AP.IOI B-1902-15

-VULCAN B Mk.2

AIRCREW MANUAL RESTRICTED

'

t

'

2nd Edition

April 1970

RESTRICTED

VULCAN B

AIRCREW MANUAL

BY COMMAND OF THE DEFENCE COUNCIL

Prepared by Procurement Executive, Ministry of Defence

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AP lOlB-1902-1 5

(Formerly AP 4SOSB-PN)

(AL9)

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NOTES TO USERS

1. The Flight Reference Cards (AP 101B-1902-14) are issued separately from this manual and are subject to a separate amendment procedure but they are complementary to this manual and reference is made to them where necessary.

2. The Manual is divided by Marker Cards as follows:

Part 1. Description and Management of Systems Part 2. Limitations Part 3. Handling Part 4. Emergency Procedures Pan 5. Illustrations

Each Part is divided into Chapters as listed on its marker card. Each sheet is identified by a Part, Chapter, Page reference at the foot of the page. Thus a page bearing 1-3 Page 1 is Page 1 of Part 1 Chapter 3.

3. The limitations quoted in Part 2 are mandatory and are not to be exceeded except in an emergency. Instructions containing the word 'must' are also mandatory.

4. The Manual and its associated Flight Reference Cards aim to provide the best operating instructions and advice currently available. Although they provide guidance for most eventualities, they are not substitutes for sound judgement and good airmanship; moreover, they assume an adequate knowledge of the pertinent volumes of AP 3456. Furthermore, circumstances might require aircrew to depart from or modify the prescribed procedures and drills. Consequently the Manual and Flight Reference Cards should not be regarded as documents which are to be adhered to inflexibly at all times-other than as explained in the preceding paragraph.

5. Amendment lists will be issued as necessary and each amendment list instruction sheet will state the main purpose of the amendment and will include a list of modifications covered by the amend­ment. The list of pages will also be updated at each amendment. New or amended matter of importance will be indicated by triangles positioned in the text thus: ~- ... ..... ... ~ to show the extent of amended text and thus: ~ ~ to show where text has been deleted. Sheets issued by an amendment list will bear the AL number at the bottom of the odd-numbered pages and any amendment marks on either side of the sheet will relate to that amendment. However, when a new chapter is issued with an amendment list, or an existing chapter is completely revised, this fact will be noted within the heading of the chapter and amendment marks will not appear on those pages, apart from the AL number.

6. The following conventions are observed throughout the manual:

a. The actual markings on the controls are indicated in the text by capital letters.

b. Unless otherwise stated, all airspeeds, mach numbers, accelerometer readings, temperatures and altitudes quoted are indicated values.

c. WARNINGS are inserted only when the serious consequences of not following a procedure might otherwise be overlooked.

d. Information needing emphasis is printed in italics.

e. Nores are inserted to clarify the reason for a procedure or to give information which, while not essential ro the understanding of the subject, is useful to the reader.

7. Modification numbers are only referred to in the text when it is necessary to differentiate between the pre- and post-mod states. For ease of reference, a list of modifications mentioned in the text is included in the preliminary pages, with a cross-reference to the location in the text of the modification details.

IMPORTANT

Comments and suggestions should be forwarded to the Officer Commanding, Royal Air Force

Handling Squadron, Boscombe Down, Salisbury, SP4 OJF.

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•

RESTRICTED AP 101B-1902-15

AMENDMENT RECORD SHEET

To record the incorporation of an Amendment List in this publication,

sign against the appropriate AL No. and insert the date of incorporation

AL No. 1AfENDJD BY DATE AL No. AMENDED BY DATE

j q;~; ~ 34

~ ....................................... .. ·· ···· ·~ ............................................................. . 35

36

:······ ··· ·~············ :l;~: ············ ~····· . z ~=~ l:J:l;,. ····· ·····:·············· ·· · ······~····· ······ ·· ······ ··· · ··········;p~· · · ..... .. ................................ ~ ...................... 1/jf/!.?. ..... .

.......... !~ ············ · · ··· ····~····· ··· ······ ·· · ······~~~.·:

37

38

39

40

41

42

43

11 44 .................................................................................... ...........................

12 45

13 46

14 47 .......................................... .. .................................. .. ......... .......................

IS 48

16 49 ................................................... ....... .. ................... ........ .. .......... .. ...........

17 so 18 51

19 52

20 53 ................................... .. ....... ....... ....... ....... ...............................................

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22 55 .................... ·················································· ······························•····· ····

23 56 ························································································· ·······················

24 . 57

25 58 ... .................................. .............................. ...................... .......................

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27 60 . ·················· ..............................................•................ ............................ ... ................... ........ ..................................................................................

28 61

29 62 .......................................... ... ...................................................................

30 63

31 64 ..................... .................................................................................... .......

32 65

33 66

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ANA No 1

RESTRicrED

ANA INCORPORATED

z 13 1±1

AP 101B-1902-15 Preliminaries

ANA are to be incorporated in manuscript and the ANA serial number is to be recorded above.

LIST OF PAGES AT AL 10

Pages which have manuscript amendments are marked with an asterisk.

Pag6 Issued by Page Issued by Page Issued by

Title page AL 9 1-5 Marker 1-12 Marker Card AL 1 Card AL 1

Amendment 1-5 page 1 AL 7 record Initial

3• Initial 1- 12 page 1 AL10

Prelim 1 AL10 5 AL 5 3 AL10 page

3 AL 3 7 AL 9 5 ALIO

5 AL10 9 AL 7 7 AL10

7 AL 7 1-6 Marker 9 AL10

9 AL9 Card AL 1 11 AL10

11 AL 7 1-6 page 1 AL 8 13 AL10

3 AL 3 15 AL 4

1-7 Marker 17 AL10

Part 1 Card AL 1 19 ALIO

Marker Card Initial 1-7 page 1 AL 4 1-13 Marker 1-1 Marker AL 1 3 AL 8 Card AL 1

Card !5 AL10 1-13 page 1 AL 8 1- 1 page 1 AL10 7 AL10 3* AL 6

3 Initial 1-8 Marker 5 AL 9 s AL10 Card AL 1 1-14 Marker 7 AL10 1-8 page 1 AL 9 Card AL 1

3 AL 2 1-14 page 1 AL 8 l-2 Marker Card AL 1 5 AL 9 3 AL10

7 AL 7 5 AL10 1-2 page 1 AL10

3* AL 3 9 AL 6 1-15 Marker

11 AL 8 Card AL 1 5 AL 9

7 AL 9 13 AL 9 1-15 page 1 AL 10

1--9 Marker 3 Initial

9 AL 9 Card AL 1 5 AL10

11 AL 9 1-9 page 1 AL 10 7 AL 5

3 AL10 1-3 Marker 5 AL10 Card AL 1 Part 2

1-3 page 1 AL10 1-10 Marker Card* AL 1

Marker Card Initial

1-10 page 1 AL 9 2-1 Marker Card AL 1

1---4 Marker 3* AL 2 Card AL 1 5

2- 1 page 1 AL 9 AL 9

3 AL 8 1---4 page 1 AL 6 7 AL 9

2-2 Marker 3* AL 3 9 AL 8 Card AL 1 5 AL10 1-11 Marker 2-2 page 1 AL10 7 AL 6 Card AL 1 2-3 Marker 9 AL 7 1-11 page 1 AL10 Card AL 1

