize surizes, s, or - whenever whenever the N1 differenc difference e between the the left and right engine engine exceeds exceeds 45 percent.


B737-300/-400/-500

If the rudder pressure reducer valve falls to switch back to the full pressure mode, the FLT CONTROL LOW PRESSURE light will illuminate. To reduce nuisance annunciations, a 5 second delay is incorporated before the yaw damper coupler will sense to illuminate or extinguish the light.



B737-300/-400/-500 27-20 STA 1104 FWD

NORMAL POSITION


HIGH PRESSURE POSITION

Figu Figure re 47

Rudd Rudder er Pres Pressu sure re Redu Reduce cerr



B737-300/-400/-500 27-20






B737-300/-400/-500 27-20

Figur Figure e 48

Rudde Rudderr Pre Press ssure ure Reduc Reducer er Circ Circui uitt



27-20 RUDDER PEDAL ASSEMBL ASSEMBLY Y Purpose The captain and first officer are each provided with a pair of rudder pedals used for controlling the airplane about the vertical axis. Location The rudder pedals are located below the captain’s and first officer’s instrument panels. Rudder pedal support and quadrant assemblies are in the lower nose compartment. Physical Description Each pair of pedals consist of right and left pedals mounted on a shaft. The pedal shaft is attached to the upper end of the pedal arm assembly. The lower end of the pedal arm assembly is mounted on a support shaft attached to structure below the floor. Fore and aft movement of the pedals is transmitted by two pushrods to a jackshaft yoke. The rotary motion of the yoke is passed to the forward quadrant by means of the jackshaft. Both sets of control pedals respond equally because they are bussed together by means of a pushrod connecting the two jackshaft assemblies.


B737-300/-400/-500

Control Each set of rudder pedals can be adjusted independently by means of a rudder pedal adjust knob located on the instrument panel, forward of the control wheel. The knob must be pulled aft to permit crank rotation. Rotation of the crank operates a flexshaft to drive a jackscrew which moves the yoke fore and aft. Rudder pedal adjustment crank and crank handle stops are installed to prevent the rudder pedal adjustment screw from being backdriven by heavy foot pressure applied simultaneously to both rudder pedals.



B737-300/-400/-500 27-20

RUDDER STANDBY RUDDER ACTUATOR RUDDER AFT CONTROL QUADRANT

RUDDER POWER UNIT

FEEL AND CENTERING MECHANISM RUDDER TRIM ACTUATOR

FORWARD RIGGING PIN HOLE

RUDDER PEDAL ADJUSTMENT CRANK

AFT RIGGING PIN HOLE FORWARD QUADRANT BUSROD (TO FIRST OFFICER’S PEDALS)

RUDDER PEDAL ADJUSTMENT SHAFT RA RB SEE A

RUDDER CONTROL CABLES

CAPTAIN’S RUDDER PEDALS PUSHROD TO NOSE WHEEL STEERING

BRAKE PUSHROD (TYPICAL)


RIGGING PIN HOLE (R-1 ON CAPTAIN’S PEDAL ARM, R-2 ON FIRST OFFICER’S PEDAL ARM)

A

Figu Figure re 49

Rudd Rudder er Peda Pedall Asse Assemb mbly ly



27-20 AFT RUDDER CONTROL COMPONENTS Purpose Aft rudder control components transmit rudder pedal input to the hydraulic power control units. Location The aft rudder control components are mounted in the vertical stabilizer forward of the power control units. Physical Description The aft control components consist of a cable quadrant and a torque tube that provides a dual load path for rudder control linkage inputs. These cranks are bolted to the tube. The lower crank is connected to the input rod from the aft control quadrant and to the feel and centering mechanism. The center crank is connected to the main rudder power control unit input linkage. The upper crank is connected to the standby power control unit input linkage. Operation Rudder pedal input through cables actuates the aft control quadrant which rotates the torque tube via an input rod. This results in a simultaneous input to the main power control unit, standby power control unit, and the feel and centering unit. Rudder trim is input through the feel an centering unit which rotates the torque tube. Trim inputs to both power control units simultaneous with driving the aft control quadrant to position the rudder pedals.


B737-300/-400/-500



B737-300/-400/-500 27-20

STANDBY POWER INPUT ROD

SEE

A UPPER CRANK TORQUE TUBE TRIM AND FEEL ROD CENTER CRANK

AFT QUADRANT RUDDER POWER CONTROL UNIT INPUT ROD LOWER CRANK (BELL CRANK) QUADRANT INPUT ROD


CABLE RB

A CABLE RA

Figu Figure re 50

Aft Aft Rudde Rudderr Contr Control ol Comp Compon onen ents ts



27-20 MAIN RUDDER POWER CONTROL UNIT Purpose The main rudder power control unit moves the rudder right or left when actuated by rudder pedal input, rudder trim input, or yaw damper input and provides wind gust snubbing when the airplane is parked. Location The unit is located in the vertical fin, The body is fixed to fin structure and the piston head to the rudder. Physical Description The main power control unit is a single tandem actuator with two pistons on a single rod. The unit contains two separate chambers for the two hydraulic systems, two bypass valves, and a dual control valve operated by an internal input crank. The internal crank is operated by an external input crank that is driven by an external summing lever connected to the input rod from the torque tube. Yaw damper components incorporated in the main power unit include a solenoid operated shutoff valve, a transfer valve, a yaw damper actuating piston, and a rate sensor. Power Normal operation of the main power control unit is by both hydraulic systems A and B. Either system acting alone will provide full rudder control.


B737-300/-400/-500

Control Hydraulic power from an operating system opens a bypass valve and is delivered to the control valve. When a system is off the bypass valve is springloaded to bypass. Both sides of the piston are connected to either line from the control valve to prevent a hydraulic lock.

Operation Input from the rudder pedals or rudder trim causes the torque tube to drive an input rod that positions the power unit input crank through the external summing lever. The external crank actuates the internal crank to position the control valve which ports hydraulic pressure to one side of the actuating pistons. The piston strokes to position the rudder. The external summing lever is carried by the piston to return the external crank to neutral and stop the rudder at the desired position. The amount of control valve movement is also governed by yaw damper input. Pilot input and yaw damper input are summed algebraically by the summing levers connected to the primary and secondary control valves. Pressure from the transfer valve drives the yaw damper actuator to position the control valve via the summing levers. Yaw damper is limited to 3 left or right rudder movement.



B737-300/-400/-500 27-20 TRANSFER VALVE

CAGING SPRING

YAW DAMPER SHUTOFF VALVE PRIMARY VALVE PB

ELECTRICAL CONNECTOR

SECONDARY VALVE

P C RB

TANDEM ACTUATOR

C R

YAW DAMPER POSITION TRANSMITTER

BYPASS VALVE

COMPENSATOR SPRING

SECONDARY SUMMING LEVER COMPENSATOR

YAW DAMPER ACTUATOR


PRIMARY SUMMING LEVER

RA INPUT SHAFT

Figu Figure re 51

INTERNAL INPUT CRANK

PA THERMOSTATIC VALVE OR PLUG (OPTIONAL)(2 PLACES)

Main Main Rudd Rudder er Powe Powerr Unit Unit



27-20 RUDDER FEEL AND CENTERING MECHANISM Purpose The rudder feel and centering mechanism provides artificial feel to the rudder pedals, centers the rudder, and transmits trim inputs to the aft control components. Location The eel and centering unit is located below the rudder power unit in the vertical fin. Physical Description The feel and centering unit consists of a support shaft, a feel and centering crank, two frames, an arm and roller, a spring assembly, and a cam. The support shaft is bearing mounted on structure. The feet and centering crank is fixed tc the support shaft and is connected through the trim·and feel rod to the lower crank on the torque tube. The two frames are bearing mounted on the support shaft. The arm and roller and the spring assembly attach to and rotate with the two frames. The cam is fixed to and rotates with the support shaft. The rudder trim actuator is connected to the forward side of the two frames.


B737-300/-400/-500

Operation When the rudder pedals are displaced, the torque tube rotates causing offset of the trim and feel rod and, in turn, rotation of the feel and centering crank, the support shaft, and the cam. As the cam rotates, the arm and roller are displaced out of the detent position to compress the spring assembly and provide artificial feel. Rudder trim input through the rudder trim actuator rotates the two frames with the arm and roller and spring assembly. The force of the spring assembly holds the arm and roller in the cam detent and causes the cam to rotate. Rotation of the cam causes subsequent motion of the support shaft, the feel and centering crank, the trim and feel rod, the lower crank, and the torque tube. This causes input to the rudder power unit to position the rudder and, at the same time, the lower crank causes input to the rudder control system to position the rudder pedals.



B737-300/-400/-500 27-20

CONTROL ROD

RUDDER PEDAL (TYP)

AFT QUADRANT

PILOT’S INPUT

FORWARD QUADRANT

BUS ROD FEEL AND CENTERING UNIT

TRIM ACTUATOR


TORQUE TUBE

FIN STRUCTURE

Figur Figure e 52

Rudde Rudderr Feel Feel and and Cen Cente teri ring ng Mec Mecha hanis nism m

LOWER CRANK (BELL CRANK)



B737-300/-400/-500 27-20

RUDDER TRIM CONTROL SYSTEM RUDDER TRIM INDICATOR

Purpose The rudder trim system provides a means of positioning the rudder for directional trim of the airplane. Location The rudder trim actuator is mounted between vertical fin structure and the case of the rudder feel and centering mechanism. The rudder trim control switch and rudder trim indicator are installed on the aft section of the pilot’s control stand. Physical Description / Features The rudder trim actuator consists of an acme screw driver ram, coupled to a motor-brake and position sensor (RVDT) through a gear reduction, and limit switches enclosed by an aluminum housing. The motor drive is controlled by the rudder trim switch or by activation of an internal limit switch. Monitor The rudder trim indicator is driven electrically by a rotary variable differential transformer (RVDT) in the actuator. It is graduated up to 17 units left and right to display up to the 16 " maximum left and right right rudder movement by the trim system. Loss of power to the indicator causes a black tape to cover the pointer and an off flag to appear at the left of the dial.

RUDDER TRIM KNOB

15

The Rudder Trim Indicator - indicates indicates the the position position of the RVDT RVDT of the actuato actuatorr y l n O s e s o p r u P g n i n i a r T r o F

10

RUDDER TRIM 5 0 5

LEFT

10

15

RIGHT

AILERON

 in units of trim. - may be adjusted adjusted to the rudder’s rudder’s neutral neutral position position by use of an adjustment adjustment screw located at the rear of the indicator module.

Maintenance advice: Rudder trim should only be used when the rudder PCU is supplied hydraulically and the indicators’ electrical power supplied.

CONTROL STAND

RUDDER TRIM INDICATOR

NOSE RIGHT

NOSE LEFT R U

LEFT WING DOWN

RIGHT WING DOWN

D D E R

RUDDER TRIM KNOB



B737-300/-400/-500 27-20

CONTROL ROD FWD FEEL AND CENTERING CRANK

TORQUE TUBE

RIGGING PIN HOLE

RUDDER TRIM ACTUATOR FEEL AND CENTERING UNIT ARM AND ROLLER

CENTERING SPRING SHAFT

CAM ROLLER ARM CRANK

RUDDER CONTROL INPUT FROM QUADRANT CAM ROLLER FIN


RUDDER TRIM ACTUATOR

M1125 RUDDER TRIM ACTUATOR

TORQUE TUBE

Figu Figure re 53

Rudd Rudder er Tri Trim m Sys Syste tem m

FWD



27-20 RUDDER TRIM Power The rudder trim actuator motor is powered by 115 volts ac from transfer bus. The rudder trim indicating system is powered by 28 volts ac from transfer bus 1. Control Operation of the rudder trim switch actuates a pair of contacts. Control power is applied through the upper set of contacts and the appropriate set of limit switches to the motor and brake and to ground through the lower contacts. The electromagnetic brake prevents actuator overrun and internal mechanical stops prevent overtravel in case of limit switch malfunction. Operation Trim commands from the trim switch cause the actuator to extend or retract which rotates the feel and centering mechanism. This drives the power unit input rods to position the rudder to a new neutral if hydraulic power is available. The pedals are positioned to a new neutral position corresponding to rudder position. Monitor The rudder trim indicator is driven electrically by a rotary variable differential transformer (RVDT) in the actuator. It is graduated up to 17 units left and right to display up to the 16 " maximum left and right right rudder movement by the trim system. Loss of power to the indicator causes a black tape to cover the pointer and an off flag to appear at the left of the dial.


B737-300/-400/-500



B737-300/-400/-500 27-20

15

RUDDER TRIM 5 0 5 10

10

LEFT

15

RIGHT NOSE LEFT

AILERON

LEFT WING

RIGHT WING

DOWN

DOWN

NOSE RIGHT

M159 28V AC XFR BUS 1

RUDDER TRIM INDICATOR (P8)

C788 RUDDER TRIM INDICATOR

115V AC XFR BUS 1

C787 RUDDER TRIM CONTROL


NOSE RIGHT

NOSE


LEFT M1126 126 AILE AILERO RON N AND AND RUDDER TRIM CONTROL MODULE (P8)

M1125

Figur Figure e 54

RUDDER TRIM ACTUATOR

Rudde Rudderr Tri Trim m Ele Elect ctri rica call Sche Schema matic tic



27-20 STANDBY RUDDER AC TUATOR Purpose The standby rudder actuator provides standby hydraulic pressure to operate the rudder when either system A, system B, or both are not available. Location The standby rudder actuator is located above the main power control unit in the vertical fin. Physical Description The actuator consists of a bypass valve, control valve, and the actuating cylinder. The piston rod is attached to fin structure and the opposite end of the actuator housing is attached to the rudder. Power The standby actuator is not normally powered. When selected by the A or B flight control switches or automatic operation, the actuator is powered through the standby rudder shutoff valve. At least one side of the main power control unit is not powered when the standby actuator is powered. No more than two hydraulic systems can be used to operate the rudder. Control Inputs form the rudder pedals or trim actuator are simultaneous to the main power control unit and to the standby actuator. The bypass valve is in the bypass position when standby pressure is not available. This connects both piston chambers to the same port of the control valve to prevent a hydraulic lock.


B737-300/-400/-500

Operation When standby rudder operation is activated, standby pressure opens the bypass valve and connects the actuator chambers to separate control valve ports. Control inputs, operating the external crank, position the control valve to apply pressure in one chamber and open the other to return. The actuator housing strokes on the piston to position the rudder and null the control valve.



B737-300/-400/-500 27-20

INPUT CRANK SHAFT

CONTROL VALVE

STANDBY ACTUATOR

R

SEE A BYPASS VALVE (PRESSURE ON CONDITION)

P

RUDDER ATTACH POINT

FIN ATTACH POINT B

FIN STRUCTURE

INPUT CRANK A

STANDBY ACTUATOR SEE B


TORQUE TUBE INPUT ROD INDEX MARKS INPUT CRANK RUDDER

Figu Figure re 55

Stan Standb dby y Rud Rudde derr Actu Actuat ator or

FWD



27-20 STANDBY RUDDER SHUTOFF VALVE Purpose The standby rudder shutoff valve controls hydraulic pressure to the standby actuator. Location The valve is mounted in the standby module on the aft wall of the main gear wheel well. Physical Description The valve is an electrically operated spool and sleeve contained in a cartridge that fits into a cavity in the standby module. The electric motor is splined to a cam which converts rotary motor action into linear spool action in the sleeve. Power The valve is electrically operated by 28 volts dc supplied by the battery bus. Control Operation of the standby rudder shutoff valve electrically is by either the Flight Control A or B switches on the overhead panel. The valve is closed when both switches are ON or OFF. Moving either to STBY RUD applies power to open the valve and allow standby pressure to the standby rudder actuator. The valve can also be opened automatically whenever low pressure is detected at either flight control low pressure switch, the trailing edge flags are not up, and the airplane is either in the air or on the ground with wheelspeed above 60 knots.


B737-300/-400/-500

Monitor The valve is equipped with a manual override lever and position indicator. The lever is at position 1 when the valve is closed and position 2 when open. This lever can be used to manually position the valve when electrical power is off.



B737-300/-400/-500 27-20

LEADING EDGE STANDBY STANDBY RUDDER ACTUATOR SHUTOFF VALVE

DRIVE SHUTOFF VALVE STANDBY RUDDER ACTUATOR SHUTOFF VALVE


STANDBY SYSTEM MODULAR UNIT

AFT BULKHEAD MAIN WHEEL WELL

Figur Figure e 56

Standb Standby y Rudde Rudderr Shutof Shutofff Valve alve



27-20 RUDDER HYDRAULIC SYSTEM Operation/Control Sequence Control inputs to the rudder hydraulic power units are from either pilot’s rudder pedals or the trim control switch. Any input causes rotation of a torque tube in the vertical fin that results in simultaneous input to both the main and standby hydraulic power units. The pressurized power units deflect the rudder up to a maximum of 26" in either direction, 16" by trim. Normal Sequence Normal operation of the rudder is by hydraulic systems A and B applied to the dual tandem piston in the main power unit through shutoff valves in the respective flight control modules. These valves are controlled by the applicable flight control (A and B) switches on the pilot’s overhead panel. The switches are normally guarded ON. When positioned OFF the respective valve closes, blocking system pressure to the aileron, elevator, and rudder. One system can move the rudder full travel but at approximately half the normal rate. Manually operated isolation valves can selectively isolate the rudder from either primary hydraulic system for test purposes.


B737-300/-400/-500

Backup Sequence Either or both primary hydraulic systems can be replaced by the standby power unit, pressurized by the standby hydraulic system through the standby rudder shutoff valve. This valve is operated by positioning either flight control switch (A or B) to STBY RUD. This closes the respective flight control shutoff valve so that no more than two systems operate the rudder at the same time. The standby rudder shutoff valve is automatically opened and the standby pump is started to pressurize the standby actuator whenever either primary flight control low pressure switch is low, the trailing edge flaps are not up, and the airplane is either in the air or on the ground with wheelspeed above 60 knots.



B737-300/-400/-500 27-20

STANDBY MODULAR PACKAGE

RUDDER STANDBY POWER UNIT

M TO ELEVATOR SYSTEM TO STANDBY RESERVOIR

FLT CONTROL A B

FROM STANDBY PUMP

PS

RUDDER STANDBY SYSTEM SCHEMATIC

STANDBY HYD

STDBY RUD OFF

STDBY RUD OFF

A ON

LOW PRESSURE

OFF LOW LOW PRESSURE PRESSURE

UP OFF

OFF

ON

ON

YAW DAMPER YAW DAMPER

PILOTS’ OVERHEAD PANEL

HYDRAULIC FUSE (NOT ON ALL AIRPLANES)

RUDDER POWER UNIT

”B” SYSTEM PRESSURE TO OUTBOARD SPOILERS

M

M

PS

TO AILERON POWER UNIT ”A”

TO AILERON POWER UNIT ”B”

PS

M

M

FEEL DIFF PRESS SPEED TRIM FAIL MACH TRIM FAIL

”A” SYSTEM MODULAR PACKAGE

AUTO SLAT FAIL

OFF ON

R B

CHECK VALVE (NOT ON ALL AIRPLANES)

TO INBOARD SPOILERS

ARM DOWN

OFF

R A

”A” SYSTEM PRESSURE

ALTERNATE FLAPS


P B

LOW QUANTITY

B ON

SPOILER A B

P A

TO ”A” SYSTEM RETURN

COMPENSATOR

Figu Figure re 57

HYDRAULIC ISOLATION SHUTOFF VALVES MODULE

FROM AILERON POWER UNIT ”A”

Rudd Rudder er Hydr Hydrau auli lic c Sche Schema mati tic c

”B” SYSTEM MODULAR PACKAGE FROM AILERON POWER UNIT ”B”

COMPENSATOR TO ”B” SYSTEM RETURN


FLIGHT CONTROLS ELEVATOR & TAB

27-30

27-30

ELEVATOR & TAB

ELEVATOR ELEV ATOR SYSTEM INTR ODUCTION Purpose The elevator system provides primary pitch control about the airplane lateral axis. System Description Two elevators are hinged to the aft end of the left and right horizontal stabilizer sections. The elevators are powered by two independent hydraulic systems in response to either control column inputs or the autopilot. In the event of dual hydraulic failure, a manual reversion mode allows the elevators to be driven directly through a mechanical control system. Three balance panels and a balance tab on each elevator reduce air loads to assist elevator movement, particularly during manual reversion mode. Artificial hydraulic feel is provided at the control column.


B737-300/-400/-500




B737-300/-400/-500 27-30

Figu Figure re 58

Elev Elevat ator or and and Tab Tab Sys Syste tem m Intro Introduc ducti tion on


FLIGHT COPNTROLS ELEVATOR & TAB

27-30 ELEVATOR ELEV ATOR CONTROL SYSTEM General Component Locations Forward control components of the elevator system are located beneath the control columns in the lower nose compartment. Cable runs connect them to the aft control components located inside the tail cone.

General Operation The elevators can be moved by four different inputs: - Pilo Pilott comm comman and d - Auto Autopi pilo lott

General Subsystem Features The two forward elevator control quadrants are connected by a torque tube which provides the same input from either control column. These inputs are provided to cable quadrants mounted on the lower (input) torque tube inside the empennage. Two parallel hydraulic power control units drive the upper (output) torque tube to which the elevators are attached. A feel actuator mounted on the feel and centering unit connected to the lower torque tube receives hydraulic pressure from an elevator feel computer and provides artificial feel on the control columns.

- Mac Mach tri trim m - Neut Neutra rall shif shiftt Normal operation of the elevators is hydraulic, using system A and system B to independent power control units. One operating hydraulic system is sufficient to power the elevators. Manual reversion provisions are incorporated in the event both hydraulic systems are inoperative. Maximum elevator travel is limited by the actuator stroke. Wind gust protection is provided by hydraulic dampening.

System Interfaces The elevators are commanded to move by operation of the stabilizer through neutral shift rods that connect the two systems. Whenever the stabilizer is moved from 3 to 17 units the elevators are gradually moved up. The mach trim system commands the elevators to move at excessive mach numbers. This is accomplished by an actuator mounted on top of the feel and centering unit.


B737-300/-400/-500



B737-300/-400/-500 27-30

SEE

A

AUTOPILOT B ACTUATOR

ELEVATOR TAB ELEVATOR

OUTPUT TORQUE TUBE

NEUTRAL SHIFT SENSOR

EA EB

MACH TRIM ACTUATOR AUTOPILOT A ACTUATOR

ELEVATOR ELEVATOR A PCU ELEVATOR ELEVATOR FEE L AND CENTERING UNIT ELEVATOR B PCU

STABILIZER POSITION SENSOR


INPUT TORQUE TUBE ELEVATOR FORWARD CONTROL QUADRANT

ELEVATOR AFT CONTROL QUADRANTS FWD

A

Figu Figure re 59

Elev Elevat ator or Cont Contro roll Sys Syste tem m

ELEVATOR POSITION SENSOR



27-30 ELEVATOR ELEV ATOR FORWARD FORWARD CONTROL CONTR OL QUADRANTS QUADRANT S Purpose The elevator forward control quadrants are installed to transmit pilot input from the control columns to the aft elevator control components through cables. Location The forward quadrants are suspended from the column connecting torque tube, adjacent to the bottom of each control column, in the lower nose compartment. Physical Description/Feature The elevator forward control quadrants are aluminum alloy forgings. A control column balance weight is attached to each quadrant. The elevator control cables are located in grooves along the lower surface of the quadrants. Each cable is secured at the quadrant. Elevator up cables (EB) run directly aft. Elevator down cables (EA) start forward, run through turnaround pulleys and proceed aft. A pitch force transducer is installed in each forward control quadrant. Transducer operation will be covered in Chapter 22, Auto Flight. Operation Forward and aft motion of either control column is transmitted to the torque tube through mating face splines. This actuates the forward quadrants, tensions the cables, and operates the aft control components to drive the elevators.


B737-300/-400/-500



B737-300/-400/-500 27-30 MICROPHONE SWITCH

MEMORY DEVICE CONTROL WHEEL

AUTOPILOT DISENGAGE SWITCH

STABILIZER TRIM CONTROL SWITCH

CAPTAIN’S CONTROL COLUMN

STALL WARNING CONTROL COLUMN SHAKER WIRE HARNESS

CABIN FLOOR

FIRST OFFICER’S CONTROL COLUMN

DUST COVER ELEVATOR FORWARD CONTROL QUADRANT

QUADRANT TORQUE TUBE ELEVATOR ELEVATOR FORWARD CONTROL QUADRANT


WEIGHT BLOCKS

CONTROL CABLE EA

CONTROL CABLE EB AUTOPILOT PITCH TRANSDUCER

Figu Figure re 60

Elev Elevat ator or For Forwa ward rd Cont Control rol Quad Quadra rants nts



27-30 INPUT TORQUE TUBE Purpose The elevator control input torque tube transmits inputs from the pilots, autopilot, mach trim, or neutral shift to the hydraulic power control units. It also transmits the reaction from the feel and centering unit to the control columns.

ELEVAT ELEV ATOR OR POWER CONTROL UNITS U NITS Purpose Two elevator power control units provide actuation of the elevators in response to inputs from the control columns, autopilot actuators, mach trim, or neutral shift.

Location The input torque tube is the lower of the two torque tubes in the empennage.

Location The control units are mounted vertically in the empennage above the aft elevator control quadrant torque tube.

Physical Description/Feature Elevator aft control quadrants, consisting of four individual segments, are mounted on the input torque tube. The two upper segments are connected to the elevator up cables (EB) and the lower segments to the elevator down cables (EA). Cranks on the input torque tube provide input to the hydraulic power control unit and the feel and centering mechanism.

Physical Description/Feature The upper mounting terminals of the housing attach to lugs on the output torque tube which is directly linked by pushrods to each elevator. The piston rod is attached to structure. Elevator power control units are identical to the aileron power control units.

Operation Pilot input through the cables to the aft quadrants or autopilot actuator inputs rotate the input torque tube. This provides a simultaneous input to both hydraulic rower control unit and to the feel and centering unit. The power control units stroke on their pistons to position the elevators and the feel and centering unit provides artificial feel at the column. Neutral shift or mach trim input rotates the entire feel and centering unit. This rotates the input torque tube resulting in an input to the hydraulic actuators and back through the cables to both control columns.


B737-300/-400/-500

Power The two control units operate independently from separate hydraulic systems, the left unit from hydraulic system A pressure and the right unit from hydraulic system B pressure. Operation Forward or aft movement of the control columns rotates the input torque tube and actuates the power control unit input crank. This displaces the main control valve slide and directs hydraulic pressure to the main actuator piston. The power control unit body extends or retracts and positions the elevator. Movement of the power control unit returns the input crank to neutral and closes the main control valve, which stops the elevator at the desired position. The summed output of the two autopilot actuators is applied to a crank on the input torque tube. This causes an input to the power control units as previously described for a pilot input.



B737-300/-400/-500 27-30

FLEXIBLE HYDRAULIC LINES

OUTPUT TORQUE TUBE

POWER CONTROL UNIT

SEE

A

POWER CONTROL UNIT

SEE

UPPER MOUNTING BOLT

INPUT CRANK CLEVIS

B ELECTRICAL CONNECTOR

SLEEVE

LOWER MOUNTING INPUT ROD BOLT INPUT TORQUE TUBE


MOUNTING BRACKET

FWD FWD

RIG PIN E-5 A

Figure Figure 61

Elevat Elevator or Input Input Torque orque Tube Tube and Power Power Contr Control ol Unit Units s

B



27-30 OUTPUT TORQUE TUBE Purpose The elevator output torque assembly is provided to transmit the motion of the power control units to the elevators. Location The output torque tube is the upper of the two torque tubes mounted in the empennage. Physical Description/Feature The elevator power control units are at!ached to an output torque tube. This is a dual torque tube system, with the inner torque tube made up of two half tubes. The two half tubes are bolted to the outer tube at the middle, therefore, input from either or both power control units will always result in an equal output to both elevators during normal operating function. The secondary pickup provides for a redundant output to the elevator if an inner torque tube should break.


B737-300/-400/-500




B737-300/-400/-500 27-30

Figur Figure e 62

Elev Elevat ator or Outpu Outputt Tor Torque que Tube



27-30 ELEVATOR ELEV ATOR TAB CONTROL MECHANISM Purpose A tab control mechanism is installed for each elevator to position the tab in the direction opposite to elevator movement. Location The mechanisms are located inside the horizontal stabilizer elevator balance bays, aft of the stabilizer rear spar. Physical Description Each tab control mechanism consists of a pushrod and bellcrank assembly. The bellcrank is attached at the forward end to the inside upper surface structure of the stabilizer and hinged to the elevator front spar. Pushrods inside the elevator connect the crank and tab. Control The elevator tab functions as a balance tab at all times. Elevator movement causes the pushrods to drive the tab in the opposite direction by an amount in proportion to elevator travel. Material Tab control rods are made either from aluminum or titanium. However, each pair of tab control rod has to consist of the same material.


B737-300/-400/-500



B737-300/-400/-500 27-30

SEE


1

Figu Figure re 63

Elev Elevat ator or Tab Cont Contro roll

1




B737-300/-400/-500 27-30

ELEVATOR BALANCE PANELS The Balance Panels (3) - are mounted mounted between between the structure structure of the Elevato Elevators rs and the structure structure of the stabilizer. - are used during during manual manual reversion reversion - are accessible accessible through through a panel. panel. - carry the the Balance Balance Weights Weights for for the tab balanc balancing. ing.



B737-300/-400/-500 27-30 IDLER HINGE SEAL

ELEVATOR ELEVATOR NOS E

IDLER HINGE

STABILIZER TRAILING EDGE RIB

SEE A BULB SEAL

STABILIZER SUPPORT BEAM

BALANCE PANEL FLAT SEAL ELEVATOR ELEVATOR (REF)

AFT MOUNTING BOLTS HINGE SEAL

FORWARD MOUNTING BOLTS

FWD

INBD

STABILIZER ACCESS PANEL (REF) STABILIZER REAR SPAR

ELEVATOR NOSE

A


VENT GAP

IDLER HINGE

HINGE SEAL

Figu Figure re 64

BALANCE HINGE BALANCE WEIGHT SEAL PANEL

Elev Elevat ator or Bala Balanc nce e Pan Panel el

ELEVATOR TAB ADJUST WEIGHTS (INSTALLED IN ELEVATOR NOSE IN BALANCE BAY NO. 2 AS REQUIRED



27-30 ELEVATOR FEEL COMPUTER Purpose The elevator feel computer supplies controlled system A and system B hydraulic pressure to the elevator feel control unit. Location The computer is mounted on structure in the compartment aft of the rear pressure bulkhead, to the right of the stabilizer jackscrew assembly. Physical Description/Feature The computer is a dual unit, the housing divided to accommodate identical components for system A and for system B. Internal components for each system include a q-diaphragm, force balance valve, relief valve and a stabilizer actuated cam. A feel differential pressure switch is installed in the housing.


B737-300/-400/-500



B737-300/-400/-500 27-30

FEEL COMPUTER

PITOT PROBE

STABILIZER INPUT

PITOT LINE

FEEL COMPUTER SEE

DRAIN PLUG

A

DRAIN LINE

PITOT LINE


FWD

A

Figur Figure e 65

Elev Elevat ator or Feel Feel Compu Compute terr Loc Locat atio ion n



27-30 ELEVATOR FEEL COMPUTER SCHEMATIC Input/Output Inputs to the feel computer are hydraulic system A and hydraulic system B pressures, pitot pressure, static pressure, and a stabilizer position mechanical input. Pitot pressures are directed to each system from individual pitot tubes on the vertical stabilizer. The outputs from the feel computer are two controlled hydraulic pressures to a dual feel actuator mounted on the feel and centering unit. The two output pressures are not additive. The highest of the two computed pressures will be reacted against. Feel forces increase as airspeed increases. Operation The two output pressures from the computer are controlled by pitot pressure and stabilizer position. Pitot pressure acting upon the q-diaphragm meters hydraulic pressure by displacing the force balance valve. As pitot pressure increases, the force balance valve will be actuated to increase pressure to the system until the droop spring and stabilizer actuated cam is contacted. Changes in stabilizer position rotates the stabilizer actuated cam and repositions the droop spring. A relief valve provides protection against excessive pressure in the feel system. Monitor The feel differential pressure switch monitors both computer output pressures. The switch closes when a difference of 25 per cent between system A and system B output pressure is detected. The elevator feel differential warning light on the forward overhead, P5, panel will illuminate when this difference exists and the trailing edge flaps are fully retracted.


B737-300/-400/-500



B737-300/-400/-500 27-30 NEUTRAL SHIFT MECHANISM

MACH TRIM ACTUATOR CENTERING LINKAGE

STABILIZER CENTERING CAM ELEVATOR FEEL AND CENTERING UNIT SYSTEM B RETURN

SYSTEM A RETURN

DUAL FEEL ACTUATOR

PITOT PORT

STATIC PORT Q DIAPHRAM DROOP SPRING

from Flight Control Module

STABILIZER ACTUATED CAM

PRESSURE PORT

SYSTEM B

RETURN PORT

RELIEF VALVE FORCE BALANCE VALVE

PRESSURE PORT RETURN PORT

CONTROL COLUMN


A 1

2

MACH TRIM SWITCH 2 MUST SENSE TRAILING EDGE FLAPS FULL UP FOR GROUND TO BE AVAILABLE

FEEL DIFFERENTIAL WARNING LIGHT - P5

2 2

TE FLAPS UP

PRESSURE DIFFERENTIAL SWITCH

MACH TRIM 1 SWITCH 2 MAIN WHEEL WELL

TE FLAPS NOT UP

CONTROL CABLES EB

MASTER DIM AND TEST

AFT CONTROL QUADRANT

STRUCTURAL PIVOT CONTROL CABLES EA

Figur Figure e 66

SYSTEM A

2

A

FORWARD CONTROL QUADRANT

FEEL COMPUTER

Elev Elevato atorr Fee Feell Com Comput puter er Sche Schema mati tic c



27-30 ELEVATOR ELEV ATOR FEEL AND CENTERING UNIT Purpose The elevator feel and centering unit provides artificial feel to the pilot and centering for the elevator control system. Location The feel and centering unit is mounted to structure between the input and output torque tubes in the empennage. Physical Description / Features The unit consists of a centering cam and roller assembly, centering linkage, and an externally mounted dual feel actuator. The feel actuator receives hydraulic inputs from the feel computer that act on pistons within separate chambers of the free-floating actuator case. A rod from one piston attaches to the feel and centering unit case. A rod from opposing piston attaches to the centering cam and roller assembly by means of a linkage.


B737-300/-400/-500

Operation Control column input rotates the cam against the spring-loaded roller. This causes a reaction through the centering linkage against the highest pressure at the feel actuator. The reaction force is felt at the control column. The feel and centering unit neutral position changes as stabilizer attitude varies. Changes in stabilizer attitude are transmitted through the neutral shift mechanism to rotate the feel and centering unit about the lateral axis. The cam follows the roller causing an input to the power control units and the control columns. A mach trim actuator is mounted on top of the feel and centering unit. The actuator rod is linked to the horizontal stabilizer through the neutral shift mechanism and functions as part of the neutral shift mechanism except when the mach trim system is operating. Mach trim command signals position the actuator output rod. This rotates the case of the feel and centering unit and inputs to the elevator system.



B737-300/-400/-500 27-30 AUTOPILOT B ACTUATOR OUTPUT TORQUE TUBE NEUTRAL SHIFT SENSOR

MACH TRIM ACTUATOR

AUTOPILOT A ACTUATOR ELEVATOR A PCU

FEEL AND CENTERING UNIT ELEVATOR B PCU

STABILIZER POSITION SENSOR


INPUT TORQUE TUBE AFT CONTROL QUADRANTS

FWD RIG PIN E-5

Figur Figure e 67

Elev Elevat ator or Feel Feel and and Cen Cente teri ring ng Unit Unit



27-30 ELEVATOR ELEV ATOR HYDRAULIC SYSTEM Normal Operation The elevators are normally operated by system A and system B hydraulic power. Hydraulic pressure to the power control units, feel computer, and autopilot actuators is applied through the flight control shutoff valves in tee main gear wheel well. These valves are electrically operated by their respective switches on the overhead panel. The power control units position the elevators. Elevator feel forces are provided by the dual feel actuator mounted on the feel and centering unit. The actuator is provided pressure controlled by airspeed, modified by stabilizer position. One hydraulic system is capable of sustaining normal elevator control. Backup Operation A manual reversion mode allows full elevator control by mechanical means with both hydraulic systems off.


B737-300/-400/-500



B737-300/-400/-500 27-30 ELEVATOR DUAL FEEL ACTUATOR

DIFFERENTIAL PRESSURE SWITCH

FEEL COMPUTER TO RUDDER POWER UNIT

TO RUDDER POWER UNIT

P

P

R

R

ELEVATOR POWER CONTROL UNITS

”A” SYSTEM PRESSURE

”B” SYSTEM PRESSURE

TO INBOARD SPOILERS

TO OUTBOARD SPOILERS M


M

PS

TO AILERON POWER UNIT ”A”

1

TO AILERON POWER UNIT ”B”

PS

M

M

HYDRAULIC ISOLATION VALVES

TO ”A” SYSTEM RETURN

FROM AILERON POWER UNIT ”A”

Figu Figure re 68

FROM AILERON POWER UNIT ”B”

Elev Elevat ator or Hydra Hydraul ulic ic Sche Schema mati tic c

TO ”B” SYSTEM RETURN



27-30 HYDRAULIC ISOLATION VALVES VALVES Purpose Manually operated isolation valves provide a means of isolating the elevator power control units, the elevator feel computer, and the rudder power unit for ground leakage and flow tests. Location A modular assembly consisting of six hydraulic isolation valves is mounted on the right side of the fuselage, inside the stabilizer compartment, aft of the rear pressure bulkhead. Physical Description An isolation valve is installed in each of the system A and system B hydraulic lines to the feel computer, the elevator power unit and tab actuator, and the rudder power unit. A placard adjacent to each valve identifies the system effected. A cover plate which is bolted in place maintains the valves in the open position. Operation The cover plate must be removed to close one or more of the valves. The cover plate cannot be installed unless all the valves are open. The valves must be open and cover plate installed for flight.


B737-300/-400/-500



B737-300/-400/-500 27-30

ISOLATION VALVES MODULE LOCKING BAR

N0. 6 ELEVATOR RIGHT PCU and A/P

TAIL COMPARTMENT SYSTEM B

RUDDER

N0. 5

and YAW DAMPER FEEL COMPUTER N0. 4

N0. 3 ELEVATOR LEFT PCU and A/P SYSTEM A


N0. 2

RUDDER

FEEL COMPUTER N0. 1 VALVES (POSITIONED OPEN)

FWD

Figur Figure e 69

Hydr Hydrau auli lic c Isol Isolat ation ion Valve alves s



27-32

27-32

STALL WARNING SYSTEM

STALL WARNING INTRODUCTION GENERAL Two stall warning systems alert the pilots of an approaching stall condition. The warning is accomplished by vibrating the pilots’ control columns. This warning is also used if airspeed falls too low. Each stall warning system operates independently and includes an angle of airflow sensor, a flap position transmitter, a digital stall warning computer, a self-test switch, and a control column shaker. The angle of airflow (AOA) sensor detects the airplane angle of attack and the flap position transmitter senses the position of the trailing edge flaps. These signals are passed to the stall warning computer along with information about airplane flight status. At predetermined combinations of flap position and airplane angle of attack, the computer outputs a stall warning signal to activate the control column shaker. The point at which the warning occurs is also influenced by:  engine speed,  airspeed, and any  asymmetry between leading edge position and the flap selection. The stall warning computers provide information to the ground proximity warning computer for wind shear warning, to the Electronic Flight Instrument System (EFIS) for pitch limit and speed tape display functions, and to the digital flight data acquisition unit (DFDA MRU).


B737-300/-400/-500

Two different types of Stall Warning Computers are listed in the AMM:  CONFIG 1 : DSWC (Digital Stall Warning Computer), operating in the older types of aircrafts,  CONFIG 2 : SMC (Stall Management Computer), operating in the newer types of aircrafts. Both types are identical in their stall warning calculation, but the CONFIG 2 additional contains the Autoslat-Function, so there is no need for a separate Autoslat Computer. The SMC is able to operate as a DSWC but not vice versa.



B737-300/-400/-500 27-32

Stall Warning Test Module

AOA Sensor left

N1

EADI

N2


Stall Warning Comp. #1 Flap Position XMTR right

Figu Figure re 70

Stal Stalll War Warni ning ng Intro Introdu duct ctio ion n

Stall Management Computer #1



B737-300/-400/-500 27-32





B737-300/-400/-500 27-32

D


F

Figur Figure e 71

Stal Stalll Warni arning ng Com Compon ponen entt Loca Locatio tion n



27-32 GENERAL DESCRIPTION Inputs The following signals are linked to the stall warning computers:  Angle of attack from the on-side AOA Sensor,  Flap position from the off-side Flap Position Transmitter,  N1 revolution from the on-side N1 Indicator,  N2 revolution from the off-side N2 Indicator.  Air/GND Relays,  Takeoff/Go Around from thrust levers  Asymmetry signal from Leading Edge Indication Module,  ADC,  IRS and  FMC

Outputs Several different airplane flight conditions can cause the stall warning computer to activate the control column shaker. shaker. The shaker will be activated by whichever condition gives the earliest warning of an approaching stall. The stall warning computers also provide information to the  GPWS, ground proximity warning computer for wind shear warning,  Electronic Flight Instrument System (EFIS) for pitch limit and speed tape display functions,  and to the digital flight data acquisition unit (DFDA MRU). y l n O s e s o p r u P g n i n i a r T r o F

B737-300/-400/-500

Test An Operational Test may be started from the overhead panel separately for each System.



B737-300/-400/-500 27-32

STAL L WARNING / COMPUTER

STALL MANAGEMENT COMPUTER

left AOA

STAL L WARNING / STALL MANAGEMENT

left N1 Ind.

COMPUTER

#1 right N2 Ind.


right Flap Pos. XMTR STALL WARNING ASYM MODE

LE FLAPS LE FLAPS TRANSIT EXT A A

Figur Figure e 72

to System #2 from System #2

Stall Stall War Warni ning ng Sys Syste tem m Bloc Block k Diag Diagra ram m



27-32 OUTPUT SIGNALS In addition to activating the control column shaker, the stall warning computer provides information to other airplane flight systems. Ground Proximity Warning The stall warning computer transmits angle of attack and flap position data to the ground proximity warning computer (GPWC) to assist the GPWC in computing a wind shear warning. Pitch Limit An airplane pitch limit is calculated by the stall warning computer from data received from the IRU, the DADC, and the AOA sensor. This pitch limit information is transmitted to the EFIS where it is displayed on the pilots’ electronic attitude director indicators. Speed Tape Information for the speed tape display of the electronic attitude director indicator (EADI) is computed in the stall warning computer from data received from the IRU, the DADC, the FMC, the flap position transmitter, and the AOA sensor. The control column shaker activation speed speed and minimum operating airspeed are included in the speed tape display. The speed tape display is activated when connectors D8489P and D8489J on the E2-3 electronics shelf are connected. Digital Flight Data Acquisition Unit Information from the stall warning computer is transmitted to the digital flight data acquisition unit so that it may be recorded.


B737-300/-400/-500




B737-300/-400/-500 27-32

Figu Figurre 73

Outp utput Si Signal nals



27-32 ADJUSTMENT /TEST Additional to the normal ‘‘Operational Test‘‘ (activating from the cockpit) and the ‘‘BITE Test‘‘ Test‘‘ (activating at the computer), the AMM contains the following sequences: CONFIG 1  System Performance Test  Landing Gear Relay Test  Asymmetry Mode Test  High Thrust Mode Test  Speed Floor Test  Speed Tape and Pitch Limit Test. CONFIG 2  SMC Discrete Input Test  SMC ARINC Input Test  SMC 1 AOA Sensor Input Test  SMC 2 AOA Sensor Input Test  SMC Flap Position Input Test  SMC Engine Tachometer Inputs Test  SMC to EFIS Interface Test  SMC Asymmetry Mode Test.

OPERATIONAL TEST (CONFIDENCE TEST) y l n O s e s o p r u P g n i n i a r T r o F

B737-300/-400/-500

Test in short form:  Push and hold the STALL WARNING TEST No.1 (2) switch on the stall warning test module (P5 overhead panel).  Make sure that captain‘s (first officer‘s) control column shaker operates.  Release the switch.  Make sure that the shaker stops. ADVICE: Some important checks for the AOA-Sensor will be found in chapter 27-32-11 /601 (Inspection /Check) of the AMM .



B737-300/-400/-500 27-32

SENSORS

CONFIG 1

STALL WARNING AC

STALL WARNING DC

STALL WARNING ASYM MODE

LEADING EDGE MODULE

DSWC

CONFIG 2 SENSORS SMC SNSR EXC AC


(open) SMC CMPTRDC

STALL WARNING ASYM MODE

LEADING EDGE MODULE AUTO SLAT DC

SMC

Figu Figure re 74

Oper Operat atio iona nall Tes Testt



27-32 BITE TEST OF DSWC (CONFIG 1) Preparation for the Test  Open CB’s for - AUTOSL AUTOSLAT AT System System,, - STALL STALL WARN WARN ASYM ASYM MODE MODE  Put a value for gross weight in the FMC. BITE Test  Erase the stall warning computer fault memory: - Push and hold hold the START START BITE switch on the front of stall warning warning computer. - Make sure sure that all lights lights on the computer computer come come on for 2 to 4 seconds. seconds. - Look to see see if one or more more lights lights stay on on after 4 seconds seconds.. NOTE:


B737-300/-400/-500

IF A LIGHT STAYS STAYS ON AFTER 4 SECONDS, THE COMPUTER MEMORY HOLDS A FAILURE FOR THE FUNCTION. THE LIGHT WILL GO OFF AFTER APPROXIMATELY 25 SECONDS. - When one or more more lights lights starts to flash flash after 10 to 20 seconds, seconds, release release the START BITE switch. The memory is now clear.  Do a BITE test for stall warning computer: - Push and hold hold the START START BITE switch on the front of stall warning warning computer. - Make sure sure that all all lights lights on the compute computerr come on for 2 to 4 seconds. seconds. - Release Release the the START START BITE BITE switch. switch. - Look to see see if one or more more lights lights stays on on after 4 seconds seconds.. NOTE:

IF A LIGHT STAYS STAYS ON AFTER 4 SECONDS, THE COMPUTER MEMORY HOLDS A REAL-TIME FAILURE FOR THE FUNCTION.