11 AL 9 3* Initial 2-3 page 1 AL 9

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Page Issued by

Part 3

Marker Card AL 8

3-1 Marker Card AL 1

3-1 page 1 AL 8

3 AL 9

5 AL 9

3-2 Marker Card AL 1

3-2 page 1 AL 9

3 AL 9

5 AL 9

3-3 Marker Card AL 8

3-3 page 1 AL 9

3 AL10

5 AL 8

3-4 Marker Card AL 8

3-4 page 1 AL 9

3* AL 8

3-5 Marker Card 3-5 page 1

Part4

Marker Card

4-1 page 1

3

5

Part 5

Marker Card 5-1 page 1

3

5

7 9

11

AL 8

AL 9

Initial

AL 9

AL 9

AL 9

Initial

Initial

Initial

Initial

Initial

Initial

Initial

Prelim Page 1 (AL 10)

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MODIFICATION NUMBERS MENTIONED IN TEXT

Mod No

2193

2321

2330

2393

2465

Effect

Pressurisation of radio equipment ...

Nitrogen pressudsation for hydraulic system

Redesigned pressure-head ...

Combined static linejoxy hosejmictel release coupling at

rear crew station ...

New Tacan aerial .. .

CM 0291/Mod 2505 Transfer of event marker supply to F95 camera ...

SEM012 .. . Manual operation of wing and fuselage tank fire

extinguishers

~ SEM 027 I Mod 2504 Abandon aircraft switch automatically operates cabin light

(SEM 057 in MRR aircraft)

SEM 041/Mod 2503 RAT field switch on AEO's panel lOP

SEM 045/Mod 2502 F95 Mk 4 camera (MRR only)

SEM 04t6/Mod 2500 IF BIAS voltmeter at navfradar station

Prelim Page 2

SEM 058 Dimmer facility on rear crew TFR fault lights . ..

STI 410 Airbrake setting at high drag changed to 77° -+- 3°

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Location in text of Mod details Pt, Chap, Para

1, 13, 24

) 1, 10, 6 l1, 13, 21

1, 12, 88

,4, 1, 7

1, 14, 22

1, 3, 7

1, 6, 9

4, 1, 7

1, 4, 22

. 1, 3, 7

1, 14, 33

I, 15, Table 1

1, 7, 69

Prelim. Fig. 1 Vulcan B Mk. 2

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AP lOlB-1902-15 Preliminaries

Prelim. Page 3 (AL 3)

DOPPLER BENEATH PORT WING AFT OF MAINWHEEL BAY.

FORWARD>~h /~• RED

1 SHRIMP UNDERSIDE OF STARBOARD ENGINE NACELLES FROM REAR.

GLIDEPATH ILS .

LOCALISER (ILS)

BLUE DIVER

TACAN

RA /ALT Mk.7 (BENEATH AIR

INTAKE) ALT 6 (Tx)

(BENEATH INTAKE EACH SIDE)

Prelim. Fig. 2 Vulcan B Mk. 2-Aerial Location

Prelim. Page 4

E2B COMPASSES

CENTRE PANEL

RETRACTABLE CONSOLE

Prelim. Fig. 3 Cockpit Layout

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UHF LOWER

SAGA

SENSE AE RAD/COMP

RESTRICTED AP IOlB-1902-15 Preliminaries

INTRODUCTION

~General

1. The Vulcan B Mk 2 delta-wing, all .... metal aircraft may be powered by four Olympus 200 series engines, each developing 17,000 lb static thrust at sea level in ISA conditions, or by Olympus 301 engines, developing 18,000 tb static thrust. Provision is made for air-to-air refuelling. A ram air turbine and an airborne auriliary power plant provide for emergency electrical supplies. Equipment bays outside the pressure cabin contain power distribution and fuse panels in addition to various components of the main services :

a. Nose section.

b. Nosewheel bay.

c. Main undercarriage bays.

d. Bomb bay.

e. Power compartment (aft of bomb bay).

Crew

2. The aircraft is operated by a crew of five: 1st pilot, co-pilot, navigator/radar, navigator/plotter and air electronics officer (AEO). 11he pilots sit side-by-side in ejection seats on a raised platform at the front of the cabin normally referred to as the cockpit. Behind the pilots, facing ak, the rear crew members sit on bucket-type seats (the outer two swivelling) facing one long table, behind which is a crate carrying their equip­ment. A prone bombing position is in a blister below the pilots' floor.

Instrument Layout

3. The pirots' instruments and controls are on the front panel (divided into four sections), the port and starboard consoles, the centre (retraotable) ronsole and the throttle quadrant (see Fig 3 opposite).

4. Magnetic indicators and warning lights for the vital systems are grouped across the top of the pilots' front panel and consist of :

a. Two amber MAIN WARNING lights, one at either side, which indicate failure of the PFC units, feel units or autostabilisers (except the yaw dampers).

b. One red alternator failure (ALT FAll.) light which illuminates foll~g a single alternator failure and flashes when two or more alternators fail.

c. Eight magnetic indicators (fuur each side of the alternator light) for PFC units, artilidal feel, auro­sta'bilisers, airbrakes, bomb doors, canopy, entrance door and pressure-bead heaters.

5. On the coaming above the pilots' front panel are four engine fire warning lights which also serve as fire extinguisher pushbuttons.

In addition, two red fire warning lights for the wing fuselage and bomb bay fue1 tanks are at the top of <the co-pilot's instrument panel. Post-SEM 012, only the bomb bay fire warning light is fitted.

6. A list of panels accessible in flight and their nomenclature is given below; the pilot's and some rear crew members' panels are shown in Part 5.

lP Pilots' instrument panel (four sections)

2P Fuel contents panel (forward of throttle quadrant)

3P Fuse and relay panel (port side of cabin)

4P Fuse and relay panel (starboard side of cabin)

SP Retractable console

6P Port console

7P Statboard console

8P Prone bombing panel

9P Bombing control panel (nav fradar)

lOP Alternator control panel (AF.O)

HP Radio and radar supplies panel (rear pressure bulkhead)

12P llSV 1600Hz switch panel (nav /radar)

24P No 1 llSV 3-phase distribution panel (star­board side cabin, under rear crew floor)

ZSP No 2 llSV 3-phase distribution panel (star­board side cabin, under rear crew floor)

48P 28V DC sub-distribution panel (rear pressure bul:kihead)

SOP Secondary supplies switch panel (AEO)

54P Intercom panel (port side cabin under crew floor)

68P Failure warning distribution panel (below pilots' floor)

69P Air ventilated suit fuse panel (below rear table)

70P AAPP control switch panel (AEO)

71P MFS power panel (port side, below pilots' floor)

72P MFS power panel (starboard side)

75P 200V sub--distribution panel (starboard side cabin, under crew floor)

81P ECM switch panel (AEO)

lOOP Gold film windscreen panel ~

RESTRICTED Prelim

Page 5 (AL 10)

Prelim Page 6,~

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ABBREVIATIONS USED IN THE TEXT

AAPP

ADD

ADF

ADP

AEO

ARI

AVS

BCF

BRSL

CSDU ...

DG

PV

ECM

EHPP

FRC

GPI

HF/SSB

HFRT ...