B737-300/-400/-500 27-32

DSWC

Figu Figure re 75

BITE BITE Test est (co (conf nfig ig.. 1)




B737-300/-400/-500 27-32

BITE TEST OF SMC (CONFIG 2) General You will use the buttons on the front of the stall management computer (SMC) to do a BITE test and to monitor the inputs to the SMC’s. The SMC will allow entry into BITE only when one or more of these conditions exist: - Engines Engines OFF OFF (N1 < 15%, 15%, N2 < 50%). 50%). - Airspeed Airspeed < 60 knots knots AND AND flaps fully retracted. retracted. BITE Control and Display  The maintenance control panel on the front of the SMC has two 8-character displays and 6 buttons for fault isolation and for checks during ground maintenance. - The ON/OFF ON/OFF button activate activates s and deactivates deactivates the BITE feature feature.. All buttons on the SMC are normally inactive until the ON/OFF button is pushed to activate BITE. - The MENU button button displays displays the the current current menu at any point point in BITE. - The up and down down button button errors are are used to sequence sequence through through the items items under each menu. - If a button is held down, down, approximatel approximately y 4 items/options items/options per second second will be shown in sequence. - The up-sequen up-sequence ce ends by releasin releasing g the button button or at the end of the menu menu with display showing BEGIN OF LIST or END OF LIST. - The YES and and NO buttons buttons are are pushed pushed to answer answer a menu option. option. - If the NO button button is held down, down, options options will be shown shown in sequence sequence at approximately 4 options per second.  To enter BITE, push the ON/OFF button. A BITE entry test will automatically be performed by the SMC. - If the entry test test fails, the display display will show TEST FAIL momentarily, momentarily, followed by nn FAULTS, FAULTS, where nn is the number of faults detected during the entry test. - The display display will show show the first fault. fault. Other Other faults (if (if they exist) exist) will be shown as the down button is pushed repeatedly.

- If the entry passes, passes, the the display will will show SYS OK. OK. Push the menu menu button to access the options of the main menu. These are the main menu options: A. PRESENT PRESENT FAUL FAULTS TS ? - PRESENT FAULTS FAULTS provides the continuous updated updated (real time) status of interface signals. B. SELF SELF TEST TEST ? - The SELF SELF TEST option option run run a test test of the SMC SMC LRU iitself. tself. During the test, the display shows TEST IN PROGRESS, followed by either SMC LRU OK or SMC LRU FAIL. FAIL. For a failed test , after SMC LRU FAIL shows momentarily, nn FAULT shows, then DISP FAULT ? shows. If you answer YES to DISP FAULTS FAULTS ? the display shows a list of faults internal to the SMC. C. FAULT FAULT HISTORY HISTORY ? - FAULT FAULT HISTORY shows shows any faults detected by SMC during the previous previous flight leg. D. GROUND GROUND TEST TEST ? - GROUND GROUND TEST allows allows testing and trouble trouble shooting shooting of system inputs inputs on the ground. System inputs are listed under one of three menu options :  DISCRETE INPUTS ?  ARINC TEST ?  SENSOR TEST ? E. SYSTEM SYSTEM CONFI CONFIG G?  The SYSTEM CONFIG option gives information on the airplane model and airplane configuration for which the SMC is currently programmed. F. CLR AND RETEST RETEST ? - The CLEAR CLEAR AND RETEST RETEST option allows allows the operator operator to clear clear any twoflight transition faults in the SMC RAM.



B737-300/-400/-500 27-32

BUTTON ON SMC MAINTENANCE CONTROL PANEL

ON/ OFF

DISPLAY ON SMC MAINTENANCE CONTROL PANEL

BITE INSTRUCTIONS

TEST IN PROGRESS TEST TEST PASS PASS TEST FAIL FAIL SYSTEM OK

TEST FAIL

SMC MAINTENANCE CONTROL PANEL TO TEST FAIL

MENU

PRESENT FAULTS?

YES

TO PRESENT FAULTS MENU

YES

TO SELF TEST MENU

NO SELF TEST?

MENU

ON/ OFF

YES

NO

NO FAULT HISTORY?

YES

TO FAULT HISTORY MENU

YES

TO GROUND TEST MENU

NO GROUND TEST?


STALL MANAGEMENT COMPUTER

NO

PART NO. 65-52822 SYSTEM CONFIG?

YES

TO SYSTEM CONFIG MENU SERIAL NO. BOEING

NO CLR AND RETEST?

NO

YES

TO CLR AND RETEST MENU

____ NOTE: TO EXIT FROM ANY POINT IN SMC BITE, PRESS THE ON/OFF BUTTON.

Figu Figure re 76

BITE BITE Test est (co (conf nfig ig.. 2)


FLIGHT CONTROLS HORIZONTAL STABILIZER

27-40

27-40


TRIM CONTROL SYSTEM INTRODUCTION Purpose The horizontal stabilizer trim control system provides longitudinal trim of the airplane by varying the angle of attack of the horizontal stabilizer. System Description The horizontal stabilizer assembly consists of a left and right section attached to a center section. The center section is connected to a jackscrew assembly that drives the stabilizer. The jackscrew is operated by either of two electric actuators (main electric or autopilot) or manually by cables. The cable system also operates trim position indicators, adjacent to the trim wheels on the control stand, to provide continuous indication of Stabilizer position. General Component Locations The stabilizer jackscrew, aft drum, and both electric actuators are located in a compartment forward of the stabilizer. A forward cable drum assembly and a column switching module are located in the lower nose compartment. Manual trim wheels, main electric trim switches, and position indicators for each pilot are located in the flight compartment.


B737-300/-400/-500

General Subsystem Features The maximum travel limit of the stabilizer is 17 units. The position indication scale is calibrated from 0 to 17 units with 3 units the neutral position. This provides a maximum of 3 units airplane nose down and 14 units airplane nose up trim. The stabilizer must be positioned within the green band on the position indicator for takeoff or the takeoff warning horn will sound. System interfaces A speed trim system is incorporated in the autopilot system. The autopilot commands the autopilot servo motor to trim the stabilizer, when required, to compensate for unstable flight conditions experienced during the low speed, high thrust conditions at takeoff. This system will be covered in autoflight. The stabilizer is connected to the elevator control linkage through neutral shift rods. Whenever the stabilizer is moved from 3 to 17 units the elevators are gradually moved up.



B737-300/-400/-500 27-40

JACKSCREW WITH BALLNUT

GEARBOX AND ACTUATORS



POSITION INDICATOR AND MANUAL CONTROL MECHANISM

Figu Figure re 77

Stabi Stabili lize zerr Trim Trim Contr Control ol Syst System em Loca Locati tion on



27-40 STABILIZER TRIM CONTROL SCHEMATIC General Operation The horizontal stabilizer is operated three ways and in this order of priority: - Manual Manual - trim trim wheels wheels on the the control control stand. stand. - Main Electric Electric - thumb thumb switches switches on both both control control wheels. wheels. - Autopilot Autopilot - pitch pitch channel channel or speed speed trim trim system. system. Normal stabilizer operation is electrical by either the autopilot, main electric actuator or on aircrafts with primary stabilizer trim actuator by the primary actuator driving the jackscrew through a gearbox. Manual control is accomplished by driving the aft cable drum on the jackscrew gearbox through cables. The manual system remains engaged at all times and is thus back-driven by the two electric actuators. This ensures correct indication on the position indicator. Manual system operation will disengage both electric actuators. Normal electric trimming of the stabilizer is done at one of two rates as controlled by flap position. Trim rate with flaps retracted is 1/3 the trim rate with flaps extended. The autopilot actuator also trims at one of two rates as controlled by flap position. High speed autopilot rate is equal to the normal electric low speed rate. The low speed autopilot rate is 1/2 the rate of the high speed autopilot rate. The column switching module prevents the electric actuators from moving the stabilizer in a direction opposite to pilot control column movement if the column has been moved a predetermined amount. A column override switch is available to bypass the column switching module. Cutout switches can remove all electric power from either actuator. CAUTION:


B737-300/-400/-500

IN THE EVENT OF SIMULTANEOUS ACTUATION OF THE TRIM CONTROL SWITCHES FOR OPPOSITE DIRECTIONS OF TRIM, THE SWITCHING WILL CAUSE BOTH ELECTROMAGNETIC CLUTCHES TO ENGAGE AND RESULT IN MOTOR STALL WHICH MAY DAMAGE THE MOTOR DUE TO OVERHEATING.




B737-300/-400/-500 27-40

Figur Figure e 78

Stabi Stabili lize zerr Syst System em Sche Schema matic tic



27-40 STABILIZER FORWARD FORWARD CONTROL MECHA NISM

COLUMN SWITCHING MODULE

Purpose The stabilizer forward control mechanism provides the means of manual control for the stabilizer in the event of electrical malfunction.

Purpose The column switching module provides the capability to stop electric operation of the stabilizer when the control column is moved in the direction opposing stabilizer movement.

Location The forward control mechanism extends from a stabilizer trim wheel on each side of the control stand to a forward trim mechanism in the lower nose compartment. Physical Description/Feature The stabilizer trim wheels and a sprocket are splined to a control wheel shaft that extends through the control stand. Rotation of the stabilizer trim wheels transmits motion to the forward trim mechanism sprocket by a chain assembly. The forward trim mechanism sprocket drives the forward cable drum which is connected by cables to an aft cable drum on the stabilizer jackscrew and gearbox assembly. Operation Manual control is accomplished by rotating either trim wheel on the control stand. Operation of the forward control mechanism drives the jackscrew and gearbox assembly via the cables to position the stabilizer. During electrical operation of the stabilizer the jackscrew and gearbox assembly drives the forward control mechanism, through the cables, to provide stabilizer position indication and rotate the trim wheels.


B737-300/-400/-500

Monitor A stabilizer position indicator provides continuous indication of stabilizer trim. Operation of the forward control mechanism drives a jackshaft through a flexible cable and transmits motion to a linkage that positions the indicator. A scale on the control stand is calibrated in units of trim and has an area outlined to indicate the proper takeoff stabilizer range, referred to as the GREEN RANGE. Maintenance Practices The forward cable drum assembly is mounted in a housing which is attached to structure by one horizontal and two vertical suspension points. Turnbuckles at these suspension points can be adjusted to align the forward mechanism and obtain proper chain and cable tension.

Location The module is located in the lower nose compartment on the right side beneath the first officer’s position. Physical Description / Feature The column switching module contains relays and switches that are in the stabilizer trim electric circuits between both the main electric actuator and the autopilot actuator and their respective control components. A mechanical linkage to operate the switches is connected to the torque tube that interconnects both elevator quadrants.



B737-300/-400/-500 27-40 STABILIZER TRIM CONTROL SWITCH

COLUMN ACTUATED CUTOUT SWITCH OVERRIDE (CONTROL STAND)

STABILIZER TRIM

CONTROL COLUMN

STAB TRIM

STABILIZER TRIM CONTROL SWITCH

STABILIZER TRIM WHEEL

CAB DOOR

APL NOSE DOWN

STABILIZER POSITION INDICATOR

OVERRIDE CAB DOOR UNLOCKED A

0 3 0

NORM 5 (COLUMN SWITCHING MODULE)

ELEVATOR TORQUE TUBE

2 0 1 0

10 COLUMN ACTUATED CUTOUT SWITCH 15 APL NOSE UP INDICATOR FLEXIBLE SHAFT AFT SUPPORT LINK STB CABLE 1

CHAIN ASSEMBLY


STA CABLE

FORWARD CABLE DRUM

FORWARD SUPPORT LINK 1

APPROXIMATLEY 4-1/2 WRAPS OF STB CABLE FROM TOP OF DRUM AND 28-1/2 WRAPS OF STA CABLE FROM BOTTOM OF DRUM WITH STABILIZER POSITIONED AT FULL AIRPLANE NOSEDOWN

FORWARD TRIM MECHANISM SPROCKET

TURNBUCKLE

Figu Figure re 79

FORWARD TRIM MECHANISM GEAR HOUSING

Stab Stabil iliz izer er Forw Forwar ard d Contr Control ol Mech Mechan anis ism m

1

T A K C E G O - F % F M A C



27-40 STABILIZER JACKSCREW AND GEARBOX ASSEMBLY Purpose The stabilizer jackscrew and gearbox assembly positions the stabilizer by converting electric actuator, autopilot actuator, or cable drum rotary motion to linear motion. Location This assembly is located in a compartment, aft of the rear pressure bulkhead, that is accessible through a door on the lower left side of the fuselage. Physical Description/Feature The stabilizer jackscrew and gearbox assembly consists of a ball nut and jackscrew, a gimbal assembly, a cable drum, and a gearbox consisting of gearing and brakes. The gearbox is connected to a bulkhead in the fuselage by a lower gimbal which allows fore and aft angular motion as the stabilizer is positioned. An upper gimbal connects the ball nut to the stabilizer front spar fitting. A safety rod is installed in the jackscrew shaft to support the stabilizer in the event of jackscrew failure. Power Two electric actuator are mounted on the assembly and drive the gearbox through output shafts. Both are two speed, three phase, 115 volt ac motors. Normal operation is the main electric actuator controlled by the pilot’s control wheel switches. Autopilot control of the stabilizer is through the autopilot actuator.


B737-300/-400/-500



B737-300/-400/-500 27-40 SAFETY ROD

JACKSCREW BALLNUT STABILIZER STRUCTURE GREASE FITTING

UPPER GIMBAL GEARBOX

ELECTRIC ACTUATOR (NORMAL)

LOWER STOP (UPPER STOP NOT SHOWN BUT EQUIVALENT) UMBRELLA ASSEMBLY COVER PLATE

AUTOPILOT ACTUATOR LOWER GIMBAL


STB CABLE

STA CABLE FWD GEARBOX CABLE DRUM

Figure Figure 80

Stabil Stabilize izerr Jacksc Jackscrew rew and Gear Gearbox box Asse Assembl mbly y (confi (config. g. 1) 1)



B737-300/-400/-500 27-40





B737-300/-400/-500 27-40 SAFETY ROD

JACKSCREW BALL NUT

STABILIZER STRUCTURE UPPER GIMBAL

PRIMARY STABILIZER TRIM ACTUATOR

GREASE FITTING LOWER STOP (UPPER STOP NOT SHOWN BUT EQUIVALENT)

GEARBOX

UMBRELLA ASSEMBLY COVER PLATE

LOWER GIMBAL ACTUATOR COVER


STB CABLE

FWD

STA CABLE

Figure Figure 81

Stabil Stabilize izerr Jacksc Jackscrew rew and Gear Gearbox box Asse Assembl mbly y (confi (config. g. 2) 2)

GEARBOX CABLE DRUM



27-40 STABILIZER JACKSCREW AND GEARBOX DIAGRAM Control The primary brake system in the stabilizer trim gearbox prevents any aerodynamic loads on the stabilizer from rotating the jackscrew when the control system is not being operated. When the jackscrew shaft is driven by the jackscrew gearbox in a direction which increases the airloads on the stabilizer, a brake plate rotates through a ratchet producing a clicking sound in the assembly. The auxiliary brake system is provided in case of primary brake system failure. When the jackscrew is turned by the jackscrew gearbox, the auxiliary brake system is released through a gear driven by the gearbox. If the primary brake system fails, the jackscrew shaft rotation applies the auxiliary brake to prevent any additional shaft rotation. Operation When the main electric actuator is energized, a uni-directional motor drives the trim input servo gear in the direction of the engaged magnetic clutch. The clutch drive gear drives the drive gear shaft through the engagement of the upper clutch member and the spring-loaded lower clutch member. The drive gear powers the jackscrew through the brake unlock gear. Because the lower clutch member is splined to the cable drum, the drum rotates to operate the position indicator and trim wheels. The autopilot motor operates in the same manner described for the main electric actuator through the autopilot trim servo input gear. Manual input from the trim wheels causes the lower clutch member to cam down against the spring and disconnects the shaft from the clutch drive gear. This effectively disconnects both electric actuators.


B737-300/-400/-500



B737-300/-400/-500 27-40 ELECTRO MAGNETIC CLUTCH

JACKSCREW AUXILIARY BRAKE ASSEMBLY

SAFETY ROD

MOTOR

SEE

A

BRAKE UNLOCK GEAR RATCHET PAWL

DRIVE GEAR CLUTCH DRIVE GEAR

REDUCTION GEAR

PRIMARY BRAKES TORQUE LIMITER CLUTCH


SHOE BRAKE ARM

UPPER CLUTCH MEMBER

TRIM SERVO INPUT GEAR

LOWER CLUTCH MEMBER

BRAKE UNLOCK GEAR

RELEASE CAM AND ROLLER ASSEMBLY (4 PLACES)

JACKSCREW A

CLUTCH SPRING BRAKE DRUM

CABLE DRUM

Figu Figure re 82

Stabil Stabilize izerr Jack Jacksc scre rew w and and Gea Gearbo rbox x

AUTOPILOT TRIM SERVO INPUT GEAR



27-40 STABILIZER TRIM LIMIT SWITCHES Purpose Four stabilizer trim limit switches limit the up and down travel of the horizontal stabilizer leading edge during either main electric or autopilot operation. Location The four cam operated microswitches are mounted in a vertical row on brackets attached to structure at the left of the jackscrew attach fitting. Two stabilizer takeoff warning switches are included in this row to make a total of six switches. Features All switches are operated by the same cam which is mounted by a support tube to the horizontal stabilizer center section jackscrew attach fitting. Operation The cam moves with the stabilizer to actuate the limit switches to the open position at the desired stabilizer travel limits. Opening of a switch removes power to the respective actuator to terminate stabilizer travel. Main electric nose down travel is limited by one of two limit switches as selected by flap position. Takeoff warning switch operation will be discussed under takeoff warning.


B737-300/-400/-500



B737-300/-400/-500 27-40 STAB TAKEOFF WARNING SWITCH (S546) (REF 31-26-24)

STAB AUTOPILOT AND ELEC ACTUATOR LIMIT SWITCH (S145)

STAB ELEC ACTUATOR LIMIT SWITCH (S844)

SWITCH MOUNTING BRACKETS

LIMIT SWITCHES (AIRPLANE NOSE DOWN)

CAM LIMIT SWITCHES (AIRPLANE NOSE UP)


STAB TAKEOFF WARNING SWITCH (S132) (REF 31-26-24)

FWD

Figu Figure re 83

STAB ELEC ACTUATOR LIMIT SWITCH (S115)

Stabil Stabilize izerr Tri Trim m Swit Switch ches es Loc Locat atio ion n

STAB AUTOPILOT LIMIT SWITCH (S144)




B737-300/-400/-500 27-40

MAIN ELECTRIC ACTUATOR OPERATION OPERATION Operation Normal control and operation of stabilizer trim is accomplished with electrical power. Control power for the relays, the electromagnetic clutches is 28-volts dc. The actuating power for the stabilizer trim motor motor is 115-volt, 115-volt, three-phase ac. When the three-phase power source has the correct phase sequence, assuring correct direction of motor rotation, the phase sequence relay contacts will be closed. The clutch and relay control circuits are ready for normal electrical operation only when the main electric trim cutout switch is at the NORMAL position. With power supplied and the trim cutout switch at the NORMAL position, the stabilizer safety relay coil will be supplied 28-volt dc power to hold the contacts closed. This will allow the trim control relay to control switching power directly to the motor for normal operation. The actuation of either trim control switch will switch control power to energize the trim control relay and the appropriate electromagnetic clutch. This action connects three-phase power to the motor, which rotates in one direction only, and engages the clutch which controls the direction of actuator drive. With the electrical trim motor actuating the jackscrew, the trim wheels will be turned and the trim indicators positioned through the manual system as the manual system is driven from the cable drum on the jackscrew gearbox. The trim actuator will continue to drive the jackscrew in the direction selected until the trim control switch is released, the limit switch is actuated, or the cutout switch is positioned to CUTOUT. CUTOUT. Limit switch actuation interrupts power to the clutch which disengages the motor. The motor will not be de-energized by limit switch actuation. The trim wheel rotation will stop and trim position indicator will show that the trim action has stopped at the extreme end of travel. Release of trim control switch will cause the control relay to drop out, interrupting motor power. Switching trim cutout switch to the CUTOUT position disconnects all normal control circuits, shorts all relay and clutch solenoids to ground, and de-energizes the safety relay to open the trim motor power circuit.

Normal Sequence When either pilot actuates his control wheel switches, control power is applied to the circuit through the column switching module and cutout switch. The selected directional clutch is energized through the respective limit switch. The stab trim control relay is energized through the phase sequence relay. The motor is driven by ac power to move the stabilizer in the selected direction at the speed determined by flap position. The limit switch opens and stops stabilizer movement when the preset travel limit is reached. CAUTION:

CAUTION: DO NOT EXCEED MAIN ELECTRIC ACTUATOR DUTY CYCLE OF 2 MINUTES ON AND 13 MINUTES OFF. OFF.

Backup Operation The stabilizer is operated manually by the pilots’ trim wheels when electric operation is not possible. The cutout switch is actuated to CUTOUT to remove electric power from the circuit and prevent motor operation, when required. The switches inside the module provide a path for electric power between the two actuators and their control components when the control column is in the neutral range. When the control column is moved out of the neutral range, switches in the module open the electrical circuit to both actuators in the direction opposite of column movement. Relays and another set of switches allow the actuators to operate the stabilizer in the same direction as column movement. A switch on the control stand, when positioned to OVERRIDE, provides a path for electric power to bypass the column switching module. The stabilizer can then be electrically operated in both directions regardless of column position.



B737-300/-400/-500 27-40

LOW SPEED

115V AC

HI SPEED STAB TRIM ACTUATOR

STAB TRIM SAFETY RELAY R66

STAB TRIM CONTROL RELAY R64

PHASE SEQ RELAY R63

SPEED CHANGE RELAY

28V DC STAB TRIM CONTROL

STABILIZER TRIM SHIELD

NOSE UP

NOSE DN STAB TRIM SW S135

NOT UP STAB ELEC ACT LIMIT APL NOSE UP S115

NOSE UP CLUTCH

NORMAL COLUMN OVERRIDE SWITCH S847 NOT DOWN STAB ELEC ACT LIMIT APL NOSE DOWN FLAPS UP S844

RIGHT WHEEL COLUMN COLUMN AFT NOSE UP

NOSE DN CLUTCH

NOT DOWN STAB ELEC ACT APL NOSE UP S145 CUTOUT


NOSE DN STAB TRIM SW S134 LEFT WHEEL COLUMN

STAB TRIM CUTOUT SWITCH S272 COLUMN FWD

STAB TRIM ACTUATOR M222

FLAPS UP RELAY R335

CONTROL STAND

COLUMN ACTUATED CUTOUT SWITCH M1201

FLAPS NOT UP S245 1

1

AIRPLANES WITH FLAPS UP RELAY

2

ALL EXCEPT

FLAPS NOT UP S245

1

2

Figur Figure e 84

Stabi Stabili lize zerr Tri Trim m Cont Contro roll Cir Circu cuit it



FLIGHT CONTROLS FLAPS

B737-300/-400/-500 27-50

27-50

FLAPS

TRAILING EDGE FLAPS INTRODUCTION Purpose The trailing edge flap system is installed to provide additional lift during takeoff and landing by increasing the camber of the wing.