HRS

"IFF/SSR:

ILS ... .. MFS

NBC . ... NBS : -.... :

NHU ... NMBS ...

NRV

0/H

PDJ

PEC

-PESJ

PFC

PFCU

PSP ... RAT ... TASU ...

~ TBC

TFR

TFRU ...

TRU

VSI

'.

...

...

'.

Airborne auxiliary power plant

Airstream direction detector

Automatic direction finding

Azimuth director pointer

Air electronics officer

Air radio installation

Air ventilated suit

Bromochlorodifluoromethane

Bomb release safety lock

Constant speed drive unit

Directional gyro

Direct vision

Electronic countermeasures

Emergency hydraulic power pack

Flight reference cards ·

Ground position indicator

High frequency, single side-band

High frequency radio telephony

Heading reference system

· Identification friend or foe, secondary surveillance radar

· Instrument landing system ·

Military flight · system

· Navigation & bombing computer

Navigation & bombing system . Navigatot's heading ti.nit ·· ::

Normal ;'l!aximum bra~ing speed

Non-retUrn valve

Overheat

Pilot direction indicator

Personal equipment connector

Pilots emergency stores jettison

Powered flying controls

Powered flying control unit

Personal survival pack

Ram air turbine

True airspeed unit

Tail brake parachute

· Terrain following radar

Terrain following radar unit

Transformer/rectifier unit

Vertical speed indicator·

RESTRICTED AP lOlB-1902-15 Preliminaries

LIST OF ASSOCIATED AIR PUBLICATIONS AND AIR DIAGRAMS

Title

Aircraft Assisted-Escape Systems:

~ Ejection Seats Types 3KS1, 3KS2, Mk 4

Time-delay Firing Unit

Canopy Jettison Equipment

Aircraft Automatic Stabilisers

Aircraft Pressure Fuelling Equipment: TTansfer Valve Mk 9 ...

Fuelling Indicator Mk 7

Air Publications

Aircraft Pressurising and Air Conditioning Equipment ...

Aircraft Vapour Cycle Heat Transfer Equipment

Pressure Regulating Valve

Aircraft Refuelling in Flight

Airstream Direction Detectors

Armament (excluding Gmded Weapons) ...

Automatic Pilot Mk 10 B/C

Weapon Response Simulator Type 105 Mk 1

Decca Doppler

Engine Starting System, Rotax Type CT1303

Electrical Equipment

Fire Protection

Green Satin

Hydraulic Equipment: Accumulators

Reservoir

Reducing Valve

Thermal Relief Valve

Instruments

Guided Weapons ...

~Navigation and Bombing Systems

Olympus 20101, 20201 and 20301 ECU's

Olympus 30101 ECU

Pneumatics: High Pressure Relief Valve, Dunlop

Pressure Relief Valves Hymatic

~ Powered Flying Controls and Equipment

Radar Manual

Radio Altimeter

Ram Air Turbine, Plessey

Rover AAPP

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AP No

109B-0122-1

109C-0203-1

109F-0114-1

112C-2001-1

106D-0207-16C

106D-4701-1

1 07B-OOO 1-1

107B-0104-16A

107B-0106-16A

106D Series

112G-0113-1

110 Series

112C-0800-1B

110M-0101-1A

114E-1600-16

103D-0103-3A

113 Series

107E-0102-1

114E-0300-1

105B-0313-16C

105B-0318-16C

105B-0731-1

lOSB-0738-16

112G Series

118 Stties

114 Series

102C-0402-2

102C-0403-l

lOSC-0514-1

105C-0550-16

105D Series

114A Series

116B-0203-1

105C-1005-16A

1 02F-0204-16A

Prelim Page 7 (AL 7)

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Air Publications (continued) Title

Safety Equipment:

~ Parachute Assembly, Type B, Mk 42

Parachute Assembly, Back Type, Mk 46 and 49

Pressure Jerkin Mk 3

Pressure Jerkin Mk 4

~ Anti-G Suits Mk 7B

Life Preserver Mk 17/17 A

SARBE ...

Terrain Following Radar Unit, Type AN/APN-170

UHF ARC 52 (ARI 18124)

V /UHF PTR 175 (ARI 23143) ...

Undercarriage Equipment Dowty

Undercarriage Equipment Dowty-Rotol

Mainwheels, Dunlop

Nosewheels, Dunlop

Vulcan B Mk 2, Aircraft Servicing Manual

Access Panels and Walk-ways

Cabin and Aircraft Pressurisation

De-icing Systems ...

Electrical Installation

Electrical Location

Emergency Equipment and Exits ...

Engines: 200 Series

300 Series

AAPP

Escape Drill

Fire Extinguishing Systems

Flying Controls and Lubrication

Fuel System

Hydraulic System ...

Jet Efflux, Danger Areas

Lubrication

Pneumatic System ...

Powered Flying Controls

Radio Installation

Radio Location

Prelim Page 8

Title Air Diagrams

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. .

AP No

108C-0116-l

108C-0117-1

108F-0301-l

108F-0302-1

108F-0401-l

108F Series

116B-0901 Series

114N-0100-l

116D-0105-l

116D-0116-1

1803E

1803V

104F-1017-1

104G-1000-1

101B-1902-1 SeJJies

AD No

101B-1902-Dll

7086V

101B-1902-D27

101B-1902-Dl2A 101B-1902-Dl2B

101B-1902-Dl3

7086N

102C-0402-D

102C-0403-D

102F-0204-D

7568 101B-1902-D20

101B-1902-D2

101B-1902-D4

101B-1902-D3

101B-1902-D6

lOIB-1902-D 1

101B-1902-D28

IOSD-1900-Dl

101B-1902-Dl4A, B

lOIB-1902-DlSA, B

• Principal Dimensions

Overall length ...

Wing span

Height to top of fin

Wheel track

Wheel base

Fuei System

Type of fuel:

~ Normal use

NATO Code Designation F34 Avtur F40 Avtag

Operational necessity F3S Avtur

F44 Avcat

Fuel Capacities

Tank Group Tank No

1 and 4 1 (outboard, port 4 and starboard) s

7

Total each group

Total both groups

2 and 3 2 (inboard, port 3 and starboard) 6

Total each group

Total both groups

Total fuel

Tank

A (fwd saddle) ...

E (aft saddle) ...

Cylindrical (each usable)

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Leading Particulars

UK Spec

DERD2453 DERD 2454

DERD2494

DERD2452

Feet Inches

105 6

111 0

27 1

31 1

30 1•5

us Spec and Type

MIL T-83133 (JP8) MIL T-5624 (JP4)

MIL T-5624 (JP5)

Main Tanks

Avtur (8 lb I gall)

Gallons Pounds

2X 610 2 X 4880 2X 630 2 X S040 2X SIS 2 X 4120 2X S6S 2 X 4S20

2 X 2320 2 X 18,S60

4640 37,120

2X 93S 2X 7480 2X 630 2X 5040 2 X 74S 2X 5960

2 X 2310 2 X 18,480

4620 36,960

9260 74,080

Bomb Bay Tanks

AP lOlB-1902-15 Preliminaries

} Contain FSII

-No FSII, may contain Hi tee

- Contains FSII, may re-quire Hitec ~

Avtag (7 · 7 lb I gall)

Gallons Pounds

2X 620 2X 4774 2X 640 2X 4928 2X S2S 2X 4042·5 2X 57S 2x 4427·5

2 X 2360 2 X 18,172

4720 36,344

2X 94S 2X 7276·5 2X 640 2 X 4928 2X 7S5 2X 5813·5

2 X 2340 2 X 18,018

4680 36,036

9400 72,380

Pounds Gallons

. .. 718

. .. 721

. .. 99S

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Avrur

5744

5768

7960

Avtag

5529

5552

7662

Prelim Page 9 (AL 9)

Nitrogen Pressurisation Cylinders Charging pressures

Power Units Name Type Number Static thrust (200 series) Static thrust (Mk 301) 000

External starting air

Engine Oil System Type Oil

Tank capacity (200 series)

· Tank capacity (Mk 301)

Constant-Speed Drive Oil System Type Oil

Tank capacity

Starting System Low-pressure air starter motor 0 0 0

Rapid start

Engine Driven Auxiliaries Hydraulic pumps Constant-speed drives and alternators

Airborne Auxiliary Power Plant Type Number 000

Oil System Oil

Sump capacity Starting oo,

Ram Air Turbine Type Output

Hydraulic System Pumps VVorking pressure Reservoir capacity System capacity 0 0 0

Fluid

Prelim Page 10

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3 (60/1825) 2 at 3000 PSI 1 at 1800 PSI

Olympus 200 series or Mk 301 Turbo-jet Four 17,000 lb, sea level, ISA 18,000 lb, sea level, TSA Oo8lbfsec at 40 PSI

Integral with engine OX-38 (34A/9100591, NATO 0-149, D Eng RD

2487) 4~- galls oil, 1! galls air space System capacity 2! galls 3i galls oil, 1 i galls air space System capacity 2:! galls

Attached to engine OX-38 (34A/9100591, NATO 0-149, D Eng RD

2487) 11 pints oil, 3-! pints air space, system capacity

12! pints

Rotax CT0806/1 (one per engine) CT1303

Three (Nos 1, 2 and 3 engines) Four, 40 kVA, one per engine

Gas turbine (Mk 10301) One Sump type OX-38 (34Af9100591, NATO 0-149, D Eng RD

2487) 4! pints (total oil in system 5 pints) Cartridge or electric

Plessey TRA 170/26 17 kW, 200 volts, 3-phase, 400 Hz

Three (engines Nos I, 2 and 3) 3600-4250 PSI 2t galls 12 galls OM-15 (34Bf9100572, NATO H-515)

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•

Hydraulic Power Pack Type Operating pressure Motor Capacity Fluid

Undercarriage Emergency Lowering System Cylinders Charging pressure

Pneumatic System (Nitrogen) Entrance Door System

Cylinders Charging pressure

H2S scanner Cylinder Charging pressure

NBS Cylinder Charging pressure

Air Conditioning System Turbine unit Oil Capacity ...

Flying Controls Power Units

Outer elevons (four) Inner elevons (four) Rudder (two) Oil Motor gearbox oil

Capacity (each) Outer elevons .. . Inner elevons .. . Rudder

Air brakes Oil (gearboxes)

Bomb Aimer's Windscreen De-icing Tank capacity Fluid

Aerofoil and Engine Anti-Icing Thermal system, using engine air

RESTRIGrED

E8908Y 3500 to 3900 PSI AC, 3-phase 11 pints OM-15

Two Mk 5 3000 PSI

~ Three Mk SF ~ 2000 PSI

One (60 /9429885) 1800 PSI

One (60/9429885) 1800 PSI

Type BT15, Mk 2A

M lOIB-1902-15 Preliminaries

OX-38 (34A/9100591, NATO 0-149) 210 cc

PBS P132 P138 OM-15 (34B/9100572, NATO H-515) EEL3

4 pints 4 pints 8 pints

OX-14 (34B/9100589, NATO 0-147)

12 gails + t gall air space AL-8 (HB/9100475, NATO S-738)

RESTRICTED Prelim

Page 11 (AL 7)

Fire Extinguishers Engines

Type ... Number

Wing tanks Type ... Number

Fuselage tanks Type ... Number

Leading edge Type ... Number

Crew's cabin Type ... Number

AAPP Type ... Number

External compartment Type ... Number

Bomb bay Type ... Number

Oxygen System Regulator Cylinders Number ... Charge pressure ...

Safety Equipment Ejection seats:

1st pilot Co-pilot Life raft

Electrical System Type Alternators

Number Type ... Voltage

RAT alternator AAPP alternator Transformer rectifier units

Battery Transformers

Frequency changers

Prelim Page 12

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Mk 13A, methyl-bromide Four

Mk 14A, methyl-bromide Twelve

Mk 13A, methyl-bromide Four

Mk 13A, methyl-bromide Six

34H, BCF Five

Mk 4AX, methyl-bromide (27N/152) One

34H, BCF One

Mk 13A, methyl-bromide Eight

Mk 21A, 21B or 17F demand system Mk lOA (6D/9429900) (2250 litres) Twelve 1800 PSI

Type 3KS1 Mk 4 Type 3KS2 Mk 4 MSS

AC

Four Type 175, 40 kVA 200 volt, 3-phase, 400 Hz, neutral earth 22 kVA 40 kVA Two, 7 · 5 kW, 28 volt, earth return One, 0·75 kW One, 112 volt One, type K, 24 volt, 40 AH Two, 3 kVA, 115 volt, 400 Hz, 3-phase One, 1 kVA, 115 volt, 400 Hz, 3-phase One 40 VA, 115 volt, 400 Hz Two, 3 kVA, 115 volt, 1600 Hz, single phase

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e RESTRICTED AP lOlB-1902-15 Preliminaries

MODIFICATION NUMBERS MENTIONED IN TEXT

Mod No

2193

2321

2330

2393

SEM 012

Effect

Pressurisation of radio equipment ...

Nitrogen pressurisation for hydraulic system

Redesigned pressure-head ...

Combined static linejoxy hosejmictel release coupling at

rear crew station ...

Manual operation of wing and fuselage tank fire

extinguishers

SEM 027 j Mod 2504 Abandon aircraft switch automatically operates cabin light

STI 410 Airbrake setting at high drag changed to 77°+3°

RESTRICTED

Location in text of Mod details Pt, Chap, Para

1, 13,24

~ 1, 10, 6 ~ 1, 13, 21

1, 12, 88

4, 1, 7

1, 6, 9

4, 1, 7

1, 7, 69

Prelim Pa.ge 13 (AL 9)

Prelim Paee 14

~

RESTRICTED

ABBREVIATIONS USED IN THE TEXT

AAPP Airborne auxiliary power plant

ADD Airstream direction detector

ADF Automatic direction finding

ADP Azimuth director pointer

AEO Air electronics officer

ARI Air radio installation

AVS Air ventilated suit

BCF Bromochlorodifluoromethane

BRSL Bomb release safety lock

CSDU ... Constant speed drive unit

DG Directional gyro

DV Direct vision

ECM Electronic countenneasures

EHPP Emergency hydraulic power pack

FRC Flight reference cards

GPI Ground position indicator

HF/SSB High frequency, single side-band

HFRT ... High frequency radio <telephony

HRS Heading reference system

IFFJSSR Identification friend or foe, secondary surveillance radar

ILS Instrument landing system

MFS Military flight system

NBC Navigation & bombing computer

NBS Navigation & bombing system

NHU Navigator's heading unit

NMBS ... N onnal maximum braking speed

NRV Non-return valve

0/H Overheat

PDI Pilot direction indicator

PEC Personal equipment connector

PESJ Pilots emergency stores jettison

PFC Powered flying controls

PFCU Powered flying control unit

PSP Personal survival pack

RAT Ram air turbine

TASU ... True airspeed unit

TBC Tail brake parachute

TFR Terrain following radar

TFRU ... Terrain following radar unit

TRU Transfonnerjrectifier unit

VSI Vertical speed indicator

RESTRICTED

AP lOlB-1902-15

PART 1

DESCRIPTION AND MANAGEMENT OF SYSTEMS

List of Chapters

Title Chap.