System Interfaces Leading edge flaps and slats on each wing operate in conjunction with the trailing edge flaps to provide increased lift.

System Description The trailing edge flap system consists of an inboard and an outboard assembly on each wing that are each composed of three mechanically linked segments which extend and separate to form a triple for added lift. A hydraulic motor drives all trailing-edge flaps by means of a torque-tube drive system connected to two ball bearing drive screws on each flap assembly. This motor has an automatic shutoff feature in case of either flap asymmetry or cable tension loss. An electric motor serves as a backup for trailing edge flap extension and retraction. The normal cruise position of the flaps is retracted with the trailing edge flaps nested together to form a continuous surface. Flaps are extended for takeoff and landing to increase the effective wing area. Takeoff Takeoff flap positions provide high lift and relatively low drag. Landing flaps produce high lift and high drag which aids in deceleration to low approach speeds. A flap load limiter system protects the trailing edge flaps from excessive airloads by automatically retracting flaps from the fully extended landing position when a predetermined airspeed is exceeded. When airspeed is reduced the flaps automatically return to the fully extended position.

General Operation Normal operation of the trailing edge flaps is system B hydraulic power control led by the flap lever. Alternate operation is by an electric motor controlled by two switches on the overhead pane!.

General Component Locations The flap control lever is on the pilots’ control stand. The flap control unit and flap power unit are in the main gear wheel well. A position transmitter is mounted on each wing rear spar and the indicator is in the flight compartment. General Subsystem Features Each flap assembly travels on two flap tracks. The torque tube drive system operates two transmission assemblies attached to the mid flap.



B737-300/-400/-500 27-50

FLAP CONTROL LEVER

WFB CABLE RIGHT ANGLE GEARBOX

FLAP CONTROL UNIT

WFA CABLE

ANGLE GEARBOX

FLAP POSITION TRANSMITTER INBOARD


TRAILING EDGE FLAP OUTBOARD FLAP TRACK FAIRINGS OUTBOARD TRAILING EDGE FLAP FLAP POWER UNIT

Figu Figure re 85

Trail railin ing g Edge Edge Fla Flap p Syst System em



27-50 FLAP LEVER Purpose The flap control lever provides the pilots a means of regulating the position of the trailing edge flaps using hydraulic power. Location The flap control lever is located on the upper right side of the control stand. Physical Description/Features The lever assembly consists of a spring loaded telescoping handle that rotates a cable drum around a shaft in the control stand. The lever rotates around a quadrant which has detents at the flap positions, graduated in units. The lever is spring-loaded to lock in each detent. Operation Lifting the lever releases the lock and allows rotation. The quadrant contains gates at the 1 and 15 unit detents which prevent inadvertent lever movement past these detent positions.  Position 1 to check flap position for one engine inoperative go-around  Position 15 - to check flap position for normal go - around. The lever must be lowered into the detent and passed under the gate before further rotation can occur. The flap control lever rotates the cable drum to actuate cables which position a flap control quadrant above the right wheel well.


B737-300/-400/-500



B737-300/-400/-500 27-50

FLAP CONTROL LEVER FLAP QUADRANT CONTROL STAND

FWD


Figure 86

Flap Le Lever



27-50 FLAP CONTROL UNIT Purpose The flap control unit is installed to contain the flap control valves and associated mechanical linkages that regulate hydraulic operation of the flaps. Location The flap control unit is mounted on the right wheel well ceiling near the aft bulkhead. Physical Description/Features The flap control unit incorporates the mechanical linkage that operates both the trailing edge flap control valve and the leading edge flap control valve. These valves are mounted on the control unit, as well as a cable operated follow-up drum that operates three cams. Seven cam operated electric switches are mounted on the forward side of the control unit. Control Normal hydraulic control of the trailing edge flaps is by cables from the flap control lever that operate the input linkage to the trailing edge flap control valve. A flap control cable tension switch is located below the cabin floor, near the flap control unit quadrant. Tension Tension on the flap control cables causes the spring loaded cable support lever to hold a target near a reed switch. A cable break allows the spring to pull the target away and activate the switch. This closes the flap bypass valve and prevents hydraulic operation of the flap power unit.


B737-300/-400/-500

Operation The trailing edge flaps are normally hydraulically operated. Rotation of the control lever actuates the flap control unit linkage through the cables. This positions the trailing edge control valve to port System B hydraulic pressure to the flap hydraulic motor. The hydraulic motor powers the torque tub’ drive system to position the flaps. Hydraulic motor operation positions cables that rotate the follow-up drum on the control unit. The follow-up drum positions three cams. One cam returns the trailing edge control valve to null and stops the flaps at the desired position. The second cam positions the leading edge control valve and the third actuates the respective electric switches.

Monitor Actual trailing edge flap position is monitored by seven cam operated switches on the flap control unit actuated by the follow-up system. These include alternate drive limit switches, a takeoff warning switch, a landing warning switch, leading edge indication switch, a mach trim switch, and a stall warning switch. Individual switch operation will be discussed when the specific circuit is covered.



B737-300/-400/-500 27-50

FLAP CONTROL QUADRANT

CABLE TENSION SWITCH

SEE A INPUT SHAFT FOLLOW-UP SHAFT FOLLOW-UP DRUM FLAP 10-DEGREE SWITCH S584

WFFA CABLE BELLCRANK

CAM ROLLER

WFFB CABLE TE FLAP CONTROL VALVE

FLAPS UP LIMIT SWITCH S245 STALL WARNING S856 FLAPS DOWN LIMIT SWITCH S246 TAKEOFF WARNING SWITCH S130


CAM ROLLER LANDING WARNING SWITCH S138

FLAP LOAD LIMITER SOLENOID VALVE

MACH TRIM SWITCH S290 TE FLAP CONTROL VALVE LINK

LE FLAP CONTROL VALVE SWITCH SENSED BAR (MAGNET)

SUPPORT

SPRING

A LEVER

SUPPORT

Figu Figure re 87

Flap Flap Cont Contro roll Unit Unit



27-50 FLAP POWER UNIT Purpose The flap power unit transfers mechanical energy from the flap hydraulic motor or alternate drive electric motor to the trailing edge flap drive system. Location The flap power unit is mounted in the center of the aft wall of the main gear wheel well. Physical Description/Features The flap power unit is an aluminum housing containing the gearing necessary to drive the output shaft and the follow-up system. The flap hydraulic motor and the alternate drive electric motor are mounted on the power unit. Operation During normal flap system operation, power from the hydraulic motor is transmitted through a pinion gear to a reduction gear splined to the output shaft. During flap alternate drive operation, power from the electric motor is transmitted through a second pinion gear to the same reduction gear. The reduction gear drives the flap torque tube drive through the output shaft. Both motors operate the same worm gear and worm wheel to drive the follow-up drum which is connected by cables to the flap control unit follow-up drum. CAUTION:


B737-300/-400/-500

DURING GROUND OPERATION, DO NOT OPERATE THE FLAP ALTERNATE ALTERNATE DRIVE UNIT MORE THAN 4 MINUTES OPERATION AND 25 MINUTES OFF. YOU CAN CAUSE DAMAGE TO THE FLAP ALTERNATE DRIVE UNIT.



B737-300/-400/-500 27-50

HYDRAULIC MOTOR

SEE

SHAFT

A

SHAFT RETAINER


FLAP POWER UNIT

FWD A

ALTERNATE ELECTRIC DRIVE UNIT

Figur gure 88

Flap Pow Powe er Un Unit



27-50 FLAP HYDRAULIC DRIVE SYSTEM Purpose

A reversible hydraulic motor attached to the flap power unit drives the flap system during normal operation. Location The hydraulic motor drive shaft mates with the power unit input shaft. It is on the aft wall of the main gear wheel well. Physical Description/Features The nine cylinder piston-type motor converts hydraulic pressure to mechanical energy. Two hydraulic lines from the flap control valve connect to inlet ports on the motor. A case drain line connecting the motor to the hydraulic fluid return system through a check valve allows lubrication of the motor. Power The hydraulic motor is powered by System B, available at the control valve, through a priority valve and a flow limiter. The flow limiter controls the speed of flap movement by regulating fluid flow to the motor at 9.0 to 10.5 gpm. The priority valve stops fluid flow to the flap hydraulic motor when B system demand causes a pressure drop below 2400 psi at the valve. This would most likely occur when the landing gear was retracted after takeoff with the left engine hydraulic pump inoperative. Control Hydraulic pressure to the motor is controlled by the position of the trailing edge flaps control valve and the flap bypass valve.


B737-300/-400/-500

Operation Movement of the flap control lever tensions the cables and drives the control unit input linkage to position the control valve. System B pressure is then applied through the normally open bypass valve to one port of the hydraulic motor. The pistons are actuated to rotate the output shaft and drive the power unit. Rotation of the follow-up drum by power unit operation drives the follow-up drum on the control unit to actuate the linkage and return the control valve to null.

Normal Sequence The trailing edge flaps are normally operated by hydraulic power. When the flap control lever is rotated, cables displace the flap control unit linkage. This positions the trailing edge flap control valve to port pressurized fluid from hydraulic system B to the flap hydraulic motor. The motor drives the flaps through a torque tube drive and transmission assemblies. As the flaps move, cables from a follow-up drum on the power unit rotate a follow-up mechanism on the control unit. A cam in the follow-up mechanism returns the control valve slide so that it nulls and stops the flow of hydraulic fluid when the desired flap position is reached. The operating speed of the trailing edge flaps is controlled by a flow limiting valve installed in the pressure line to the flap control valve. A priority valve stops fluid flow to the hydraulic motor when B system demand causes pressure to drop below 2400 psi at the valve. A motor operated bypass valve in the hydraulic lines between the control valve and the flap power unit controls operation of the hydraulic motor. The normal position of the bypass valve directs pressure thru one line to the motor and opens the other line to return. This allows the motor to run when the control valve is actuated.



B737-300/-400/-500 27-50

FOLLOW UP CAM FOLLOWER


Figur Figure e 89

Flap Flap Hydr Hydrau auli lic c Driv Drive e Sche Schema mati tic c



27-50 FLAP LOAD LIMITER Purpose The flap load limiter system protects the trailing edge flaps against excessive airloads Location A solenoid operated secondary valve slide is incorporated in the trailing edge flap control valve. The solenoid is mounted on the valve. Physical Description/Features The load limiter system consists of the solenoid valve, two flap lever actuated 40 unit switches and two airspeed switches (or two printed circuit boards comprising an ARINC 429 receiver and power supply). The airspeed switches are installed in the lower nose compartment on the right side. Input to the airspeed switches is from the auxiliary pitot static probes on the left and right sides of the fuselage adjacent to the flight compartment. The system can be tested by using a three position test switch and green indicator light in the Electrical/Electronic Equipment Compartment.


B737-300/-400/-500



B737-300/-400/-500 27-50

FLAP LEVER 40 UNIT SWITCHES (2)


Figu Figure re 90

Flap Flap Load Load Limi Limite terr Compo Compone nent nt Loc Locat atio ion n



27-50 FLAP LOAD LIMITER (CONFIG. 1) Control The load limiter system automatically retracts the flaps from 40 to 30 units when airspeed exceeds a range of 152 to 162 knots. Power is applied to the load limiter solenoid valve thru two separate airspeed switches and two 40 unit flap lever switches. The load limiter system operates only when hydraulic power is available. Operation The flap 40 unit switches are actuated by a cam when the flap control lever is moved to the 40 unit position. Electric power is 28 volts dc from bus 2. When the airspeed switches senses an airspeed in excess of a range of 152 to 162 knots, the hydraulic solenoid valve attached to the trailing edge flap control valve is energized. Energizing the solenoid valve positions the trailing edge flap control valve to the 30-unit position, allowing pressurized fluid to flow to the hydraulic motor. As airspeed decreases to below a range of 147 147 to 157 knots, the solenoid valve is de-energized, the control valve is positioned to the 40-unit position, and the flaps extend to 40 units. The flap load relief light comes on when the flap load limiter system is operating the flaps. BITE A three position FLAP LOAD LIMITER test switch and a green test light are located on a panel mounted on the inboard forward stanchion of the E2 rack. The switch is powered through the ground position of the R277 air/ground sense relay contact.


B737-300/-400/-500

Moving the switch to SYSTEM TEST energizes the solenoid through the flap lever 40 unit switches. The flaps are observed to move from 40 units to 30 units. Moving the switch to SWITCH TEST checks the airspeed switches in the low position. Electric power is applied through the low position of the airspeed switches and the flap lever switches not in the 40 unit position and turns on the green light.



B737-300/-400/-500 27-50

28V DC BUS 2

AIRSPEED

160 KT

MASTER A

FLAP

DIM

A

FLAP LOAD

LOAD RELIEF

RELIEF LIGHT P2 CENTER INSTRUMENT

P6-2

PANEL L814

FLAP LOAD LIMIT AIRSPEED SWITCH S560

AIRSPEED

160 KT

AIR FLAP LOAD RELIEF LIGHT SW TEST

RELAY R161

(KEYWAY)

GND M338 LANDING GEAR MOD

FLAP LOAD LIMIT AIRSPEED SWITCH S562


DC

NC

FLAP LOAD

NO

TEST SW SYS TEST

FLAP

C FLAP

NO

G

DC

HANDLE

HANDLE

LOAD LIMITER

40 S5 S 564

40

TEST LT L558

S566

AIRPLANES WITH AIRSPEED SWITCHES

Figur Figure e 91

SOLENOID VALVE V94

NC C

LIMITER

FLAP LOAD LIMITER

Flap Flap Load Load Limi Limite terr System System Cir Circu cuit it (con (config fig.. 1)



27-50 FLAP LOAD LIMITER (CONFIG 2) Control The load limiter system automatically retracts the flaps from 40 to 30 units when airspeed exceeds a range of 152 to 162 knots. Power is applied to the load limiter solenoid valve thru two separate ARINC 429 receiver switch cards and two 40 unit flap lever switches. The load limiter system operates only when hydraulic power is available. Operation The flap 40 unit switches are actuated by a cam when the flap control lever is moved to the 40 unit position. Electric power is 28 volts dc from bus 2. When the airspeed function in excess of a range of 152 to 162 knots was derived (via ADC no.1), the hydraulic solenoid valve valve attached to the trailing edge flap control valve is energized. energized. Energizing the solenoid valve positions positions the trailing edge flap control valve to the 30-unit position, allowing pressurized fluid to flow to the hydraulic motor. As airspeed decreases to below a range of 147 to 157 knots, the solenoid valve is de-energized, the control valve is positioned to the 40-unit position, and the flaps extend to 40 units. The flap load relief light comes on when the flap load limiter system is operating the flaps.


B737-300/-400/-500

BITE A FLAP LOAD LIMITER test pushbutton and a green test light are located on a panel mounted on the inboard forward stanchion of the E2 rack. The pushbutton is powered through the ground position of the R277 air/ground sense relay contact. Pressing the switch, together with a airspeed signal provided by the ADC no.1 (when respective test switch was selected on the front of the ADC) energizes the solenoid through the selected flap lever 40 unit switches. The flaps are observed to move from 40 units to 30 units.



B737-300/-400/-500 27-50

28V DC BUS 2 PC BOARD

MASTER

(POWER SUPPLY

DIM A

FOR ARINC FLAP

429

LOAD RELIEF

RECEIVER)

FLAP LOAD RELIEF LIGHT

P6-2

M1452

P2 CENTER INSTRUMENT PANEL L814

PC BOARD (ARINC 429 RECEIVER)

AIR

GND

M1453 FLAP LOAD LIMITER SOLENOID VALVE V94 NO

FLAP LOAD


DC

E11 LANDING GEAR LOGIC SHELF

R277 GND SENSING SQUAT RELAY LIMITER

NC

TEST SW

NC SYS TEST

NC C

NO

C NO

DC

FLAP

HANDLE

HANDLE

LOAD LIMITER

40 S5 S564

40 S5 S566

TEST LT L558

AIRPLANES WITH ARINC 429 RECEIVER

Figur Figure e 92

G

FLAP

Flap Flap Load Load Limi Limite terr System System Cir Circu cuit it (con (config fig.. 2)

A



27-50 FLAP TRANSMISSION TRA NSMISSION INSTALLATION Purpose The flap transmission assembly converts flap torque tube rotation into linear motion to extend or retract the trailing edge flaps. Location The transmission assemblies are mounted on the flap tracks just aft of the landing gear support beam and the wing rear spar. Physical Description/Features The flap drive system contains eight transmission assemblies, numbered from left to right, 1 through 8. Each transmission assembly is enclosed in a flap track fairing and consists of a transmission gearbox, a universal joint, and a ball nut and jack screw actuator. Each transmission gearbox incorporates a torque limiter; two springs wound together, bevel gears and input and output shafts inside a housing. If a flap jams, excessive torque on the screw actuator will will cause the springs to expand and bind against the housing. Excessive torque is absorbed by the torque limiter until the flap hydraulic motor stalls. The torque limiter operates in either direction of flap travel. The inboard transmission assembly on each flap incorporates a no-back friction brake. At extended flap positions, this brake prevents flap retraction due to airloads. The brake disengages when the jackscrew is operated.


B737-300/-400/-500

Operation The actuator consists of a jackscrew and a recirculating ball bearing nut. The jackscrew is connected to the transmission transmission gearbox by a universal joint which allows angular deflection of the jackscrew during flap operation. The ball bearing nut is attached to the mid flap through a gimbal assembly. During flap operation, torque tube drive rotation is transmitted through the transmission gearbox to the jackscrew. The ball bearing nut is restrained from turning and travels fore and aft on the rotating screw to extend and retract the flaps.



B737-300/-400/-500 27-50

FWD

INBD NO-BACK BRAKE (inboard gearboxes only)

TORQUE LIMITERS (each transmission) ANGLE GEARBOX NO. 3

ANGLE GEARBOX NO. 4

NO. 5

NO. 6

NO. 7

NO. 2 FLAP POWER UNIT GEARBOX

INBD MIDFLAP

NO. 1

RIGHT ANGLE GEARBOX


NO. 8

INBD MIDFLAP

RIGHT FLAP RIGHT ANGLE GEARBOX (REVERSE TURNIG DIRECTION)

TRANSMISSION ASSEMBLY (8 PLACES)

OUTBOARD MIDFLAP

OUTBOARD MIDFLAP

Figu Figure re 93

Flap Flap Trans ransmi miss ssio ion n Sche Schema matic tic



B737-300/-400/-500 27-50





B737-300/-400/-500 27-50

SEE

A

TORQUE TUBE

UNIVERSAL JOINT

OUTBOARD mid flap


FLAP TRACK TORQUE LIMITER HOUSING TRANSMISSION GEARBOX BALL NUT AND SCREW ACTUATOR FWD

INBD

Figu Figure re 94

Flap Flap Trans ransmi miss ssio ion n Asse Assembl mbly y

A



27-50 OUTBOARD MID FLAP DRIVE MECHANISM Purpose The mid flap drive mechanism positions each flap section by positioning the mid flap. Physical Description/Features Two flap tracks, curved forged steel beams mounted on the lower surface of the wing, are installed for each flap segment. The flap is supported by flap carriages attached to the mid flap that travel on these tracks. Rollers on each carriage assembly support the flap on the track and side load rollers provide lateral alignment. A preload spring unit is mounted on the mid flap to simulate flight loads on the flap assembly when the flaps are retracted on the ground. Operation The three flap sections nest together with the flaps retracted. A sequencing carriage attached to the foreflap has a toggle set in a detent in the carriage. When the jackscrews drive the mid flap out of retract, the carriage rolls aft on the track. The foreflap is carried aft by the sequencing carriage during the first segment of flap extension.


B737-300/-400/-500




B737-300/-400/-500 27-50

Figur Figure e 95

Outbo Outboar ard d Mid Mid Fla Flap p Dri Drive ve Mech Mechan anis ism m



27-50 OUTBOARD FORE FLAP DRIVE MECHANISM Purpose The foreflap drive mechanism is installed to separate the foreflap from the mid flap during extension. Physical Description/Features Three curved support beams extend through the foreflap lower surface and connect to three foreflap tracks. The tracks ride on roller bearing mounted to mid flap structure. Operation As the mid flap extends the foreflap toggle assemblies ride on the mid flap carriages until a position corresponding to 8 units. At this point, aft movement of the toggle assembly is stopped by a lug protruding from the upper surface of the track. The forward roller bearing on the toggle assembly drops in a detent on the track, locking the foreflap into position. The mid flap continues to roll away on the foreflap tracks.


B737-300/-400/-500




B737-300/-400/-500 27-50

Figur Figure e 96

Outbo Outboar ard d Fore Fore Flap Flap Drive Drive Mech Mechan anis ism m



27-50 OUTBOARD FLAP FAIRING DRIVE Purpose The outboard flap fairing drive positions the hinged section of the flap track fairing to maintain a clearance between the flap and fairing as the flap changes position. Physical Description/Features The outboard flap track fairings consist of a fixed section and an aft moveable section that rotates about a hinge. The drive mechanism consists of a fairing support arm with rollers that ride in two fairing cam tracks. Operation Motion of the mid flap move the fairing support arm rollers in the cam track, causing the aft fairing to pivot.


B737-300/-400/-500




B737-300/-400/-500 27-50

Figu Figure re 97

Outb Outboa oard rd Fla Flap p Fair Fairin ing g Driv Drive e



27-50 OUTBOARD AFT FLAP DRIVE Purpose The outboard aft flap drive separates the aft flap from the mid flap based on mid flap position. Physical Description/Features A bellcrank is mounted on and pivots about the fairing support arm. A pushrod connects one end of the bellcrank to the aft flap. A roller on the other end rides in a cam track in the center of the fairing. Operation As the fairing pivots in response to carriage movement on the track, the bellcrank cam arm repositions the bellcrank. This moves the aft flap pushrod to actuate the aft flap for its initial movement. When the mid flap approaches full travel, the bellcrank and pushrod move the aft flap again, increasing the slot between it and the mid flap.


B737-300/-400/-500



B737-300/-400/-500 27-50

OUTBOARD FOREFLAP

FLAP TRACK FORWARD FAIRING

OUTBOARD MIDFLAP

FOREFLAP SEQUENCING CARRIAGE FOREFLAP

MIDFLAP TRACK

OUTBOARD OUTBOARD TRACK FAIRING AFTFLAP SEE

A

MIDFLAP CARRIAGE

MIDFLAP

FLAP TRACK FAIRING SUPPORT ARM

AFTFLAP PUSHROD

BELLCRANK


AFTFLAP BELLCRANK CAM TRACK A FAIRING CAM TRACK

FLAP TRACK AFT FAIRING

Figu Figure re 98

Outb Outboa oard rd Aft Aft Flap Flap Driv Drive e

INBOARD FLAP TRACK FAIRING



27-50 INBOARD MID FLAP DRIVE MECHANISM Purpose The inboard mid flap drive mechanism positions each of the three flap sections by positioning the midflap. Physical Description/Features The mid flap is supported by two flap carriages which ride on main flap tracks. A tubular support at each end of the mid flap attach it to the carriage. Another tubular support at each end of the fore flap attach it to the toggle assembly. A cam track in which the forflap toggle assembly rides is attached to the carriage. Operation The operation of the inboard mid flap drive mechanism is the same as that described for the outboard flap. One difference in the mechanism is that the main carriage cam continues aft by riding on rollers attached to the fore flap toggle assembly.


B737-300/-400/-500




B737-300/-400/-500 27-50

Figur Figure e 99

Inboa Inboard rd Mid Mid Flap Flap Drive Drive Mech Mechan anis ism m



27-50 INBOARD AFT FLAP DRIVE MECHANISM Purpose The inboard aft flap drive mechanism separates the aft flap from the midflap as programmed by mid flap position. Location The inboard aft flap drive mechanism is located at the inboard side of the flap and inside the mid flap. Physical Description/Features The aft flap drive mechanism consists of a cable drum with cam slot, a bellcrank with cam follower, a slave bellcrank, two pushrods, an actuating mechanism boom, an actuating cable support fitting and connecting cables. The boom, which pivots about a bracket on the inboard flap track, is not structurally attached to the mid flap. The cable support fitting is mounted on the inboard midflap carriage. A cable mounted at each end of the boom passes over the cable support fitting and attaches to the cable drum inside the midflap. A cam follower on the inboard bellcrank rides in the cable drum cam slot. The slave bellcrank is connected to the inboard bellcrank by two cables. Both bellcranks are connected to the aft flap by pushrods. Operation The boom rotates down as the midflap moves in the extend direction and vice versa. The aft flap drive mechanism is actuated by the change in relative position between the boom and the cable support fitting. The bellcranks provide aft flap motion through the pushrods. The aft flap makes two distinct movements, near the beginning and at the end of mid flap travel.


B737-300/-400/-500



B737-300/-400/-500 27-50

FWD INBOARD AFT FLAP DRIVE MECHANISM

INBD MIDFLAP CARRIAGE

AFTFLAP ACTUATING CABLES SLAVE BELLCRANK

INBOARD BELLCRANK

AFTFLAP ACTUATING CABLES


AFTFLAP PUSHROD

INBOARD CABLE DRUM

CAM SLOT

CAM FOLLOWER

ACTUATING CABLE SUPPORT FITTING

AFTFLAP ACTUATING CABLE

INBOARD AFTFLAP CLUTCH INBOARD AFTFLAP

AFTFLAP ACTUATING MECHANISM BOOM

AFTFLAP PUSHROD EXHAUST GATE

Figur Figure e 100 100

AFTFLAP ACTUATING CABLES BOOM ROLLER

Inboa Inboard rd Aft Aft Flap Flap Dri Drive ve Mech Mechan anis ism m



B737-300/-400/-500 27-50






B737-300/-400/-500 27-50

FLAPS UP


FLAPS 40 UNITS

Inboa Inboard rd Aft Aft Flap Flap Dri Drive ve Mech Mechan anis ism m



27-50 INBOARD AFT FLAP CLUTCH Purpose A clutch is installed in the master bellcrank assembly to stop aft flap retraction when a foreign object is lodged between the midflap trailing edge and the aft flap leading edge. Location Clutch discs are installed between the bellcrank and the cam arm containing the cam roller, inside the mid flap. Physical Description/Features The clutch assembly contains a cam arm clutch disc, a bellcrank clutch disc and bellcrank springs and attach fittings. Operation If a force greater than 300 pounds is imposed on the bellcrank arm during flap retraction the springs will be overcome and the clutch discs will disengage. The cam follower will still operate but no motion will be transmitted to the bellcrank assembly. The clutch mechanism will automatically reset when the flaps are extended to 40 units. Aft flap track mechanical stops will hold the aft flap as the cam follower rotates to the correct position.


B737-300/-400/-500




B737-300/-400/-500 27-50

Figu Figure re 102 102

Inbo Inboar ard d Aft Aft Flap Flap Clut Clutch ch



27-50 EXHAUST GATES Purpose Exhaust gates are installed to provide an aerodynamic surface that can be rotated out of the path of the engine plume at high flap settings. Location An exhaust gate is hinged to the inboard end of each outboard mid flap and to the outboard end of each inboard mid flap. Physical Description/Features There is an exhaust gate actuation mechanism installed on the inboard end of the outboard mid flap. On some airplanes the exhaust gate actuation mechanism consists of two pushrods and a mechanism assembly. The mechanism assembly is attached to the mid flap at the front and rear spars. The forward pushrod that is attached to the foreflap, moves the forward part of the mechanism. When the forward part of the mechanism moves past a certain position, it moves the aft part of the mechanism. As the aft part of the mechanism mechanism moves, it moves the aft pushrod which moves the exhaust gate. On other airplanes the exhaust gate actuation mechanism consists of three pushrods, a link, a mechanism housing, a bellcrank, and a cam. The housing assembly is attached to the mid flap between the front and rear spars. The forward pushrod that is attached to the fore flap, moves the link that is attached to the cam through the center pushrod. As the cam turns past a certain position, a follower on the bellcrank moves the bellcrank and the aft pushrod moves the exhaust gate.


B737-300/-400/-500

Operation A pushrod attached to the fore flap actuates the link. The link is attached to the cam through a second pushrod. As the cam rotates, a follower on the bellcrank moves the bellcrank causing the attached pushrod to actuate the exhaust gate. As the mid flap and fore flap separate between the 25 and 40 unit flap positions, the mechanism progressively opens the exhaust gate. The exhaust gate is raised to a maximum of 30 degrees at the 40 unit flap position.



B737-300/-400/-500 27-50

EXHAUST GATE ACTUATION MECHANISM MIDFLAP

EXHAUST GATE A

A EXHAUST GATE PUSHROD

FOREFLAP CAM

BOLT

SHEAR RIVET

FLAP-UP RIG PIN HOLE

MIDSPAR CUTOUT

BOLT

FOREFLAP PUSHROD INBD

MIDFLAP

EXHAUST GATE

AIRPLANES WITH TWO PUSHRODS ACTUATING MECHANISM


PUSHROD

BOLT PUSHROD SHEAR RIVET

MIDFLAP

FOREFLAP INBD AIRPLANES WITH THREE PUSHRODS A-A

Fig Figure 103

Exhaust Gate Gates s



27-50 TRAILING EDGE BYPASS VALVE Purpose The trailing edge flap bypass valve interconnects the flap hydraulic motor pressure and return ports to prevent it from operating during operation of the flap alternate drive system. Location The valve is mounted on the aft wall in the right main gear wheel well. Physical Description/Features The flap bypass valve is a motor-operated two position valve containing three ports from which hydraulic lines extend to the flap control valve and the flap hydraulic motor. Power The valve is operated by 28 volts dc from bus number 1. Control The valve is normally controlled by the alternate flap arm switch. It is also controlled by the flap asymmetry shutoff relay. A position indicator and manual lever allows positioning the valve manually when electric power is removed.


B737-300/-400/-500



B737-300/-400/-500 27-50

SEE A

FLAP ALTERNATE DRIVE UNIT

MACH TRIM SWITCH S290

LANDING WARNING SWITCH S138 WFFB CABLE

FLAP 10DEGREE SWITCH S584

TORQUE TUBE

WFFA CABLE

TE FLAP BYPASS VALVE

TE FLAP CONTROL VALVE


FLAP LOAD LIMITER SOLENOID VALVE FLAPS UP LIMIT SWITCH S245 FLAPS DOWN LIMIT SWITCH S246

A

FLAP CONTROL UNIT LE FLAP CONTROL VALVE TAKEOFF WARNING SWITCH S130 STALL WARNING S856


Trail railin ing g Edge Edge Bypa Bypass ss Val Valve ve Loca Locati tion on

FLAP POWER UNIT




B737-300/-400/-500 27-50

TRAILING EDGE BYPASS VALVE Operation Moving the alternate flaps arm switch to ARM applies electric power to position the valve to bypass. The down line from the flap control valve is blocked and the flap hydraulic motor ports are connected together to allow fluid circulation within the motor. Returning the alternate flaps arm switch to OFF moves the bypass valve to normal and restores hydraulic operation of the flap system. The flap asymmetry shutoff relay is energized to move the valve to bypass and stop flap operation by the hydraulic motor when left and right wing flap movement is not symmetrical, or cable tension to the control valve is lost.



B737-300/-400/-500 27-50

FLAP SHUTOFF VALVE

S1 ALT FLAP MASTER ARM SWITCH ARMED NORMAL

28V DC OFF

FLAP POS IND

BYPASS R123 FLAP ASYMMETRY SHUTOFF RELAY

28V AC TRANSFER BUS NO. 2 P6 PNL


S54 FLAP POSITION INDICATOR COMPARATOR SWITCH


SENSED BAR (MAGNET)

S816 WFA FLAP CONTROL CABLE TENSION SWITCH

Trail railin ing g Edge Edge Bypa Bypass ss Val Valve ve Circ Circui uitt

V52 TE FLAPS BYPASS VALVE




B737-300/-400/-500 27-50

ALTERNATE ALTERNA TE FLAP D RIVE UNIT Purpose The flap alternate drive unit uses electrical power to drive the flap system if a failure prevents normal hydraulic operation of the flap system. Location The alternate drive unit is mounted on the flap power unit located on the aft wail in the main gear wheel well. Physical Description/Features A 115 volt ac motor and a gearbox are the primary alternate drive unit components. The gearbox incorporates a double planetary reduction gear train and a disconnect and overload clutch. The output shaft drives the same torque tube that is normally driven by the hydraulic motor. The ring gear of the first planet is fixed to the motor housing. The input planet system carrier gears rotate around an input sun gear, cut on the motor drive shaft. The output sun gear is attached to the input planet system carrier gears. Rotation of the output sun gear drives the second planet system carrier gears attached to the output shaft. The output ring gear is a floating gear held fixed by a cable assembly during motor operation. The two terminals on the cable assembly are held by a spring mechanism consisting of a sensing spring, return spring, and a spring collar. Since the normal spring load on the cable is not sufficient to prevent rotation of the output ring gear, a mechanism is provided to accomplish that function. It consists of a yoke, energizing pin, bellcrank, and a solenoid. One arm of the bellcrank is attached to the solenoid and the energizing pin is inserted into a hole in the other arm of the bellcrank. The opposite end of the energizing pin is attached to the yoke. The ends of the yoke are attached to the cable terminals. Control Operation of the alternate drive motor is controlled by actuation of the alternate flaps arm switch to ARM and toggling the control switch.

Operation The solenoid is energized by rectified ac power simultaneously with application of power to the motor. Energizing the solenoid causes the bellcrank to drive the yoke, compressing the return springs and increasing cable tension. This locks the output ring gear, allowing the motor to drive the output shaft. If binding occurs in the gear train, damage to the unit is prevented by the load sensing spring compressing to relieve tension in the cable, allowing the output ring gear to slip. When the·flap system is being driven by the hydraulic motor, the output shaft is rotating. Since the alternate unit solenoid is de-energized, the cables are not· under tension and the output ring gear is allowed to rotate with the carrier gears. This prevents transmission of motion to the electric motor shaft.



B737-300/-400/-500 27-50

LOAD SENSING SPRING

SOLENOID

CABLE TERMINAL

RETURN SPRING

YOKE MOTOR

ENERGIZING PIN

RETURN SPRING BELLCRANK

RING GEAR

LOAD SENSING


SPRING

INPUT PLANETARY GEAR ASSEMBLY

OUTPUT SHAFT CABLE ASSEMBLY

RING GEAR OUTPUT PLANETARY GEAR ASSEMBLY


Alte Altern rnat ate e Flap Flap Dri Drive ve Unit Unit



27-50 TRAILING EDGE FLAP POSITION INDICATING Purpose The trailing edge flap position indicating components provide visible indication of the angular position of the trailing edge flaps. Location The flap position indicator is mounted on the center instrument panel in the flight compartment. It receives signals from two position transmitters mounded on the outboard flap torque tube in each wing. The left transmitter is between the numbers one and two transmission assemblies and the right is between the numbers seven and eight assemblies. Physical Description/Features Each position transmitter assembly consists of a synchro-type transmitter mounted on a gearbox. The gearbox is driven by the outboard flap torque tubes and in turn drives the transmitter mounted on the gearbox housing. A dual synchro-type indicator registers the flap position in units. Power The position indicating system is driven by 28 volts ac power from the number 2 transfer bus. Operation During flap operation, rotation of the outboard torque tube drives the flap position transmitter synchros through the gearbox. The transmitter synchros send electrical signals to the dual indicator synchros to move the two needles and reflect the angular position of left wing and right wing trailing edge flaps.


B737-300/-400/-500



B737-300/-400/-500 27-50 SEE A

2

5

SEE B

1 10 UP

15

FLAPS

25 40

NOTE: BLACK INDICATOR DIAL WITH WHITE MARKINGS REVERSED FOR CLARITY

30

SEE B

POSITION INDICATOR A

FLAP POSITION TRANSMITTER

FWD


B

INBD

OUTBOARD FLAP TORQUE TUBE

LEFT WING SHOWN-RIGHT WING OPPOSITE GEARBOX

Figure Figure 107

Trail Trailing ing Edge Edge Positi Position on Indic Indicati ation on Compo Component nent Locati Location on



27-50 FLAP ASYMMETRY CONTROL CIRCUIT Operation/Control Sequence The flap asymmetry control system stops hydraulic operation of the trailing edge flaps when a specified difference exists between the position of the flaps on the left wing and those on the right wing.

A comparator switch inside the flap position indicator controls the flap asymmetry relay mounted on the back of the E3-2 rack in the electronic equipment compartment. An asymmetry test panel is mounted on the forward stanchion of the E3 rack. The system consists of the two flap position transmitters, the dual position indicator with comparator switch, the asymmetry shutoff relay, and the trailing edge flap bypass valve. The test panel consists of an electric circuit, an asymmetry test switch, and a green test light. The asymmetry system functions only during hydraulic operation of the flaps. During flap operation, a difference in position between the left and right wing trailing edge flaps is detected by the flap position indicating system. When the two indicator pointers separate by a predetermined amount the comparator switch closes and applies power to the asymmetry shutoff relay. The energized relay drives the trailing edge flap bypass valve to bypass and stops the hydraulic motor. The test panel is provided for self test of the asymmetry protection system. Actuation of the switch to TEST LEFT or TEST RIGHT causes the pointers on the flap position indicator to separate, the bypass valve to move to bypass, and the green light to illuminate. The green light illuminates whenever the flap bypass valve is in the bypass position. y l n O s e s o p r u P g n i n i a r T r o F

B737-300/-400/-500



B737-300/-400/-500 27-50

28V AC TRANSFER BUS NO.2

TO CONTROL PANEL LIGHTING CIRCUIT

TE FLAP POSITION IND P6-2

RIGHT

RIGHT FLAP POS TRANSMITTER TEST LEFT-UP NORM FLAP ASYMMETRY TEST SW

TEST RIGHTDN

FLAP POSITION COMPARATOR SWITCH

LEFT


LEFT FLAP POS TRANSMITTER

FLAP POSITION INDICATOR

28V DC BATTERY BUS DIM & TEST CONTROL

TO FLAP BYPASS VALVE

P6-3

FLAP ASYMMETRY TEST LIGHT 3 1 2

TO ASYMMETRY SHUTOFF RELAY (E3-2)

ELECTRONIC EQUIPMENT RACK E3

Figure Figure 108

Flap Flap Pos. Pos. Indica Indicat. t. and Asy Asym. m. Control Control Circui Circuitt




B737-300/-400/-500 27-50

T. E. FLAP SYSTEM HYD RAULIC OPERATION Normal Sequence The trailing edge flaps are normally operated by hydraulic power. When the flap control lever is rotated, cables displace the flap control unit linkage. This positions the trailing edge flap control valve to port pressurized fluid from hydraulic system B to the flap hydraulic motor. The motor drives the flaps through a torque tube drive and transmission assemblies. As the flaps move, cables from a follow-up drum on the power unit rotate a follow-up mechanism on the control unit. A cam in the follow-up mechanism returns the control valve slide so that it nulls and stops the flow of hydraulic fluid when the desired flap position is reached. The operating speed of the trailing edge flaps is controlled by a flow limiting valve installed in the pressure line to the flap control valve. A priority valve stops fluid flow to the hydraulic motor when B system demand causes pressure to drop below 2400 psi at the valve. A motor operated bypass valve in the hydraulic lines between the control valve and the flap power unit controls operation of the hydraulic motor. The normal position of the bypass valve directs pressure thru one line to the motor and opens the other line to return. This allows the motor to run when the control valve is actuated.



B737-300/-400/-500 27-50

NORMAL

28V DC DC BUS NO.1

FLAP VALVES

OFF

BYPASS FLAP ASYMMETRY SHUTOFF RELAY

ARMED ALTERNATE FLAPS ARM SWITCH

PRIORITY VALVE BYPASS

NORM

TE FLAP CONTROL VALVE

TE BYPASS VALVE


PRESSURE RETURN

FLAP POWER UNIT

Figure Figure 109

CONDITION: FLAPS EXTENDING

Trail Trailing ing Edge Edge Flap Flap System System Hydrau Hydraulic lic Operat Operation ion

SYSTEM B PRESSURE

SYSTEM B RETURN




B737-300/-400/-500 27-50

FLAP ALTERNATE DRIVE SYSTEM CIRCUIT Backup Operation A flap alternate drive unit operates the flaps electrically when required instead of hydraulic power. Two switches operate the system; the alternate flaps arm switch and the alternate flaps control switch on the pilots’ forward overhead panel. Actuating the alternate flaps arm switch to ARM supplies 28 volt dc power to the control switch and simultaneously positions the bypass valve to BYPASS. The hydraulic motor is disengaged because both lines at the motor are connected to the same port of the control valve. Moving the control switch up or down energizes the respective relay when the applicable limit switch is closed. The motor is powered by 115 volts ac through the relay contacts until the limit switch opens or the control switch is returned to OFF. When the control switch is placed down to extend the trailing edge flaps electrically, the leading edge standby shutoff valve relay is energized. The valve opens and standby hydraulic system pressure extends the leading edge flaps and slats.



B737-300/-400/-500 27-50 CABLE TENSION SWITCH NORMAL

OR FLAP ASYMMETRY

BYPASS FLAP ASYMMETRY

115V AC

SHUTOFF RELAY (E3-2)

NO. 2

TE FLAP BYPASS VALVE

TRANSFER BUS

TE FLAP DOWN LIMIT SWITCH S246

TE ALT FLAP DRIVE MOTOR P6 UP RELAY (J4) R58

DN UP

To STBY PUMP

ALTERNATE

TE FLAP UP LIMIT SWITCH S245

FLAP CONTROL

SWITCH S2

UP

OFF DOWN


28V DC DC BUS NO. 1

FLAP ALTERNATE DRIVE MOTOR

OFF DOWN

ARM

RELAY

FLAP VALVES P6

(J4) R57

ALTERNATE

FLAP ARM

OPEN

SWITCH

CLOSE

LE FLAP STANDBY SHUTOFF VALVE RELAY K3 FLIGH FLIGHT T CONTRO CONTROLS LS MODU MODULE LE

Figur Figure e 110

LE FLAP STANDBY DRIVE SHUTOFF VALVE VALVE

(P5-3) (P5-3)

Flap Flap Alte Altern rnat ate e Dr Driv ive e Syst System em Cir Circu cuit it


FLIGHT CONTROLS LIFT AUGMENTING

27-80

27-80

LIFT AUGMENTING

LEADING EDGE FLAP AND SLATS INTRODUCTION Purpose High lift leading edge devices are used in combination with the trailing edge flaps to allow airplane operation from short runways. The use of leading edge devices allows a change in wing camber which greatly increases lift. System Description/Features Three leading edge slats are installed outboard of the engine and two leading edge flaps are installed inboard of the engine on each wing. The flaps and slats are numbered from left to right as depicted on the graphic. Normal operation of the leading edge devices is by hydraulic system B. Control is with a control valve on the flap control unit positioned by the follow-up cables from the flap power unit.


B737-300/-400/-500




B737-300/-400/-500 27-80

Figure Figure 111

Leadin Leading g Edge Edge Flaps Flaps and Slat Slat Introdu Introducti ction on



27-80 LEADING EDGE FLAP SLAT SYSTEM General Component Locations Two Krueger-type leading edge flaps are installed on each wing, inboard of the engine. Hydraulic two-position actuators, used to extend and retract each flap, are mounted on the wing front spar with each rod end connected to the respective flap. Three leading edge slats are installed on each wing, outboard of the engine. Hydraulic three-position actuators, used to extend and retract each slat, are mounted on the wing front spar with the rod end connected to the respective slat. Control components are located in the main gear wheel well. General Subsystem Features Leading edge flaps are two position devices, retract and extend. The flaps are hinged to the leading edge of the wing and are retracted to the underside of the wing. A folding nose section rotates and is stored on the underside of the wing when the flap is retracted. Leading edge slats are three position devices, retract, extend (intermediate) and full extend. The slats function as the wing leading edge when retracted. They are attached to tracks which ride on rollers in the wing leading edge to extend. System Interfaces Extension and retraction of the leading edge devices is programmed based upon position of the trailing edge flaps. Normal hydraulic operation of the leading edge flaps and slats is controlled by the trailing edge flap follow-up system operating a control valve on the flap control unit.


B737-300/-400/-500

General Operation Normal hydraulic operation of the leading edge devices is by system B, controlled by the flap lever. The leading edge flaps and slats are retracted when the flap lever is at 0. The leading edge flaps go to full extend and the slats to their intermediate position, extend, when the flaps move between 0 and 1. The slats go to full extend when the flaps move between 5 and 10. This movement is reversed on retraction. Alternate operation is by the standby hydraulic system through a leading edge device shutoff valve, controlled by the alternate flap control switches. THe leading edge devices can only be extended, with no intermediate positioning of the slats, during alternate flap operation. An autoslat system is installed that will automatically extend the leading edge slats from the intermediate to full extend position, if required, to provide additional lift. Autoslat operation is normally accomplished by system B hydraulic power. A power transfer unit provides hydraulic power for autoslat operation when the right engine pump output pressure is low, the nose gear is off the ground, and the trailing edge flaps are at positions 1, 2, or 5.