AIR CONDITIONING 1

AIRCREW EQUIPMENT ASSEMBLIES AND OXYGEN SYSTEM 2

ARMAMENT AND OPERATIONAL EQUIPMENT 3

ELECTRICAL SYSTEM 4

ENGINES AND AAPP . . . 5

FIRE PROTECTION SYSTEMS 6

FLYING CONTROLS . . . 1

FUEL SYSTEM . . . 8

GENERAL AND EMERGENCY EQUIPMENT AND CONTROLS 9

HYDRAULIC SYSTEM AND EMERGENCY AIR SYSTEM. 10

ICE PROTECTION 11

MILITARY FLIGHT SYSTEM, AUTO-PILOT AND OTHER FLIGHT INSTRUMENTS 12

PNEUMATIC SYSTEMS 13

RADIO AND RADAR ... 14

TERRAIN FOLLOWING RADAR 15

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RESTRICTED AP lOlB-1902-15

PART I

CHAPTER l-AIR CONDITIONING (All text completely revised by ALlO)

Contents CABIN AIR SYSTEM

Cabin Air, General Cabin Air Conditioning Unit Cabin Pressurisation Cabin Ventilation Cabin Air Conditioning, Controls and Indicators Operation of the Pressurisation System

Para 1 4

10 15 17 23 27 28 29

Cold Air Turbine Overspeed Loss of Cabin Pressure Emergency Decompression

AIR-VENTILATED SUITS SYSTEM Normal Air-Ventilated Suits (AVS) System AVS Air-Conditioning Unit

31 32 35 36 37

A VS Components in Cabin Auxiliary A VS System ... A VS Controls

BOMB BAY CONDITIONING Bomb Bay Conditioning System Bomb Bay Overlleat Drill

WINDSCREEN THERMAL DEMISTING Windscreen Demisting Supplies

39 40

Windscreen Demisting Controls and Operation 41 42

ELECTRICAL SUPPLIES Electrical Supplies

Illustrations Cabin Pressurisation and Air Conditioning Air-Ventilated Suits System

43

Fig 1 2 3 Cabin Pressure Control

CABIN AIR SYSTEM

Cabin Air, General 1. The cabin air conditioning and pressurisation system maintains comfortable temperatures and pressures within the crew compartment. Hot pres­surised air tapped from the engine compressors, cooled by cold ram air 11nd a cooling turbine, is distributed throughout the cabin via ducting. The temperature of the conditioned air is controlled by varying the propor­tion of bot air which flows through or bypasses the air cooler or the cooling turbine. The controlled air flow out of the cabin is used to cool equipment in the radome. Provision is made for conditioning the cabin air both on the ground and in unpressurised flight.

2. The cabin pressure is determined by the amount of air allowed to flow out of the crew compartment and is maintained by two pressure controllers.

Pressurisation can be set for eitlher cruise or combat conditions. Provision is made for emergency depres­surisation.

3. The main controls for cabin heating and pressur­isation are grouped together on panel 7P on the starboard console.

Cabin Air Conditioning Unit 4. The air conditioning unit in the nosewheel bay consists of an air-to-air cooler, a temperature control valve (TCV), a cooling brake turbine unit and a water separator.

5. The cooler is supplied with cold air from a ram air intake between the cabin and cite port engine air intake. The cold air passes through a rearward-facing duct below the unit.

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Page 1 (AL 10)

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6. The brake turbine unit is an inward flow turbine coupled to a centrifugal braking compressor. The turbine passes air from the TCV to a water separator and thence to the cabin. The compressor passes filtered air from the nosewheel bay through the exhaust duct. The speed of the turbine is monitored by a pressure ratio switch which, if the turbine overspeeds, auto­matically selects a warmer setting on the TCV thereby reducing the amount of air passed to the turbine. The COLD AIR UNIT OVERSPEED WARNING MI on the starboard console shows white, reverting to black when the overspeed !has ceased.

7. The temperature control valve is electrically­operated, either automatically or manually. When maximum heat is selected, engine air passes directly from the flow valves, t'hrough the TCV, to the cabin. As the temperature setting is reduced, air is pro­gressively allowed to pass through rhe cooler en route to the TCV. At the colder settings, the air passes to t!he cabin .through the cooler, TCV 'and turbine.

8. An underheat sensing element opens a bypass valve when the air from the turbine falls below 2°C value, t>.hus allowing warm air to mix with the cold air before it reaches the water separator.

9. An overheat switch operates to move the control valve towards the cool position when the output temperature rises ro 175°C.

Cabin Pressurisation

10. Cabin pressurisation is achieved by controlling the rate at which the air fed into the cabin is allowed to escape. Each of two pressure controllers, in the nose below the pilots' floor, supplies counter­pressure to one of the bellows of the combined valve unit in the front pressure bulkhead. One controller is motorised allowing a CRUISE or COMBAT setting to be selected. The other is unmotorised and acts as a standby in the CRUISE setting. The ground test levers on the controllers must always be fully down in flight.

11. Pressurisation begins at 8000 ft and a cabin altitude of 8000 ft is maintained until the maximum differenri<d pressure is reached. In CRUISE, the maxi­mum differential pressure is 9 PSI, attained at approxi­mately 47,000 feet. In COMBAT, the maximum differential of 4 PSI is reached at approximately 19,500 feet, ~hove t>.hese altitudes, cabin altitude increases. The change from CRUISE to COMBAT setting takes place at 12 PSI per minute and from COMBAT to CRUISE at 1 PSI per minute.

12. Safety Devices. If cabin pressure fal1s ! PSI below selected pressure, a contact in the motorised controller operates a warning horn and illuminates three red

warning lights, one at each rear crew member's posmon. The warning hom may be isolated by a switch at the rear of panel 4P. To prevent the build-up of negative differential pressure, an inward relief valve on t>.he front pressure bulkhead operates at :! PSI. An outward relief valve, also on the front pressure bulk­head, operates at 9-} PSI to limit over-pressurisation. To safeguard against the cabin extractor ducting collapsing during rapid outflows of air, the combined valve unit is fitted with a duct relief valve. A decom­pression flap, visible from the cockpit on the upper port si:de of the nose fairing, allows excess air to escape from the nose radome.