B737-300/-400/-500 27-80

FLAP CONTROL LEVER

FLAP 3

FLAP 4

SLAT 4

SLAT 5

SLAT 6

FLAP 2 FLAP 1 FLAP ACTUATOR (TYP)

SLAT ACTUATOR (TYP) SLAT 3


SLAT 2 SLAT 1

HYDRAULIC FUSE FLOW LIMITING VALVE LEADING EDGE STANDBY DRIVE SHUTOFF VALVE

Figu Figure re 112

LEADING EDGE FLAP AND SLAT CONTROL VALVE

Lead Leadin ing g Edge Devi Device ces s Compon Componen entt Locat Locatio ion n




B737-300/-400/-500 27-80

LEADING EDGE OPERATION General Operation Normal hydraulic operation of the leading edge devices is by system B, controlled by the flap lever. The leading edge flaps and slats are retracted when the flap lever is at 0. The leading edge flaps go to full extend and the slats to their intermediate position, extend, when the flaps move between 0 and 1. The slats go to full extend when the flaps move between 5 and 10. This movement is reversed on retraction. Alternate operation is by the standby hydraulic system through a leading edge device shutoff valve, controlled by the alternate flap control switches. THe leading edge devices can only be extended, with no intermediate positioning of the slats, during alternate flap operation. An autoslat system is installed that will automatically extend the leading edge slats from the intermediate to full extend position, if required, to provide additional lift. Autoslat operation is normally accomplished by system B hydraulic power. A power transfer unit provides hydraulic power for autoslat operation when the right engine pump output pressure is low, the nose gear is off the ground, and the trailing edge flaps are at positions 1, 2, or 5.




B737-300/-400/-500 27-80

Figure Figure 113

Leading Leading Edge Edge Devi Device ce Bas Basic ic System System Schema Schematic tic




B737-300/-400/-500 27-80

LEADING EDGE FL AP AND SLAT CONTROL VALVE

LEADING EDGE STANDBY DRIVE SHUTOFF VALVE

Purpose A three-position control valve regulates operation of the leading edge flaps and slats.

Purpose The leading edge standby drive shutoff valve is provided to control alternate operation of the leading edge flaps and slats.

Location The control valve is mounted on the trailing edge flap follow-up mechanism located in the right wheel well.

Location The shutoff valve is on the right side of the standby hydraulic module which is mounted on the main wheel well aft wall, above the keel beam.

Physical Description/Feature The valve consists of a sliding piston enclosed in a valve housing. Drilled passages are provided in the valve housing for a pressure port, a return port, and two cylinder ports. A pushrod connects the control valve to the trailing edge flap follow-up mechanism. With the trailing edge flaps retracted, the control valve slide blocks pressure to the leading edge flap and slat actuator extend ports. When the trailing edge flaps extend to the 1- to 5-unit position, pressurized fluid is ported through one of the cylinder ports to extend the leading edge flaps and to extend the leading edge slats to the intermediate position. When the trailing edge flaps reach the 10-unit position, pressurized fluid is ported through both control valve cylinder ports to fully extend all leading edge slats.

Physical Description/Feature The shutoff valve is a 28 volt dc, two-position valve in the standby pressure line to the leading edge flap and slat actuators. Control The shutoff valve is normally closed. It opens and allows standby pressure to extend the flaps and fully extend the slats when the alternate flap master switch is at ARM and the alternate flap control switch is moved DOWN.



B737-300/-400/-500 27-80 STANDBY RUDDER ACTUATOR SHUTOFF VALVE

LEADING EDGE STANDBY DRIVE SHUTOFF VALVE

WHEEL WELL AFT BULKHEAD

B

HYDRAULIC FUSE FLOW LIMITING VALVE

LEADING EDGE FLAP AND SLAT CONTROL VALVE

LEADING EDGE STANDBY DRIVE SHUTOFF VALVE SEE B


TRAILING EDGE FLAP CONTROL VALVE

FLAP CONTROL UNIT

FWD

INBD

A

Figure Figure 114

Leadin Leading g Edge Edge Slat Slat and Falp Falp Compo Componen nentt Locat Location ion

LEADING EDGE FLAP AND SLAT CONTROL VALVE SEE A



27-80 LEADING EDGE FLAP ACTUATOR Purpose Hydraulic actuators convert available hydraulic power to mechanical energy that positions each leading edge slat to the desired position; retract, extend, or full extend. Hydraulic actuators convert available hydraulic power to mechanical energy that extends or retracts each leading edge flap. Location Each actuator is mounted with self-aligning bearings between the wing front spar and a fitting on the leading edge flap. Physical Description/Features Each actuator housing contains two identical blocking valve and a piston and rod assembly. The blocking valves will hydraulically lock the actuator piston in position, if hydraulic pressure drops below 2000 psi. Upon pressure loss, a compression spring moves the valve slide to lock pressure in the actuator. Separate blocking valves are used for the standby system and for pressure from the control valve. External ports in the actuator housing are provided for a retract pressure line, an extend pressure line, and a standby pressure line. Control Hydraulic pressure control at the actuator is by the blocking valves. When one blocking valve is open the other is closed to prevent fluid transfer between systems.


B737-300/-400/-500

Operation Hydraulic pressure from system B opens the system blocking valve and is applied to the retract side of the piston. The extend side of the piston is connected to the control valve through the open blocking valve. When the trailing edge flaps move between 0 and 1 unit, pressure is ported through the control valve to the extend side of the piston. Extend pressure working on a larger diameter piston overcomes retract pressure and drives the actuator full stroke. When the control valve is closed by the flaps returning to nearly full up, pressure from the extend side of the piston is ported to return. The leading edge flap will then retract due to hydraulic pressure on the retract side of the piston. Standby pressure, when selected, is applied to the spring side of the system blocking valve and closes it. The standby blocking valve is opened which ports standby pressure to the extend side of the piston applies and drives it full stroke to the extend position.



B737-300/-400/-500 27-80

STANDBY PRESSURE

SYSTEM B PRESSURE

EXTEND PRESSURE FROM CONTROL VALVE

FILTER STANDBY SYSTEM BLOCKING VALVE

SYSTEM B BLOCKING VALVE

VENT


SYSTEM B PRESSURE STANDBY PRESSURE EXTEND PRESSURE

Figu Figure re 115

Lead Leadin ing g Edge Edge Flap Flap Act Actua uato torr



27-80 LEADING EDGE AC TUATOR Purpose Hydraulic actuators convert available hydraulic power to mechanical energy that positions each leading edge slat to the desired position; retract, extend, or full extend. Location Each actuator is mounted between the wing front spar and the center of each slat. Physical Description/Features Each actuator housing contains two identical blocking valves, an internal mechanical locking mechanism, an outer piston, and an inner piston and rod assembly. The blocking valves function the same as that described for the leading edge flap actuator. External ports in the actuator housing are provided for a retract pressure line, two extend pressure lines, and a standby system pressure line. Power Hydraulic power is the same as that described for the leading edge flap actuator. In addition, a power transfer unit provides hydraulic pressure to the slat actuators when the right engine driven pump output pressure is low, trailing edge flaps are at positions 1, 2, 5, or 10, and the nose gear squat switch senses air.


B737-300/-400/-500

Control The blocking valves function the same as that described for the leading edge flap actuator. The slat actuator is maintained in the retract position by an internal mechanical lock. A spring-loaded locking piston holds lock segments between the inner piston and a locking stud. External pressure must be applied to compress a lock piston against spring force, which allows the lock segments to retract and unlock the actuator.

Operation System B hydraulic pressure opens the system blocking valve and is applied to the retract side of the piston. Ports C1 and C2 are open to return at the control valve, through the blocking valve. When trailing edge flaps move between 0 and 1 unit, the leading edge control valve is positioned to provide pressure at port C1. Pressure is directed to the lock piston and the inner piston. The actuator unlocks and the inner piston extends the slat to the intermediate position. When trailing edge flaps move between 5 and 10 units, pressure is directed from the control valve, through port C2, to the outer piston. It carries the inner piston as it extends the slat to full extend. Pressure from the power transfer unit, when actuated, is applied through the system B ports. During standby system operation, pressure is applied to the spring side of the system blocking valve to close it. The standby blocking valve is opened and pressure is applied simultaneously to both the inner and outer piston. The slat moves to full extend. Standby pressure cannot operate the slat to either the retract or extend positions. Monitor A reed switch inside the slat actuator sends a retract or not retracted signal to the indicating system. A magnet attached to the locking mechanism is held away from the switch when the actuator is locked retracted. When the actuator unlocks, the magnet is moved in proximity of the switch which changes the signal.



B737-300/-400/-500 27-80

STANDBY PRESSURE

SYSTEM B PRESSURE

SYSTEM B PRESSURE FROM CONTROL VALVE

C1 PORT STANDBY SYSTEM BLOCKING VALVE

C2 PORT

FILTER SYSTEM B BLOCKING VALVE

VENT

LOCKING STUD

LOCKING SEGMENTS LOCKING PISTON POSITION SWITCH


INNER PISTON SYSTEM B PRESSURE

PISTON SHOWN EXTENDING

INTERMEDIATE PRESSURE FULL EXTEND PRESSURE STANDBY PRESSURE

Figu Figure re 116

Lead Leadin ing g Edge Edge Slat Slat Act Actua uato torr

OUTER PISTON



27-80 LEADING EDGE FLAP MECHANISM Purpose A mechanism is provided to properly position each leading edge flap. Location The mechanism is connected between the underside of wing structure and the flap. Physical Description/Features Flaps 1 and 4 each have three gooseneck hinges and flaps 2 and 3 have five gooseneck hinges attached to fittings in the leading edge of the wing. A fitting is also provided on each flap to connect it to the hydraulic actuator. Two spring-loaded seal doors, one hinged to the flap and the other to wing structure, are installed on the outboard end of flaps 1 and 4. These doors are opened by the thrust reverser sleeve moving aft. A hinged fairing is installed on the trailing edge of flaps 1 and 4 and two fairings are installed on the trailing edge of flaps 2 and 3. A single linkage operates the fairing on flaps 1 and 4 and three linkages operate the two fairings on flaps 2 and 3. Each linkage consists of a link assembly between wing structure and a crank with a pushrod attached between the crank and the hinged fairing.


B737-300/-400/-500

Operation The actuator extends to drive the flap around the hinges. The hinged fairing is rotated into the airstream by the linkage. During retraction as the flap rotates around the hinges, the link assembly rotates the crank clockwise. This pulls the rod up and rotates the fairing counterclockwise to stow in the wing leading edge.



B737-300/-400/-500 27-80

LEADING EDGE FLAP ACTUATOR HINGE HINGE

SPRING

LEADING EDGE FLAP UPPER FLAP LINK ROD

SPRING

KRUEGER SEAL

LEVER SEAL INBD

LINK

CAM ROLLER

INBD

SEAL DOOR LOWER FLAP NOSE ROD FLAP NOSE


Figu Figure re 117

Lead Leadin ing g Edg Edge e Flap Flap Mech Mechan anis ism m



27-80 LEADING EDGE SLAT MECHANISM Purpose The leading edge slat mechanism is provided to position the slat as required. Location The slat and associated mechanism are located at the leading edge of each wing, outboard of the engine. Physical Description/Features Each slat is guided by two main tracks and two auxiliary tracks, which ride on rollers in the wing leading edge. A third auxiliary track is installed at the outboard end of slats 1 and 6. A three-position hydraulic actuator is attached at the center of each leading edge slat. A void between the slat inner and outer skins provide a path for thermal anti-icing. Anti-icing ducts installed in the wing leading edge connect with hot air supply lines through a telescoping duct.


B737-300/-400/-500



B737-300/-400/-500 27-80

UPSTOP (4 LOCATIONS EACH SLAT)

SLAT 1 SEE A DOWNSTOP

SLAT 2

AUXILIARY TRACK FOLLOWER BEARING

SLAT 3

LEADING EDGE SLAT ACTUATOR UPSTOP LEADING EDGE SLAT

SLAT 5

AUXILIARY SLAT TRACK ARM (INSTALLED THIS LOCATION ON SLATS 1 AND 6 ONLY)

ANTI-ICING DUCT DOOR

MAIN SLAT TRACK

MAIN SLAT TRACK


SLAT 4

UPSTOP

AUXILIARY SLAT TRACK (SLAT NO. 1 SHOWN)

OUTBD

FWD

UPSTOP

THERMAL ANTI-ICING TELESCOPE DUCT

A

Figur Figure e 118

AUXILIARY SLAT TRACK ARM

Lead Leading ing Edge Edge Slat Slat Mech Mechan anis ism m

SLAT 6



27-80 SLAT AUXILIARY TRACK Purpose Auxiliary tracks are installed to provide a means of stabilizing the slat at the intermediate position. Location Auxiliary tracks are mounted on the wing front spar, aft of the slat. Physical Description/Features An auxiliary track extension arm is attached to the slat. When the slat is moved, a roller on the extension arm rides in a cam track in the auxiliary track. Structural stops on the main track at the retracted and fully extended positions prevent the roller from bottoming out in the cam track. The slat is stabilized in the intermediate position by the roller contacting a detent arm, that is preloaded by a torsion rod attached to the front spar adjacent to the auxiliary track.


B737-300/-400/-500



B737-300/-400/-500 27-80

TORSION ROD ARM

TORSION ROD FRONT SPAR

TOP AFT AUXILIARY TRACK SUPPORT BOLT


TRUNNION TORSION ROD BOLT AFT AUXILIARY TRACK SUPPORT

DETENT ARM

AUXILIARY TRACK DOWNSTOP

Figur Figure e 119

Slat Slat Auxi Auxili liar ary y Tra Track ck and and Dete Detent nt Arm Arm



27-80 SLAT MAIN TRACK Purpose The main tracks attached to each slat act as guide units and contain the adjustable mechanical stops which limit full extend travel. Location The main tracks are attached to the stat and mounted in the wing leading edge aft of the slat. Physical Description/Features The main tracks are guided during slat movement by rollers mounted in the wing leading edge. A downstop mounted on the aft end of each main track engages an adjustable downstop on structure to limit slat extension. Upstops consist of four adjustable stop bolts in the fixed leading edge that contact stop fittings on the slat when retracted.


B737-300/-400/-500



B737-300/-400/-500 27-80

ROLLERS

FIXED LEADING EDGE MAIN TRACK COVER PLATE ADJUSTMENT BOLTS

FRONT SPAR

SLAT

DOWNSTOP (AIRPLANES WITHOUT THE DISTANCE FROM THE SLAT LOWER TRAILING EDGE TO THE FIXED LEADING EDGE PANEL

DISTANCE FROM THE MAIN TRACK COVER PLATE TO THE BOTTOM SURFACE OF THE LEADING EDGE


THREE-W AY ADJUSTABLE DOWNSTOPS) SEE

F

FRONT SPAR

SLAT SHOWN RETRACTED

F


Slat Slat Main Main Track rack

DOWNSTOP (AIRPLANES WITH THE THREE-W AY ADJUSTABLE DOWNSTOPS)



27-80 LEADING EDGE FLA PS AND SLAT OPERATION Operation/Control Sequence Normal operation of the leading edge devices is by System B. Retract pressure is direct to the actuators and both extend and full extend pressure is through a leading edge devices control valve. Control of the extend and retract sequence is by the trailing edge flap system. Alternate extension of the leading edge devices is by standby system pressure through the leading edge standby drive shutoff valve. Control is with the alternate flap master switch and the alternate flap control switch. Standby pressure can only position the flaps to extend and slats to full extend. Major/Subsystem Sequence An autoslat system is installed to automatically extend the slats from intermediate extend to full extend at high angles of attack. Autoslat control is by a dual channel autoslat control valve. A power transfer unit supplies hydraulic pressure for autoslat operation when the system B engine driven output pressure is low and the airplane is in the air with the trailing edge flaps at position 1 , 2, or 5.


B737-300/-400/-500

Normal Sequence System B pressure is delivered directly to the retract port of all leading edge device actuators. With the trailing edge flaps retracted, both extend ports are open to return through the leading edge devices control valve. As the trailing edge flaps move between 0 and 1 unit, the flap follow-up system positions the control valve to port pressure that drives the leading edge flaps to full extend and the slats to extend. Pressure is also supplied to both solenoid valves in the autoslat control valve. As the trailing edge flaps move between 5 and 10 units, the flap follow-up system positions the control valve to port pressure through the closed autoslat control valve to drive the slats to full extend. This sequence is reversed during the retraction cycle.

The slats are at intermediate extend and pressure is available at the autoslat control valve solenoids when the trailing edge flaps are at position 1, 2, or 5 units. When the airplane is off the ground and experiences excessive angle of attack, the autoslat computers signal the solenoids to open. Pressure is applied through to autoslat control valve to fully extend the slats. After the condition is corrected, the solenoid de-energizes and the valves close and the slat returns to the intermediate position. Power transfer unit pressure will accomplish the extension/retraction cycle as well as autoslat operation when the system B engine driven pump is not operating. Backup Operation Alternate extension of the leading edge devices is by standby pressure through the leading edge standby drive shutoff valve. During normal operation the shutoff valve is held closed by 28 volts dc. Positioning the alternate flaps master switch to ARM starts the standby pump and arms the control switch. When the control switch is moved to DOWN the leading edge shutoff valve relay is energized. Electric power opens the shutoff valve and hydraulic pressure is applied to the leading edge flap extend ports and slat full extend ports. A 1.5 gpm flow limiter and a hydraulic fuse rated at 280 cu. in. is installed in the standby extend line. The fuse closes and blocks hydraulic flow when volumetric capacity is exceeded. The fuse automatically resets at a delta pressure of 5 psi. The leading edge shutoff valve is held open by a holding circuit to the relay after the control switch is released from down. The master switch must be returned to OFF to close the shutoff valve.




B737-300/-400/-500 27-80


Lead Leading ing Edge Edge Flap Flap and and Sla Slatt Oper Operat atio ion n



27-81 AUTOSLAT COMPUTER (CONFIG 1) Power Autoslat computer 1 is powered by 28 volts ac from transfer bus 1 and 28 volts dc from bus 1. Autoslat computer 2 is powered by 28 volts ac fro; transfer bus 2 and 28 volts dc from bus 2. Computer 1 inputs are: - Left outboa outboard rd wheel wheel speed speed relay relay (M980) (M980) > 60 knots - Left alpha vane - angle angle of of attack attack - R321, nose nose gear gear ground ground sense relay (E-1 (E-11) 1) - R277, main main gear gear ground ground sense relay (E-1 (E-11) 1) - S814 and S815, S815, engine engine thrust thrust lever advance advance switche switches s - Autosl Autoslat at Com Comput puter er 2 - S740, S740, Trail Trailing ing Edge Edge Flap Flaps s 2" position switch. (not used) Computer 2 inputs are similar. Differences include the right outboard wheel speed relay, the right alpha vane, R344 and R343, nose and main gear ground sense relays, and autoslat computer 1. Outputs of the autoslat computers are to the respective solenoid valves in the autoslat control valve and to the autoslat fail light on the overhead panel, P5. Autoslat computer 1 with its associated inputs and outputs is designated channel 1 and autoslat computer 2 with its associated inputs and outputs is designated channel 2. Either channel independently can activate the autoslat system.


B737-300/-400/-500

Operation Each autoslat computer will signal its related solenoid to energize when excessive angle of attack is detected and that channel is not inhibited. Channels 1 and 2 are independently inhibited when: - the airplane airplane is is on the ground ground (main (main and and nose squat) squat) - the channe channell is inva invalid lid.. Angle of attack input from the respective alpha vane is compared to a trip point set in the computer. When the trip point is exceeded and the channel is not inhibited, an output signal to the autoslat control valve causes the slats to extend from intermediate to full extend.

Monitoring Each autoslat channel performs two separate monitoring routines on takeoff, static alpha compare and dynamic alpha compare, and one monitoring routine in flight. Static alpha compare begins when the airplane is on the ground, at least one thrust lever is advanced, and wheelspeed is above 60 knots. The vane angle is compared to 0 +3". If the vane is out of tolerance and the opposite vane is in tolerance, the channel is invalid, inhibited, and a signal is sent to the failure monitoring circuit. The test ends when either the nose or main gear squat switch indicates air mode. (This static alpha compare function is not available an all computers) Dynamic alpha compare begins when the main gear squat switch is in the air mode and wheelspeed is greater than 60 knots. The left and right vane signals are compared to be within 3 of each other. If disagreement exists with both systems valid, the higher and lower vanes signal the respective failure monitoring circuit, but both channels remain valid and operational. This test ends when the wheels are braked on gear retraction. In Flight Monitoring An automatic test of the autoslat system is normally conducted each flight. The test is initiated when the trailing edge flaps are moved to 15 units with the airplane in the air. Actuation of the landing warning switch, S138, on the flap control unit, in flight, causes each autoslat computer to energize its respective control valve coil. The slats are not affected since they are already at the full extend position. System failures will be reported to the autoslat indication system and latched.



B737-300/-400/-500 27-81

AUTOSLAT FAIL

P5-3



Auto Auto Slat Slat Syst System em Sch Schem emat atic ic (co (conf nfig. ig. 1)



27-81 AUTOSLAT CONTROL VALVE VALVE (CONFIG 2) Purpose The autoslat control valve provides the means of extending the slats from the intermediate to full extend position, to provide additional lift when high angle of attack is experienced. Location The autoslat control valve is located in the aft end of the right ram air duct bay. Physical Description/Features The autoslat control valve is divided into two sections, each containing two valves. One valve is solenoid operated and controlled by one of two stall management computers (SMC). Actuation of this solenoid valve opens ports which apply hydraulic pressure to operate the second hydraulic actuated valve. Actuation of the second valve ports hydraulic pressure from the slat extend line to the full extend line. Control Control of one section of the autoslat control valve is by a signal from stall management computer number 1 and the other section is by stall management computer number 2.


B737-300/-400/-500

Operation The autoslat control valve is normally closed and has no effect on normal operation of the leading edge slats. When the airplane is in the air, the stall management computers will actuate the solenoid valves if the airplane approaches a stall. The open solenoid valves direct pressure to open the pressure operated valves which direct intermediate extend pressure into the full extend line. If the slats are at the intermediate position, they will move to full extend and remain there until the autoslat computer signal ceases. The slats will then retract to the intermediate position.

BITE Operation of each set of valves within the autoslat control valve can be checked by test circuitry in the respective stall management computer. The solenoid valves can be actuated on the ground: - by positioning positioning the respect respective ive system system alpha vane up to a position position greater greater then 19.8. - In conjunction conjunction with with positioning positioning the alpha alpha vane, the GROUND GROUND sensing sensing test button on the E-11 must be depressed, - hydraulic hydraulic system system B pressure pressure availab available le - and leading leading edge devices devices and trailing trailing edge flaps flaps must be position positioned ed between 1 to 15 units.



B737-300/-400/-500 27-81 HYDRAULIC SYSTEM B MODULAR UNIT 29-00-02

B SYS RETURN MODULE 29-00-02

A

7

A RETURN MODULE 29-00-01

4

29-00-02 SYSTEM B RESERVOIR

GND SERV 29-00-02

R

POWER TRANSFER UNIT INTERMEDIATE EXTENSION

6

29-00-01 GND SERV

S

FULL EXTENSION

HYDRAULIC SYSTEM A MODULAR UNIT 29-00-01

HYD PUMP

B SYSTEM PRESSURIZED

PTU CONTROL VALVE (STA 678, LBL 12.9, WL 154)

HYD MOTOR

B SYSTEM DEPRESSURIZED

(MAIN GEAR WHEEL WELL ON KEEL BEAM)

LEADING EDGE DEVICES CONTROL VALVE

C2

C1

27-83-01 D3188 1

D3184B 26 H

2

24

C

VALVE OUT

M1222 AUTOSLAT AUTOSLAT COMPUTER COMPUTER 1 (E1-1)

A

INTERMEDIATE EXTENSION


27-83-02 FULL EXTENSION

D3190 1

D3186B 26 H

2

24

C

VALVE OUT

M1223 AUTOSLAT COMPUTER 2 (E1-1)

V132

1

AUTOSLAT CONTROL VALVE (STA 652, RBL 47)

1

ACTUATION OF EITHER AUTOSLAT VALVE WHEN LEADING EDGE IS FULL UP WILL PRODUCE NO S LAT MOTION ACTUATION OF EITHER AUTOSLAT VALVE WHEN LEADING EDGE IS MID UP WILL CAUSE SLATS TO GO FULL ACTUATION OF EITHER AUTOSLAT VALVE WHEN LEADING EDGE IS FULL EXT WILL PRODUCE NO FURTHER MOTION


Auto Auto Slat Slat Contro Controll Val Valve ve (co (confi nfig. g. 1) 1)



27-81 AUTOSLAT COMPUTER (CONFIG 1) Purpose Two autoslat computers are installed to actuate the autoslat control valve, when required. Location The autoslat computers are mounted on row 1 of the E1 electronic shelf in the electronic equipment compartment. Physical Description/Features Each autoslat computer contains the circuitry for operation of the autoslat control valve, failure monitoring, and self test. Light indicators and pushbuttons on the face of each computer are used for self test and failure monitoring.


B737-300/-400/-500



B737-300/-400/-500 27-81

VANE

1

ILLUMINATES IF VANE ASSOCIATED WITH THIS CHANNEL WAS NOT 130 JUST PRIOR TO LIFT OFF

COMPARE HIGH LOW

ILLUMINATES IF VANE ASSOCIATED WITH THIS CHANNEL WAS HIGHER AT LIFT OFF THAN OPPOSITE VANE BAC27DEX5478

AUTOSLAT COMPUTER NUMBER 2 SEE

A

AUTOSLAT COMPUTER NUMBER 1 SEE

SELF -MONITORILLUMINATES FOR COMPUTER FAILURE, VALVE COIL OPEN, OR SUCCESSFUL TEST

RESET

A

AUTOSLAT COMPUTER NUMBER 1

PWR OK

(NUMBER 2 IS IDENTICAL)

AUTO EX T

AUTOSLAT COMPUTER

P/N 65-52818-

A

H LEG TEST C LEG

S/N


1

COIL

A 1

A-A

NOT ON ALL AIRPLANES A


Auto Auto Sla Slatt Comp Compute uterr (co (confi nfig. g. 1)



27-81 AUTOSLAT COMPUTER SYSTEM (CONFIG 1) Power Autoslat computer 1 is powered by 28 volts ac from transfer bus 1 and 28 volts dc from bus 1. Autoslat computer 2 is powered by 28 volts ac fro; transfer bus 2 and 28 volts dc from bus 2. Computer 1 inputs are: - Left outboa outboard rd wheel wheel speed speed relay relay (M980) (M980) > 60 knots - Left alpha vane - angle angle of of attack attack - R321, nose nose gear gear ground ground sense relay (E-1 (E-11) 1) - R277, main main gear gear ground ground sense relay (E-1 (E-11) 1) - S814 and S815, S815, engine engine thrust thrust lever advance advance switche switches s - Autosl Autoslat at Com Comput puter er 2 - S740, S740, Trail Trailing ing Edge Edge Flap Flaps s 2" position switch. (not used) Computer 2 inputs are similar. Differences include the right outboard wheel speed relay, the right alpha vane, R344 and R343, nose and main gear ground sense relays, and autoslat computer 1. Outputs of the autoslat computers are to the respective solenoid valves in the autoslat control valve and to the autoslat fail light on the overhead panel, P5. Autoslat computer 1 with its associated inputs and outputs is designated channel 1 and autoslat computer 2 with its associated inputs and outputs is designated channel 2. Either channel independently can activate the autoslat system.


B737-300/-400/-500

Operation Each autoslat computer will signal its related solenoid to energize when excessive angle of attack is detected and that channel is not inhibited. Channels 1 and 2 are independently inhibited when: - the airplane airplane is is on the ground ground (main (main and and nose squat) squat) - the channe channell is inva invalid lid.. Angle of attack input from the respective alpha vane is compared to a trip point set in the computer. When the trip point is exceeded and the channel is not inhibited, an output signal to the autoslat control valve causes the slats to extend from intermediate to full extend.

Monitoring Each autoslat channel performs two separate monitoring routines on takeoff, static alpha compare and dynamic alpha compare, and one monitoring routine in flight. Static alpha compare begins when the airplane is on the ground, at least one thrust lever is advanced, and wheelspeed is above 60 knots. The vane angle is compared to 0 +3". If the vane is out of tolerance and the opposite vane is in tolerance, the channel is invalid, inhibited, and a signal is sent to the failure monitoring circuit. The test ends when either the nose or main gear squat switch indicates air mode. Dynamic alpha compare begins when the main gear squat switch is in the air mode and wheelspeed is greater than 60 knots. The left and right vane signals are compared to be within 3 of each other. If disagreement exists with both systems valid, the higher and lower vanes signal the respective failure monitoring circuit, but both channels remain valid and operational. This test ends when the wheels are braked on gear retraction. In Flight Monitoring An automatic test of the autoslat system is normally conducted each flight. The test is initiated when the trailing edge flaps are moved to 15 units with the airplane in the air. Actuation of the landing warning switch, S138, on the flap control unit, in flight, causes each autoslat computer to energize its respective control valve coil. The slats are not affected since they are already at the full extend position. System failures will be reported to the autoslat indication system and latched.



B737-300/-400/-500 27-81

LE FLAPS TRANSIT LIGHT

AIR

GND R375 NOSE SENSING

R282 MAIN GEAR AIR SENSING

M229 LE FLAPS AND SLATS INDICATION MODULE (E3-2)

COMPARE VALID IN ASC 1 POWER OK

LE FLAPS EXT LIGHT P2-2 PILOTS’ CENTER PANEL

COMPARE IN M1223 AUTOSLAT COMPUTER 2 GND

AUTOSLAT FAIL LIGHT

GND

R321 NOSE GEAR GND SE NSIN G

R277 MAIN GEAR GND S ENS ING

P5-3 FLIGHT CONTROL MODULE

E11 LANDING GEAR LOGIC SHELF

MASTER CAUTION LIGHT FLT CONT LIGHT P7 PILOTS’ LIGHTSHIELD

AUTOSLAT COMPUTER 2 EXTEND OUTPUT RETURN

L.O. >60KTS M980 AUTO SPEEDBRAKE MODULE E3-2 ELEC SHELF 28V DC BUS 1 AUTOSLAT 1 DC 28V AC XFER BUS AUTOSLAT 1 AC P6-2 CIRCU IT BREAKER PANEL


CX

T433 LEFT AOA SENSOR

FLAPS >10 S138 LANDING GEAR WARNING SW (R WHEEL WELL)

FULL EXT OUTPUT M1222 AUTOSLAT COMPUTER 1 E1-1 ELEC SHELF

NOTE: AUTOSLAT COMPUTER 1 SHOWN

V132 AUTOSLAT VALVE

INTERMEDIATE EXT OUTPUT LE DEVICES CONTROL VALVE (REF 27-81-00)

INT

FULL TO LE SLAT ACTUATORS (A12604)


Auto Auto Sla Slatt Sch Schem emat atic ic (con (config fig 1.) 1.)



27-81 AUTOSLAT COMPUTER FRONT PANEL (CONFIG 1) Maintenance Practices Five amber and one green light on the front of each autoslat computer assist fault isolation. - DS1 - illuminate illuminates s and latches latches if static alpha alpha compare compare fails. fails. - DS2 - illuminates and latches if vane was higher than 3 " from opposite vane during dynamic alpha compare. Opposite channel must have been valid when disagreement detected. - DS3 - illuminates and latches when high light is illuminated on opposite computer. - DS4 - illumi illuminates nates and and latches latches for computer computer failure, open valve coil, or successful test. - DS5 - illuminates illuminates when when power is good. Extingui Extinguishes shes when when ac, dc, or internal power is low. - DS6 - illuminates illuminates when when the autoslat autoslat computer computer is signalling signalling the solenoid solenoid to energize. Latched lights are cleared by depressing the reset pushbutton, S1, with the condition no longer present.


B737-300/-400/-500

BITE Two pushbuttons, S2 and S3, are provided to check the hot and cold leg signals to the solenoid valve coil. Faults must not be present to conduct this test. Press and hold S2 for one second. - D54 amber light illuminates - Clear D54 by pressing pressing S1 S1 reset reset switch. switch. - Repe Repeat at for for S3. S3. These pushbuttons can also be used to exercise the autoslat system. Press both S2 and S3 and hold - NO RESP RESPON ONSE SE Press “GRND sensing” test switch on front of E-11. - D56 illuminates illuminates and and slats move from from mid-extend mid-extend to full if positioned positioned at mid-extend and hydraulic power is available.



B737-300/-400/-500 27-81

DS1

DS 1

A

DS2 A

COMPARE

HIGH ILLUMINATES IF VANE ASSOCIATED LOW WITH THIS CHANNEL WAS HIGHER AT LIFT OFF THAN A OPPOSITE VANE

DS3

DS4

BAC27DEX5478 A

DS5

SELF -MONITORILLUMINATES FOR COMPUTER FAILURE, VALVE COIL OPEN, OR SUCCESSFUL TEST

DS6

G

A

PWR OK

AUTO EXT

RESET

S1

AUTOSLAT

S2

COMPUTER


P/N 65-52818S/N

H LEG TEST C LEG

S3

1 COIL


Autos Autosla latt Compute Computerr Front Front Pane Panell (confi (config g 1.)



27-81 AUTOSLAT FAILURE WARNING (CONFIG 1) Purpose Autoslat failure warning is provided to notify the crew when either or both autoslat channels are invalid or when disagreement exists during the dynamic alpha compare test. Location An amber autoslat fail light on the flight control module, P5-3, is controlled by relays inside the module. Physical Description/Features Signals from the autoslat computers operate two valid relays and one compare valid relay. Certain indications can only illuminate the fail light when the master caution relay is energized. Control The amber autoslat fail light illuminates, accompanied by master caution and the flight control annunciator when: - Both Both channe channels ls are are invali invalid. d. - One channel channel is invalid invalid and master master caution caution is recall recalled. ed. - Dynamic Dynamic alpha compare compare disagreement disagreement followed followed by master master caution recall. recall. Failures that generate an invalid signal are: - Comp Comput uter er Fail Failur ure e - Valve alve Coi Coill Open Open - Power low;-ac, low;-ac, dc, dc, or internal internal power. power. - Static Static Alpha Alpha Compa Compare re failur failure e


B737-300/-400/-500

Operation Grounds that energize the three valid relays are provided by the autoslat computers, when valid. This positions the respective relay contacts to open circuit the autoslat fail light. An invalid channel causes that computer to remove the relay ground and de-energize the respective relay. Autoslat fail light power has a direct path to ground only through· both channel valid relay contacts in the fail condition. One channel invalid will illuminate the light only when master caution recall is depressed. The dynamic alpha compare disagreement signal from both computers is interlocked to cause the compare relay to de-energize. Master caution recall must be depressed to illuminate the autoslat fail light for vane disagreement.



B737-300/-400/-500 27-81

FLT CONTROL A

STANDBY B

HYD LOW

STDBY RUD

STDBY RUD

OFF

OFF

A ON

B ON

QUANTITY

A

LOW PRESSURE

A

ALTERNATE FLAPS OFF LOW

UP

LOW

PRESSURE

A

PRESSURE

A

OFF SPOILER A B

ARM

OFF

OFF

ON

ON

DOWN

FEEL DIFF PRESS

A

SPEED TRIM FAIL

A

MACH TRIM

YAW DAMPER

FAIL

YAW DAMPER

A

AUTO SLAT FAIL

A

A

OFF ON

P5 OVERHEAD PANEL

FIRE


R

WARN BELL CUTOUT

MASTER

A

CAUTION PUSH TO RESET

FLT CONT

ELEC

IRS

FUEL

APU

OVH T/DE T

P7 PANEL, LEFT GLARESHIELD


Autos Autosla latt Failu Failure re War Warnin ning g (confi (config. g. 1) 1)



27-80 AUTOSLAT SYSTEM (CONFIG. 2) Purpose Two stall management computers are installed, to actuate the autoslat control valve, when required. Each stall management computer combines the function of an autoslat channel and stall warning system.

Computer 2 inputs are similar. Differences include the right outboard wheel speed relay, the right alpha vane, R344 and R343, nose and main gear ground sense relays, and stall management computer 1 .

SMC No. 1 controls autoslat channel No. 1 and the Captain’s stall warning system, SMC No. 2 controls autoslat channel No. 2 and the First Officer’s stall warning system.

Outputs of the stall management computers are to the respective solenoid valves in the autoslat control valve and to the autoslat fail light on the overhead panel, P5.

Location The stall management computers are mounted on row 1 of the E1 electronic shelf in the electronic equipment compartment.

Stall management computer 1 with its associated inputs and outputs is designated channel 1 and stall management computer 2 with its associated inputs and outputs is designated channel 2. Either channel independently can activate the autoslat system.

Physical Description / Features Each stall management computer contains the circuitry for operation of the autoslat control valve, failure monitoring, and self test. An amber alphanumeric display and (6) pushbuttons on the face of each computer are used for conducting self tests, provide test/failure indications and conduct BITE (built-in-test equipment). Power supply Stall management computer 1 is powered by 28 volts ac from transfer bus 1 and 28 volts dc from bus 1. Stall management computer 2 is powered by 28 volts ac from transfer bus 2 and 28 volts dc from bus 2. Computer 1 inputs are: y l n O s e s o p r u P g n i n i a r T r o F

B737-300/-400/-500

- Left outboa outboard rd wheel wheel speed speed relay relay (M980) (M980) > 60 knots -

Left alpha vane - angle angle of of attack attack Left Traili Trailing ng Edge Flap Flap position position transm transmitter itter R321, nose nose gear gear ground ground sense relay (E-1 (E-11) 1) R277, main main gear gear ground ground sense relay (E-1 (E-11) 1) S814 and S815, S815, engine engine thrust thrust lever advance advance switche switches s Stall Stall Manage Managemen mentt Comput Computer er 2

Operation Each stall management computer will signal its related solenoid to energize when excessive angle of attack is detected and that channel is not inhibited. Channels 1 and 2 are independently inhibited when: - the airplane airplane is is on the ground ground (main (main and nose nose squat) squat) - the chan channel nel is inva invalid lid.. Angle of attack input from the respective alpha vane is compared to a trip point set in the computer. When the trip point is exceeded and the channel is not inhibited, an output signal to the autoslat control valve causes the slats to extend from intermediate to full extend. In Flight Monitoring An automatic test of the autoslat system is normally conducted each flight. The test is initiated when the trailing edge flaps are moved to 15 units with the airplane in the air. Actuation of the landing warning switch, S138, on the flap control unit, in flight, causes each stall management computer to energize its respective control valve coil. The slats are not affected since they are already at the full extend position. System failures will be reported to the autoslat indication system and latched.




B737-300/-400/-500 27-80


Autos Autosla latt Syste System m Sche Schema matic tic (con (config fig.. 2)



27-80 AUTOSLAT CONTROL VALVE VALVE (CONFIG 2) Purpose The autoslat control valve provides the means of extending the slats from the intermediate to full extend position, to provide additional lift when high angle of attack is experienced. Location The autoslat control valve is located in the aft end of the right ram air duct bay. Physical Description/Features The autoslat control valve is divided into two sections, each containing two valves. One valve is solenoid operated and controlled by one of two stall management computers (SMC). Actuation of this solenoid valve opens ports which apply hydraulic pressure to operate the second hydraulic actuated valve. Actuation of the second valve ports hydraulic pressure from the slat extend line to the full extend line. Control Control of one section of the autoslat control valve is by a signal from stall management computer number 1 and the other section is by stall management computer number 2.


B737-300/-400/-500

Operation The autoslat control valve is normally closed and has no effect on normal operation of the leading edge slats. When the airplane is in the air, the stall management computers will actuate the solenoid valves if the airplane approaches a stall. The open solenoid valves direct pressure to open the pressure operated valves which direct intermediate extend pressure into the full extend line. If the slats are at the intermediate position, they will move to full extend and remain there until the autoslat computer signal ceases. The slats will then retract to the intermediate position.

BITE Operation of each set of valves within the autoslat control valve can be checked by test circuitry in the respective stall management computer. The solenoid valves can be actuated on the ground: - by positioning positioning the respect respective ive system system alpha vane up to a position position greater greater then 19.8. - In conjunction conjunction with with positioning positioning the alpha alpha vane, the GROUND GROUND sensing sensing test button on the E-11 must be depressed, - hydraulic hydraulic system system B pressure pressure availab available le - and leading leading edge devices devices and trailing trailing edge flaps flaps must be position positioned ed between 1 to 15 units.



B737-300/-400/-500 27-80

HYDRAULIC SYSTEM B MODULAR UNIT 29-00-02

B SYS RETURN MODULE 29-00-02

A

7

A RETURN MODULE 29-00-01

4

29-00-02 SYSTEM B RESERVOIR

GND SERV 29-00-02

R

POWER TRANSFER UNIT FULL EXTENSION

6

29-00-01 GND SERV

S


HYDRAULIC SYSTEM A MODULAR UNIT 29-00-01

HYD PUMP

B SYSTEM PRESSURIZED

PTU CONTROL VALVE (STA 678, LBL 12.9, WL 154)

HYD MOTOR

B SYSTEM DEPRESSURIZED

(MAIN GEAR WHEEL WELL ON KEEL BEAM)

LEADING EDGE DEVICES CONTROL VALVE

C2

C1

27-83-01 D3188 1

D3184B 26 H

2

24

C

VALVE OUT

STALL MANAGEMENT COMPUTER 1 (E1-1) A



27-83-02 D3190 1

FULL EXTENSION

2

D3186B 26 H 24

C

VALVE OUT

STALL MANAGEMENT COMPUTER 2 (E1-1)

V132

1

AUTOSLAT CONTROL VALVE (STA 652, RBL 47)

1

ACTUATION OF EITHER AUTOSLAT VALVE WHEN LEADING EDGE IS FULL UP WILL PRODUCE NO SLAT MOTION ACTUATION OF EITHER AUTOSLAT VALVE WHEN LEADING EDGE IS MID UP WILL CAUSE SLATS TO GO FULL ACTUATION OF EITHER AUTOSLAT VALVE WHEN LEADING EDGE IS FULL EXT WILL PRODUCE NO FURTHER MOTION


Auto Auto Slat Slat Contro Controll Val Valve ve (co (confi nfig. g. 2) 2)



27-80 STALL MANAGEMENT COMPUTER (SMC) (CONFIG 2) Purpose Two stall management computers are installed, to actuate the autoslat control valve, when required. Each stall management computer combines the function of an autoslat channel and stall warning system. SMC No. 1 controls autoslat channel No. 1 and the Captain’s stall warning system, SMC No. 2 controls autoslat channel No. 2 and the First Officer’s stall warning system. Location The stall management computers are mounted on row 1 of the E1 electronic shelf in the electronic equipment compartment. Physical Description/Features Each stall management computer contains the circuitry for operation of the autoslat control valve, failure monitoring, and self test. An amber alphanumeric display and (6) pushbuttons on the face of each computer are used for conducting self tests, provide test/failure indications and conduct BITE (built-in-test equipment). Character Display - indicates the respective menu menu steps (together with question mark) or the respective lists within a menu step. BITE Test / ON/OFF Button - activates activates or deactivate deactivates s the BITE function function and the push push buttons buttons of the SMC.


B737-300/-400/-500

MENU Button - indicates indicates on the display display the the chosen step step out of the menu menu at any time of the BITE. Pushing the button terminates a running system test and returns to the next higher menu step. UP Scroll Button - contro controls ls the the menu or or listin listing. g. while the button is kept depressed, the listing is indicated in a sequence (four list points / a second). At the beginning of the list the display shows ”BEGIN OF LIST”.

DOWN Scroll Button - contro controls ls the menu menu or or listing listing.. while the button is kept depressed, the listing is indicated in a sequence (four list points / a second). At the end of the list the display shows ”END OF LIST”. YES Button - confirms confirms the the respective respective menu option. option. NO Button - confirms confirms the the respective respective menu option. option. while the button is kept depressed, the listing is indicated in a sequence (four list points / a second). At the end of the list the display shows ”END OF LIST”. If the YES or NO Buttons are pressed when it is not necessary within a system test (e.g. menu or listing), it will be displayed ”BUTTON INACTIVE”



B737-300/-400/-500 27-80

BITE INSTRUCTIONS

SMC MAINTENANCE CONTROL PANEL MAIN EQUIPMENT CENTER SEE

DISPLAY

A

MENU ANGLE OF AIRFLOW SENSOR YES

ON/ OFF NO

STALL MANAGEMENT COMPUTER SEE

B STALL MANAGEMENT COMPUTER


PART NO. 65-52822

SERIAL NO. BOEING A

B

FWD


STAL STALL L MANAG MANAGEM EMENT ENT COM COMPUT PUTER ER (CON (CONFI FIG. G. 2) 2)



27-80 AUTOSLAT CHANNEL OF SMC (CONFIG 2) Power Stall management computer 1 is powered by 115 volts ac and 28 volts dc from standby bus 1, and 28 volts dc from main bus 1. Stall management computer 2 is powered by 115 volts ac and 28 volts dc from electronic bus 2, and 28 volts dc from main bus 2. Input/Output Computer 1 inputs are: - R321, nose nose gear gear ground ground sense relay relay (E-1 (E-11) 1) - R277, main main gear gear ground ground sense relay relay (E-1 (E-11) 1) - Stall Stall manage managemen mentt Comput Computer er 2 - Left alpha alpha vane vane (input (input to stall stall warning warning channel) channel) Computer 2 inputs are similar. Differences include R344 and R343, nose and main gear ground sense relays, stall management Computer 1, and right alpha vane. Outputs from the autoslat channels of the SMC’s are to the respective solenoid valves in the autoslat control valve and to the autoslat fail light on the overhead panel, P5. SMC 1 with its associated inputs and outputs is designated channel 1 and SMC 2 with its associated inputs and outputs is designated channel 2. Either channel independently can activate the autoslat system.