13. Flood Flow. Provision is made for flood flow but !'he system is inoperative.

14. Decompression. To allow decompression of the cabin in an emergency, air release valves (in the lines between the pressure controllers and the bellows of !'he combined valve unit) can be operated to remove the counter-pressure from the bellows. Operation of the valves is controlled electrically by either pilot and electrically and mechanically by the rear crew. While the combined effect of the operation of the air release valves, duot relief and decompression flaps ensures a rapid release of cabin pressure, it may take up to 30 seconds for the pressure to fall sufficiently to allow the door to open.

Cabin Ventilation

15. During unpressurised flight, the cabin can be ventilated via the ram air valve on the port side of the cabin. This allows air from the cabin conditioning ram air intake to enter the cabin but, unless the cabin switches are shut, the effect is negligible. The ram air valve should be closed before pressurising roe cabin. The cabin may :be ventilated on the ground with air supplies from the cooling air unit, via the normal control valves.

16. Individual face blowers (pll11'kah louvres) for the rear crew are powered by a 200V AC blower unit under the navigator's table, controlled by a FACE BLOWING ON/OFF switch on rhe edge of the AEO's table.

Cabin Air Conditioning, Cont.rofs and Indicators

17. General. The main cabin conditioning controls and indicators are grouped on panel 7P.

18. Engine Air Switches. The four ENGINE AIR OPEN/SHUT switches on the out'board side of the panel control the supply of engine air to the flow control valves, airframe anti-icing systems and to certain ot>.her services (see Chap 13).

1-1 Page 2 RESTRICTED

------------------------------------------------------------------------------------

TEMPERATURE CONTROL VALVE SETTINGS ~-THIS IS ONLY A DIAGRAMMATIC REPRESENTATION.IN PRACTICE THE CONTROL IS INFINITELY VARIABL£.

AP 101B-1902-15 Air Conditioning

J:( j(R TU)(COOLER CABIN ENGINES CABIN

HOT WARM COLD

RAM AIR­

INTAKE -

REAR PRESSURE BULKHEAD

VAJ.YE -

- HOT AIR SUPPlY

- COlD AIR SUPPLY

- CONDITIONED AIR

t

t

i-TO AIR

VENTILATED SUITS

DUCT REUEF VALVE

5S PSI

-

1--1 Fig. 1 Cabin Pressurisation and Air Conditioning

PRESSURE CONTROLLER

CONTROL SWITCH

Q

RAM AIR=:

CABIN AIR

SWITCH I I

SHUT-OFF VALVES

AUTO­FLOW

FROM STARBOARD ENGINES

I I

CABIN AIR

SWITCH

TO AIR VENTILATED

c~Jl~~L

SUITS -

TO CABIN AIR CONDITIONING

(FIG. 1) 1'

FROM PORT ENGINES

AIR FROM ENGINES OR A.A.P.P.

I I I I I I I I I

FLOW AUG MENTER

{,

_.1

OVER­HEAT SWITCH

REAR BULKHEAD

EXHAUST

1-1 Fig 2. Air-Ventilated Suits System

RESTRICTED

TEMPERATURE CONTROL

VALVE

TEMPERATURE CONTROL

VALVE

ON- OFF COCK

AIR TO AIR

COOLER

- HOT AIR SUPPLY

- COLD AIR SUPPLY

- CONDITIONED AIR

.1- 1 Page 3

-

~ (I)

;d

~

"d ., ~ V\

~-ol .......- -

i ,.. i ~

~ 5'

I ~

t

TO STATIC V!f\TS

~ \

RELIEF / VALVE

I RADAR COOLING

DIAPHRAGM

"-...., r--COMBIN.F,R ,"\ JALVE U"' T

~-I !l . ..

DECOMPRESSION FLAP

-

INWARD RELIEF VALVE

WATER TRAP

GROUND

-

WATER TRAP

WATER TRAP

TYPE 'F'

'I lr II

PRESSURE CONTROLLER r

_FIXED ORIFICE

I I ,,

II II

I

7 I

CABIN DECOMPRESSION CONTROL AT CREWS STATION.

-CONTROL VALVE

MOTORISED PRESSURE CONTROLLER

_ DIFFERENTIAL CAPSULE

e

~ (I)

;d

~ 0

~?d g§ 0~ P,l -·-o."' g ~ s·.!... 00 V\

RESTRICI'ED

19. Cabin Ait Switches. The two CABIN AIR OPEN /SHUT switches, beside the engine air switches, open the ·shut-off valves, allowing the engine air to pass to the air conditioning unit. One switch controls the starboard supply and one the port. When a switch is at OPEN, the auto-flow valve automatically regulates the volume of air passing to the system.

20. Cabin Temperature Controls. Cabin temperature can be controlled automatically or manually. When the CABIN TEMP CONTROL switch is moved outboard to AUTO, the temperature is automatically controlled according to the setting of the AUTO TEMP SELECTOR rotary control beneath the TEMP CON­TROL VALVE indicator. For MANUAL control, the switch is moved inboard and either forward or aft to the COLD or HOT position until the desired setting is obtained. The switch is spring-loaded from both these positions to the central (neutral) position.

21. Ram Air Valve. The ram air valve is controlled by a guarded SHUT /OPEN switch, spring-loaded to the central (neutral) position, as shown on the adjacent RAM AIR VALVE indicator.

22. Cabin Pressw·e Controls. Cabin pressurisation is controlled by a 3-position CRUISEJCOMBATJNO PRESSURE (gated) PRESSURE SELECTOR. Cabin decompression (or lack of pressurisation) can be achieved by selecting the PRESSURE SELECTOR to NO PRESSURE, or by operating either pilot's ABANDON AIRCRAFT switch, the EMERGENCY DECOMPRESSION switch on the port console or the CABIN PRESSURE RELEASE lever above the Nav /plotter's position. There is also a switch in the nosewheel bay. The ABANDON AIRCRAFT and EMERGENCY DECOMPRESSION switches are lock-toggle types which must be pulled up before selection.

Operation of the Pressurisation System

23. Before starting the engines, set the controls as given in the Flight Reference Cards.

24. After starting engines, set all ENGINE AIR switches OPEN and instruct the ASC to carry out a ducting leak check, after which all ENGINE AIR switches are to be SHUT. Except when airframe anti­icing is required or if flight safety considerations dic­tate otherwise, switches are normally to be as follows:

a. During taxying, all ENGINE AIR switches SHUT. b. For take-off, climb and descent 1 and 2 ENGINE AIR switches with the PORT CABIN AIR switch, or 3 and 4 ENGINE AIR switches with the STARboard CABIN AIR switch OPEN.

The port system is to be used on odd-numbered dates, the starboard on even-numbered ones.

c. At cruising altitude, all ENGINE AIR and CABIN AIR switches OPEN.

d. For roller landings, all ENGINE AIR switches SHUT.

25. If ram air Bow is required, set the PRESSURE SELECTOR to NO PRESSURE, the CABIN AIR switches OFF and the ram air switch to OPEN.

26. Air-conditioning is marginal when the OAT is near I 29°C. Before descent to low level use the cold air turbine to pre-cool the cabin.

Cold Air Turbine Overspeed 27. If the COLD AIR UNIT OVERSPEED WARNING indicator shows white, the pressure ratio switch has operated to stop overspeeding; the indica­tion should be only temporary. If the system is being operated in AUTO, a warmer setting should be selected to prevent the recurrence of overspeeding; in MANUAL no action should be necessary as the tem­perature control valve will back off and remain at a warmer setting. If overspeeding persists on the climb, close an ENGINE AIR switch and check that the indicator goes black; re-open the ENGINE AIR switch at altitude. If a colder setting is needed and cannot be obtained because of turbine overspeed, reduce the air supply to the turbine at cruising altitude by closing a CABIN AIR switch. When non-essentials are tripped, the overspeed magnetic indicator still functions although the temperature control valve cannot be moved. The overspeed can be terminated by selectively closing ENGINE AIR switches.

Loss of Cabin Pressure 28. If there is a serious leak in cabin pressure, or if, during the climb, the aircraft rate of climb exceeds the rate of pressurisation, the warning horn sounds.

Emergency Decompression 29. In an emergency, cabin pressure can be released by any of the following :

a. Rearward movement of the EMERGENCY DECOMPRESSION switch (1st pilot).

b. Rearward movement of either ABANDON AIRCRAFT switch.

c. Selection of NO P R E S S U R E on the PRESSURE SELECTOR (co-pilot). d. Operation ·of the CABIN PRESSURE RE­LEASE lever (nav /plotter).

Methods a. to c. are electrical and operate via a common fuse (3P 636). Method d. operates mechan­ically and electrically via a separate fuse (4P 562).

1-1 Page 6 RESTRICTED

• RESTRICTED AP lOlB-1902-15

30. To increase the rate of depressurisation, switch off the CABIN AIR switches at the same time as decompression action is taken. When re-pressurising the cabin, select COMBAT initially.

AIR-VENTILATED SUITS SYSTEM Normal Air-Ventilated Suits (AVS) System 31. The air-ventilated suits arc supplied from an air conditioning unit, similar to the cabin conditioning unit. The A VS unit uses hot air from the engines or AAPP, and cold air from the cabin system ram air intake. A ground-conditioning connection is provided, so that an external supply may be plugged into the suits.

A VS Air Conditioning Unit

32. The A VS air conditioning unit, in the nosewheel bay just aft of the cabin conditioning unit, comprises an air-to-air cooler, a turbine and fan, a flow aug­menter, a water extractor, a heat exchanger, and a filter. Hot air from the flood flow supply line passes, via an electrically-operated on-off cock, to the air-to­air cooler and then to the turbine and the water extractor. A branch line of the hot air bypasses the cooler and turbine and feeds into the water extractor via a temperature control valve, while a further line passes through another temperature control valve to the heat exchanger.

33. The temperature control valve in the hot-air line to the water extractor is controlled by a sensing element in the line between the water extractor and the heat exchanger. The temperature control valve in the hot-air line to the heat exchanger is controlled by a sensing element, set at 15°C, in the manifold inlet.

34. A tapping from the air from the cooler passes through a flow augmenter to the forward side of the water extractor, to provide additional pressure at altitude. This air, mixed with the air from the turbine, passes through the filter to the manifold in the cabin via a non-return valve.

A VS Components in Cabin 35. The temperature and pressure of the air in the manifold are regulated by a sensing element, which operates to regulate the cock supplying hot air to the conditioning unit, and a pressure relief valve and pressure controller. From the manifold, individual lines (each with an electric heater and a manual flow control valve) pass to the crew positions. The A VS hoses for the 6th and 7th seats are fed from the nav J radar's and AEO's A VS lines respectively.

Auxiliary A VS System 36. The auxiliary A VS system, for cooling, opera,tes by drawing cabin air through the crew's A VS. A 200V

· Air Conditioning

AC exhauster unit under the navigator's table is con­trolled by an A VS ON fOFF switch on the edge of the AEO's table. The pilots select auxiliary A VS by means of individual A VS CHANGEOVER COCK switches which operate changeover cocks under the navigator's table. Rear crew members have a separate line and flow cock. To prevent overheating, the e~hauster motor is not to be switched on unless at least one flow cock is open.

A VS Controls

37. a. The normal AVS system is controlled by an OPEN/CLOSE switch at the rear inboard edge of panel 7P, which controls the main cock in the hot air supply and makes power available to the indi­vidual suit heaters.

b. Temperature and flow controls are provided for each individual suit, as follows:

(1) 1st pilot and co-pilot. Rear end of port and starboard consoles. (2) Bomb-aimer. On starboard side adjacent to oxygen regulator. (3) Radar and AEO. On forward edge of the navigator's table (also control heating to 6th and 7th seat positions). (4) Plotter. Temperature control on port side and flow control on starboard side of seat.

c. Temperature is adjusted by a rotary control, consisting of two concentric knobs. The outer allows the systems to be operated automatically or manually and the inner selects the desired temper­ature, which is held automatically if AUTO is selected and the temperature selected is below 21 °C. This switch is also used as an INCREASE/DE­CREASE control if MANUAL is selected. The system is designed to maintain the temperature of the air supply between + 15°C and +46°C but, because of the length of ducting within the pressure cabin, the temperature at the suit inlet is affected by the cabin temperature and, if the cabin temperature is high, the minimum suit inlet temperature of + 15°C is not achieved.

d. The flow is controlled by an ONJOFF starwheel beside the -temperature control. The control is turned anti-clockwise from the OFF position to open the flow valves, flow increasing with increased movement of the control.

e. Overheat protection is provided for both the main air supply and the individual heaters. If the manifold temperature rises above 70°C, an overheat switch operates to shut down the air supply to the suits and to shut off the A VS beaters. Select the switch to CLOSE and allow the manifold to cool (60°C). Attempt to regain the system by selecting OPEN. If the system overheats again, select

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Page 7 (AL 10)

RESTRicrED

CLOSE and 'leave it off. A further thermostat control in each heater unit operates if the tem­perature in the heater duct becomes excessive.

38. Use of AAPP for AVS. The AAPP may be used to supply the normal AVS system on the ground, sub­ject to the limitations in Part 3. An AAPP BLEED FOR CABIN CONDITIONING SHUT jOPEN switch with adjacent MI are .on panel 7P.

BOMB BAY CONDITIONING

Bomb Bay Conditioning System

39. Twin dorsal intakes provide cold air through ducts via a cold air valve. An outlet louvre in the port bomb door exhausts t.he air to atmosphere. Controls at the navjradar's position consist of an OFF (centre)/ COLD (down) switch which controls the cold air valve, a temperature selector and a manual heat control (both inoperative) and a temperature gauge.

Bomb Bay Overheat Drill

40. a. Overheating is caused by leaks in the hot-air distribution trunking which runs through the bomb bay. The symptoms of overheat are:

(1) Temperature exceeding 30°C when the ambient temperature is 20°C or below.

(2) Temperature exceeding ambient + l0°C, when the ambient temperature is above 20°C.

(3) Possible displacement of the control columns in straight and level flight.

b. Actions

Note 1: Check each time if the temperature limits have been regained before taking the next action.

Note 2: If airframe anti-icing is ON, switching the system OFF (if practicable) may reduce the bomb bay temperature to within limits.

(1) Select the bomb bay cold air valve to COLD.

(2) Close two engine air switches (operationally, a third switch may be closed and COMBAT pressure selected.) If control deflection is noticed in the aileron sense, switch off No 3 and 4 engine air switches; if the deflection is in the elevator andjor rudder sense, switch off No 1 and 2 engine air switches.