B737-300/-400/-500

Operation Each autoslat channel of the SMC will signal its respective solenoid to energize when excessive angle of attack is detected and that channel is not inhibited. Channels 1 and 2 are independently inhibited when: - the airplane airplane is is on the ground ground (main (main and and nose squat) squat) - the channe channell is Inva Invalid lid.. Angle of attack input from the respective alpha vane is compared to a trip point set in the Computer. When the trip point is exceeded and the channel is not inhibited, an output signal to the autoslat control valve causes the slats to extend from intermediate to full extend.

Monitoring The autoslat channel, of the respective stall management computer, performs a constant monitoring routine. Each autoslat channel performs an automatic system test each landing. An automatic test of the autoslat system is normally conducted each landing. The test is initiated when the trailing edge flaps are moved to 15 units with the airplane in the air. Actuation of the landing warning switch, S138, on the flap control unit, in flight, causes each autoslat channel to energize its respective control valve coil. The slats are not affected since they are already at the full extend position. System failures will be transmitted to and stored in the fault monitoring section of the stall management computer. The faults will then be displayed, in the alphanumeric window on front of the computer, when the bite test is initiated.



B737-300/-400/-500 27-80

LE FLAPS 115V AC

TRANSIT

STBY BUS

LIGHT SMC 1 SNSR

T352 AUTO

EXC AC

LE FLAPS

TRANSFORMER

28V DC STBY BUS

M229 LE FLAPS AND SLATS INDICATION MODULE

SMC 1 CMPTR DC

EXT LIGHT P2-2 PILOTS’ CENTER PANEL

P18-2 CIRCUIT BREAKER PANEL 28V DC

MASTER CAUTION

AUTOSLAT FAIL

BUS 1 AUTOSLAT

LIGHT

1 DC

LIGHT FLT CONT LIGHT

P5-3 FLIGHT CONTROL MODULE


P7 PILOTS’ LIGHTSHIELD SMC 2 AUTOSLAT EXTEND OUTPUT

CX RETURN T433 LEFT AOA SENSOR

CX

T428 RIGHT FLAP POSITION XMTR


GND

AIR R276 AIR SENSING

R277 GND SENSING FULL EXT OUTPUT

GND R375 NOSE SENSING E11 LANDING GEAR LOGIC SHELF

M1506 STALL ____ NOTE: SMC 1 SHOWN SMC 2 SIMILAR

MANAGEMENT COMPUTER 1 E1-1 ELEC SHELF


V132 AUTOSLAT VALVE

INTERMEDIATE EXT OUTPUT LE DEVICES CONTROL VALVE

Autos Autosla latt Channe Channell of of SMC SMC (conf (config. ig. 2)

INT

FULL TO LE SLAT ACTUATORS



27-80 AUTOSLAT FAILURE WARNING Purpose Autoslat failure warning is provided to notify the crew when either or both autoslat channels are invalid or when disagreement exists during each static alpha vane test. Location An amber AUTOSLAT FAIL LIGHT on the flight control module, P5-3, is controlled by relays inside the module. Physical Description/Features Signals from the stall management Computers operate two valid relays and one alpha vane valid relay. Certain indications can only illuminate the fail light when the master caution recall relay is energized.


B737-300/-400/-500

Control The amber AUTOSLAT FALL LIGHT illuminates, accompanied by both master caution lights and the FLT CONT annunciator when: - Both channels channels are invalid invalid (light illuminat illuminates es automatical automatically). ly). or - One channel channel is invalid invalid accompanied accompanied by master master caution caution recall. recall. or - Static Static alpha vane(s) vane(s) inop accompan accompanied ied by master caution caution recall. recall. Failures that generate an invalid signal are: - Comp Comput uter er Fail Failur ure e - Valve alve Coi Coill Open Open - Power low;-ac, low;-ac, dc, dc, or internal internal power. power.

Operation Grounds that energize the three valid relays are provided by the stall management Computers, when valid. This positions the respective relay contacts to open circuit the autoslat fail light. An invalid channel causes that Computer to remove the relay ground and de-energize the respective relay. AUTOSLAT FAIL LIGHT power has a direct path to ground only through both channel valid relay contacts in the fail condition. One channel invalid will illuminate the light only when master caution recall is depressed. The static alpha vane(s) inop signal from each vane’s respective computer is interlocked to cause the relay to de-energize. Master caution recall must be depressed to illuminate the AUTOSLAT FAIL LIGHT for vane(s) failure.



B737-300/-400/-500 27-80

FLT CONTROL A

STANDBY B

HYD LOW

STDBY RUD

STDBY RUD

QUANTITY

A

LOW PRESSURE

OFF

OFF

A ON

B ON

A

ALTERNATE FLAPS OFF LOW

UP

LOW

PRESSURE

PRESSURE

A

A

OFF SPOILER A B

DOWN

ARM

OFF

OFF

FEEL DIFF PRESS

ON

ON

A

SPEED TRIM FAIL

A

MACH TRIM

YAW DAMPER

FAIL

YAW DAMPER

A

AUTO SLAT FAIL

A

A

OFF ON

P5 OVERHEAD PANEL

FIRE


R

WARN BELL CUTOUT

MASTER

A

CAUTION PUSH TO RESET

FLT CONT

IRS

FUEL

ELEC

APU

OVH T/DE T

P7 PANEL, LEFT GLARESHIELD


Autos Autosla latt Warni arning ng (confi (config. g. 2)


FLIGHT CONTROL LIFT AUGMENTING

27-80 LEADING EDGE DEVICE INDICATION Purpose The leading edge device indication system provides visual indication of the position of each leading edge flap and slat. Location Individual amber and green lights for each device are located on the annunciator module in the aft overhead panel. Two master annunciator lights, one amber and one green, are located on the captain’s instrument panel. The leading edge flap and slat indicating module, M229, is mounted on Row 2 of the E3 rack in the electronic equipment compartment. Physical Description/Feature The leading edge flap and slat indicating module contains the switching and logic cards that control the lights. The annunciator panel on P5 contains one amber transit light and one green extend light for each leading edge flap and slat, as well as one green full extend light for each slat. These lights are controlled by the position switches for each individual device. Only one light can be illuminated for each device at a given time. The master annunciator lights on the captain’s panel are controlled by all of the leading edge device position sensors and two trailing edge flap position switches on the flap control unit or by either autoslat computer. The circuit is biased so that both lights cannot illuminate at the same time. Power The M229 module is powered by 28 volts dc from bus 1. The lights are powered by 28 volts dc from master dim. All of the lights are dimmable.


B737-300/-400/-500

Operation All lights are extinguished when the leading edge flaps and slats are retracted. The individual amber lights on the annunciator panel illuminate as the respective device leaves the retract position. The master amber light, LE FLAPS TRANSIT, illuminates when the first device leaves retract. As each device reaches the extend position, the individual amber light extinguishes and the green light illuminates. The master annunciator lights on the captains panel switch from amber to green when all devices are at the extend position and the trailing edge flap switches indicate not up and not #10 units.

When the trailing edge flaps move between 5 and 10 units, the slats leave the extend position which illuminates the individual slat amber lights. The first slat in transit illuminates the master amber light. The individual amber lights extinguish and green lights illuminate as each slat reaches full extend. The master amber light remains illuminated until all slats are at full extend and the trailing edge flaps switches indicate 10 units or greater and not up. Then the amber light extinguishes and the green, LE FLAPS EXT light illuminates. The master amber light on the captain’s panel is illuminated for disagreement between trailing edge flap position and the position of any leading edge flap or slat. The lights function in reverse on retraction. Monitor Autoslat operation of the slats from extend to full extend and back can only be monitored on the overhead annunciator in flight. Either autoslat computer signals flap and slat comparator logic to hold off the master amber light and hold on the master green light when it is commanding the slats from extend to full extend. This signal remains until 13 seconds after the autoslat command ceases which is sufficient time for the slats to retract to the intermediate position. The circuit from the autoslat computers is wired through both main gear and nose gear air/ground sensing relay controls. The contacts are open on the ground, allowing the lights on the captain’s panel to function normally.



B737-300/-400/-500 27-80

LE DEVICES ANNUNCIATOR PANEL (Aft Overhead Panel)

POSITION INDICATOR

LE DEVICES FLAPS

TRANSIT

1

A

A

EXT FULL EXT

G G 1

A

G

A

G

TRANSIT

4

2

3

A

A

A

A

G 3

G

G

G

G G

G

G

G

5

4

SLATS

LE FLAPS

TRANSIT

EXTEND

EXT

G

G 2

A

A

LE FLAPS

G 6

FULL

(below Flap Position Indicator

EXT

at Pilot’s Center Panel)

SLATS

TEST

FLAPS 10-DEGREE SWITCH, S584

UNLOCK

LOCK SWITCH (ACTUATOR)

FLAP CONTROL UNIT FOLOW UP

EXTEND

> < 10_ TE FLAPS

SLAT EXTEND PROX. SENSOR

FULL EXTEND

SLAT FULL EXTEND PROX. SENSOR

STALL WARNING SWITCH, S856

TE FLAPS UP/ NOT UP LE SLAT SENSORS


INBD

1

(INHIBIT IN AIR)

RETRACT

LE FLAP RETRACT PROX. SENSOR

FLAP / SLAT POSITION INDICATION MODULE M 229 (E 3-2)

EXTEND

LE FLAP EXTEND PROX. SENSOR LE FLAP SENSORS

AUTOSLAT OR STALL MANAGMENT COMPUTER NO.1 OR 2

1

WITH AUTOSLAT SIGNAL IN AIR CENTER PANEL LIGHTS WILL NOT CHANGE FROM GREEN TO AMBER.


Lead Leadin ing g Edge Edge Devi Device ce Indic Indicati ating ng



27-80 LEADING EDGE FLAP POSITION SENSORS Purpose Leading edge flap position sensors provide a signal to operate flap position indicators. Location Two proximity sensors, retract and extend, are mounted on wing leading edge structure, above each flap. A retract actuator is mounted on one of the moveable hinges of each flap. An extend actuator is mounted on one of the link assemblies that operate each flap nose fairing. Physical Description/Features Each proximity sensor operates a switch in the leading edge flap and slat indicating module, M229. The sensors are controlled by the proximity or lack thereof to individual actuators. Proximity of sensor and actuator occurs when the flap is in that position, extend or retract. Lack of proximity indicates the flap is not in that position. The three possible flap positions are retract, extend, or in transit.


B737-300/-400/-500



B737-300/-400/-500 27-80

KRUEGER FLAP EXTEND SENSOR KRUEGER FLAP RETRACT SENSOR

ACTUATION BAR (FLAP RETRACT)


KRUEGER FLAP (SHOWN EXTENDED)


Leas Leasin ing g Edg Edge e Posi Positio tion n Sens Sensor ors s

ACTUATION BAR (FLAP EXTEND)

UPPER FLAP NOSE ROD



27-80 SLAT POSITION SENSORS Purpose Slat position sensors report slat position to the leading edge device indicating system. Location A reed switch for indication of retract position is located in the rod end of each hydraulic actuator. (config. 1) Since the read switch within the actuator was not reliable Boeing modified the position. A new reed switch is installed on the leading edge and the magnet is installed on the slat. (config 2) Two extend proximity sensors are mounted on leading edge structure adjacent to the forward end of one of the auxiliary tracks for each slat. A single actuator is attached to the extension arm roller attach bolt. Physical Description/Feature The retract switch was previously described under the slat actuator. Each extend sensor operates a switch in the leading edge flap and slat indicating module M229. They are proximity devices that operate exactly as described for the leading edge flap sensors. Four possible slat positions are retract, intermediate extend, full extend, or in transit.


B737-300/-400/-500

Operation The reed switch inside the hydraulic actuator monitors the slat in and out of the retract position. When the slat moves to the intermediate position, the auxiliary track roller moves the actuator into proximity of the aft sensor. When the slat moves to the full extend position, the roller carries the actuator away from the aft sensor and into proximity with the forward mounted sensor.



B737-300/-400/-500 27-80

PROXIMITY SENSOR SEE G

AUXILIARY SLAT TRACK

LOCKNUT F MAGNET SEAT

FWD

SLEEVE

FULL EXTEND (GREEN LIGHT) PROXIMITY SENSOR

FWD

MAGNET SPRING

SLEEVE

ROD END


LOCK INDICATOR ASSEMBLY (SWITCH)

EXTEND (GREEN LIGHT) PROXIMITY SENSOR SLAT RETRACTED

SWITCH LEAD KEY

INNER SLEEVE OUTER SLEEVE GROUND WIRE

CONNECTOR PLUG

NOTE: SLAT ACTUATOR LOCK SWITCH (CONFIG 1) IS LOCATED WITHIN THE ROD.ASSEMBLY OF THE SLAT ACTUATOR REED SWITCH (config 2)


FWD MAGNET

Slat Slat Posi Positi tion on Sens Sensor ors s



27-80 THE STALL WARNING SWITCH (S 856) Purpose The switch transmits the trailing edge flap position, trailing edge position UP or NOT UP, to the Flap / Slat Indication Module M229 to control the amber transit light. Location The switch is located at the Flap Control Unit in the main wheel well compartment on the flap control unit follow up drum.

THE FLAP 10 SWITCH (S 584) Purpose The switch transmits the trailing edge flap position, more or less than 10  extended, to the Flap / Slat Indication Module M229 to control the amber transit transit light. Location The switch is located at the Flap Control Unit in the main wheel well compartment on the flap control unit follow up drum.


B737-300/-400/-500



B737-300/-400/-500 27-80

FLAPS 10-DEGREE SWITCH, S584

FLAP CONTROL UNIT

SEE

A

RETAINING NUTS

SWITCH COVER SEE B ROLLER GUIDE (EXAMPLE)


INBD

FWD

INBD STALL WARNING SWITCH, S856

A

Figure Figure 136

Trail Trailing ing Edge Control Control Unit Switch Switch Locati Location on

B



27-80 LE FLAP / SLAT POS. INDIC. MODULE (CONFIG 1) Purpose The Module (M229) controls the function of the Leading Edge Flap/Slat Indication System. Location The Module (M229) is installed in the E + E Compartment (E 3-2). Physical Description The Module M229 receives signals from the following sensors / switches: - 2 Se Senso nsors rs LE Flap Flap Positi Position on - 3 Sen Sensor sors s LE Slat Slat Posi Positio tion n - 2 TE Flap Positio Position n Switches Switches at the Flap Flap Control Control Unit - Autoslat Autoslat signal signal (A/S Computer Computer or Stall Management Management Computer Computer)) via AIR/ GRD Relays from Nose and Main Gear.


B737-300/-400/-500



B737-300/-400/-500 27-80 LE DEVICES ANNUNCIATOR LE FLAP AND

PANEL

SLAT POSITION INDICATOR

LE FLAP AND SLAT POSITION INDICATION MODULE SEE


A

Figure Figure 137

LE Flap/S Flap/Slat lat Pos.. Pos.. Ind. Module Module M229 M229 (confi (config g 1)

A



27-80 LE FLAP / SLAT POS. INDIC. MODULE (CONFIG 2) Purpose The LE Flap / Slat Position Indication Module (M 229) - FSIM controls the function of the Leading Edge Flap/Slat Indication System. Location The LE Flap / Slat Position Indication Module (M 229) - FSIM -is located in the E + E Compartment (E 3-2). Physical Description The LE Flap / Slat Position Indication Module (M 229) - FSIM receives signals from the following sensors/switches: -

2 Se Senso nsors rs LE Flap Flap Positi Position on 3 Sen Sensor sors s LE Slat Slat Posi Positio tion n 2 TE Flap Positio Position n Switches Switches at the Flap Flap Control Control Unit Autoslat Autoslat signal signal (A/S Computer Computer or Stall Management Management Computer Computer)) via AIR/ GRD Relays from Nose and Main Gear.

BITE The LE Flap / Slat Position Indication Module (M 229) - FSIM contains the following BITE features: - non volatil volatile e memory memory of the past past faults faults - a readout readout of the gaps gaps between between the sensor sensors s and targets. targets.


B737-300/-400/-500



B737-300/-400/-500 27-80

LE FLAP AND SLAT POSITION INDICATOR

LE DEVICES ANNUNCIATOR PANEL

LE FLAP/SLAT INDICATION MODULE BOEING P/N 85-52807 S/N MOD. LEVEL: A B C D E F G H I J K OTHER FUNCTIONS ? YES PROX SENSORS ? YES AND SCROLL MENU

YES

LE FLAP AND SLAT POSITION INDICATION MODULE

ON OFF

NO

SEE


A

Figure Figure 138

LE Flap Flap / Slat Slat Pos. Pos. Ind. Ind. Modul Module e M229 M229 (config (config 2)

A


INDICATION / RECORDING SYSTEMS INDEPENDENT INSTRUMENTS

31-20

31-20

INDEPENDENT INSTRUMENTS

TAKEOFF WARNING SYSTEM Purpose

The takeoff warning System is installed to provide an aural warning to the pilot when takeoff is attempted with certain flight controls not in the proper position.

System Description An intermittent aural warning will sound in the flight compartment when the airplane is on the ground, either throttle is advanced, and the speed brake lever, horizontal stabilizer, trailing edge flaps, or leading edge flaps are not in the takeoff position.


B737-300/-400/-500




B737-300/-400/-500 31-20


Takeo akeoff ff War Warni ning ng Sys Syste tem m



31-20 TAKEOFF WARNING Purpose The takeoff warning system warns the pilot if L.E. or T.E. flaps, stabilizer, stabilizer, and speedbrake handle are not in their respective takeoff position or the parking brake is set prior to takeoff. General Component Locations The speed brake takeoff warning switch is mounted inside the pilots control stand, directly beneath the speed brake lever down detent. The trailing edge flap takeoff warning switch is mounted on the flap control unit. Two stabilizer takeoff warning switches are mounted on the aft wall of the stabilizer access compartment, left of the jackscrew. Relays energized by the extend proximity switches of numbers t and 4 leading edge flaps are on the back of the E-3 rack, row 2. Thrust lever advance switches are inside the control stand, adjacent to the thrust lever cable drum. Air-ground sensing and control circuits are in the landing gear logic shelf in the lower nose compartment. The parking brake lever switch closes the park and squat relay in the E11 shelf with airplane on ground. Contacts in the park and squat relay close to provide ground to takeoff warning circuit. The warning horn is located forward of the control stand on the first officer‘s side.


B737-300/-400/-500

General Subsystem Features The proper configuration of the flight controls for takeoff are: - Speed Brake Lever - Down Down detent detent - Stabil Stabilize izerr - Green Green Ran Range ge - Trailing Trailing Edge Edge Flaps Flaps - 1 unit throug through h 15 units. units. - Leadin Leading g Edge Edge Flaps Flaps - Exten Extend d - Parkin Parking g Brake Brake Releas Released ed General Operation If either thrust lever is advanced to the takeoff thrust range, an intermittent warning horn in the control cabin will sound if either the L.E. or T.E. flaps or stabilizer are not in the takeoff range, parking brake is not released, or the speedbrake handle is not stowed. The horn can only be silenced by correcting the configuration or retarding the thrust lever.




B737-300/-400/-500 31-20


Takeo akeoff ff War Warni ning ng Comp Compone onent nt Loca Locatio tion n




B737-300/-400/-500 31-20

TAKEOFF WARNING SCHEMATIC Operation The flight controls warning switches are parallel-wired. In series with these switches are an air/ground sensing circuit. The air/ground safety sensor deactivates the takeoff warning circuit when the airplane is airborne (weight off the landing gear). The air/ground sensing and control circuits are in the landing gear logic unit module. Flap Takeoff Warning Switch:  The flap takeoff warning switch is installed on the flap control valve follow-up mechanism. It is actuated by a cam on the control valve follow-up drum shaft when the flaps are outside the takeoff range. Leading edge flap warn relays:  The leading edge flap warn relays are installed on the E3-2 electrical shelf. One relay is used with leading edge flap No. 1 and the other relay is used with leading edge flap No. 4.  The relays are controlled by circuits in the leading edge flaps and slats position indication module located on the E3-2 shelf.  The relays are de-energized when the leading edge flaps are outside the takeoff range as controlled by proximity switches in the leading edge flap assemblies. Stabilizer Takeoff Warning Switches:  The two stabilizer takeoff warning switches are installed in the stabilizer actuator compartment and are actuated by cams on the stabilizer center section leading edge structure.  The upper switch actuates when the stabilizer is outside the takeoff (green band) range in the leading edge up (APL NOSE DN) direction.  The lower switch actuates when the stabilizer is outside the takeoff (green band) range in the leading edge down (APL NOSE UP) direction. Speed Brake Takeoff Warning Switch:  The speed brake takeoff warning switch is installed on the upper forward portion of the control stand.  The switch is actuated when the speed brake control lever is lifted out of the down and locked detent.

Park and Squad Relay (R 274):  The park and squad relay is energized when the parking brake is set  The park and squad relay is installed in the landing gear logic unit module.  The relay (R274) is controlled by the parking brake switch S100 which is actuated by the parking brake linkage.



B737-300/-400/-500 31-20

SPEED BRAKE SW

GRD = 0

(CLOSED WHEN NOT DOWN)

28V DC GATE T.E. FLAP TAKEOFF WARN SW (CLOSED WHEN FLAPS NOT IN TAKEOFF RANGE)

AURAL WARN L.G. SAFETY SW

LANDING GEAR

GND SENSORS

LOGIC SHELF

BATT BUS

ADV UPR STAB TRIM SW (CLOSED WHEN STAB

ENG 1 T/O

NOT IN GREEN BAND)

WARN SW

VOLTAGE REG

17V DC

28V DC TO PWR AMPL

RET ADV ENG 2 T/O LWR STAB TRIM SW

WARN SW

(CLOSED WHEN STAB RET

NOT IN GREEN BAND)

PILOT’S CONTROL STAND


R 274 PARK AND SQUAD RELAY ENERG. WHEN PARK BARKE IS SET

MULTI-

HORN

VIBRATOR

OSC

DRIVER

2/SEC

MODULE A2

TO L.E. FLAP EXTEND LIGHT

PWR AMPL

L.E. FLAP POS IND

CLOSED WHEN FLAP EXTENDED

28V DC (P6)

SPEAKER

L.E. FLAP AND SLAT POS IND UNIT AURAL WARNING DEVICES BOX (M315)


Takeo akeoff ff Warni arning ng Sche Schema mati tic c




B737-300/-400/-500 31-20




B737-300/-400/-500 31-20

28 V BAT. BUS

ONE OR BOTH THRUST LEVER ADVANCED AIRCRAFT ON GROUND

PARK BRAKE SET SPEEDBRAKE LEVER NOT IN DOWN DETEND TE-FLAPS <1 OR >15 UNITS LE-FLAP NO. 1 NOT EXTENDED LE-FLAP NO. 4 NOT EXTENDED STABILIZER >6.3 AIRPLANE NOSE UP STABILIZER <1.0 AIRPLANE NOSE DOWN y l n O s e s o p r u P g n i n i a r T r o F


Takeo akeoff ff Warni arning ng Log Logic ic

270604153-B737-3-ATA-27-L3-e

Recommend Documents