(3) If aircraft limitations allow, open the bomb doors ; close them when the temperature is within limits. ( 4) If the bomb doors cannot be opened, or if the temperature rises again after the doors have been

closed, close all engine air switches. Continue the sortie below FL 250, or return to base.

WINDSCREEN THERMAL DEMISTING

Windscreen Demisting Supplies 41. To demist the inside of the windscreen, hot air is supplied from a slotted duct on each side of the centre panel. Cabin air is supplied to a blower motor, below the pilots' floor on the starboard side. From the blower motor, the air passes to a 1 kW heater unit and thence to the windscreens. An overheat switch in the system cuts off electrical supplies to the heater if the temperature of the air in the ducting rises above 70°C and switches supplies on again when the temperature faills to 60°C.

Windscreen Demisting Controls and Operation

42. Windscreen demisting is controlled by an ON I OFF switch on the co-pilot's instrument panel. When switched ON, current is supplied to the blower and the heater. The system should be tested before flight by switching on and physically checking the airflow from the ducts.

ELECTRICAL SUPPLmS

Electrical Supplies

43 . 200V AC. The windscreen demisting heater and motor are supplied from No 1 busbar and the A VS heaters from .No 3 busbar.

44. ll5V AC. The cabin conditioning magnetic amplifier uses 115V AC from the starboard main transformer.

45. 28V DC. The cabin conditioning, cabin pressuri­sation, bomb bay conditioning and windscreen demist­ing system use 28V DC for switching. The cabin con­ditioning and pressurisation warning systems also use 28V DC.

46. Load Shedding. If non-essential loads are shed, the following services are lost:

a. Cabin conditioning control.

b. Cold air turbine overspeed control ; indicator still operative.

c. The bomb bay cold air valve remains in the last selected position ; control is lost. Temperature indication goes to maximum deflection.

d. All A VS heaters.

e. Auxiliary A VS and face blowers.

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RESTRICTED AP lOlB-1902-IS

PART 1

CHAPTER 2-AIRCREW EQUIPMENT ASSEMBLIES AND OXYGEN SYSTEM

Contents

General ...

EJECTION SEATS Ejection Seats, General ... Ejection Gun and Firing Handles Canopy/Scat Connection Leg Restraint Drogue Gun Barostat/g-Switch Time Delay ... Manual Separation

REAR CREW SEATS Crew Seats, General Assistor Cushions

CLOTHING AND PERSONAL EQUIPMENT CONNECTORS Personal Equipment Connectors (PEC) Air-Ventilated Suits (A VS) Pressure Jerkins and Anti-g Suits (High Altitude) Masks and Helmets Low Altitude Clothing Assemblies Rear Crew Safety Equipment

OXYGEN SYSTEM Description of Oxygen System ... Oxygen Regulators, General Oxygen Regulators, Controls and Indicators Emergency Oxygen

USE OF AIRCREW EQUIPMENT ASSEMBLIES Clothing Assembly Strapping-In Procedure Normal Exit Procedure Pressurisation Failure ... Oxygen Failure Indication Regulator Failure

IDustrations Type 3KS2 Seat-Co-pilot Seat Details Oxygen System ...

Para 1

2 8 9

10 11 13 14

17 20

22 24 25 26 27 28

31 34 37 40

42 44 46 48 49 50

Fig 1 2 3

~WARNING: The aircraft is safe for parking when safety pins are inserted in both ejection seats and in the canopy as follows :

e. Seat pan firing handle. ~

a. Canopy jettison firing unit sear. b. Guillotine sear. c. Canopy jettison and time delay trip lever. d. Ejection gun sear.

It is emphasised that pins should not be inserted in the fabric straps above the pilots' heads.

General 1. The aircrew equipment assemblies comprise the seats, the flying and safety clothing and associated

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Page 1 (AL 10)

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~connectors. The pilots arc provided with ejection seats while rear crew members have sliding, bucket-type seats, the outer two of which swivel. There are two positions for extra crew members, forward of the nav J radar and AEO seats. There are oxygen and intercom connections at all seven crew stations and static line connections at the five rear crew stations. •

E)ECI'ION SEATS

Ejection Seats, General

2. The ejection seats, type 3KS1 for the 1st pilot and type 3KS2 for the co-pilot, are similar but partially handed. Each has a horseshoe-shaped para­chute pack, a back pad with adjustable kidney pad and a personal survival pack (PSP) type ZC. The seats have a nominal ground-level capability, provided that the

~ speed is at least 90 knots. Any deviation from straight and level flight at the instant of ejection reduces the seat performance. ~

3. The seat pan is adjustable for height by means of a lever on the outboard side of the seat. The trigger in the end of the lever is pressed in to adjust the height and, when released, loeb the seat in the selected position.

4. The adjustable armrests are controlled by either of two levers on each rest, one at the forward end and one at the rear, on the side of the rest.

5. A lean-forward lever, forward on the port side of each seat, allows the occupant to lean forward, by unlocking the attachment between the shoulders and the back of the seat. The straps wind in and out, following the pilot's motion and are locked on application of negative g and/or rapid acceleration.

6. A negative-g restraint strap is attached to the seat pan and is adjusted by a downward pull.

7. The sears have pressure-demand emergency oxygen (see para 40), a Mk 42 parachute and a static-line operated guillotine (para 14).

Ejection Gun and Firing Handles

8. Each seat is fitted with an 80 ftjsec telescopic ~ejection gun fired by either the face-screen B-shaped

firing handle above the occupant's head or the seat pan firing handle in the front of the seat pan. Either handle ~ must be pulled to its full extent to fire the gun. Safety

pins for each firing handle are stowed, when not in use, in a combined stowage for all safety pins one on each side of the cockpit.

Canopy /Seat Connection

9. An interconnection between the seat-firing mechanism and the cockpit canopy enables the canopy to be jettisoned automatically when any firing handle

~is pulled. Because of this interconnection, there is always a 1-second delay between pulling an ejection seat handle and the seat ejection gun firing, even if the canopy has already gone. ~

Leg Restraint

10. Leg restraint lines ensure that the legs are drawn back and held close to the seat pan during and after ejection. The lines pass through snubbing units below the front of the seat pan and are then fastened to the cockpit floor with rivets which shear at a pull of 400 lb. The snubbing units allow the lines to pass freely dawn through the unit but prevent them passing upwards, except when released by the spring-loaded toggle at the front of each snubbing unit. The leg restraint lines are unfastened when the man portion or cover of the PEC is removed from the seat portion.

~ Note: Before ejecting, place the feet on the rudder pedals. ~

Drogue Gun

11. The drogue gun has a time-delay mechanism and fires half a second after the ejection gun has fired, with­drawing the duplex drogues to stabilise the seat. The time-delay mechanism is operated by a static trip rod, which withdraws a sear from the gun as the seat rises on the rails.

Barostat/g-Switch Time Delay

12. Automatic separation is controlled by a time delay switch, which is inhibited by a barostat and a g-switch. The time delay runs for It seconds and is started by a static line as the seat moves up the rails, provided that:

a. The height is not greater than 10,000 feet.

b. The acceleration is not greater than about 4g.

13. If either condition is exceeded, the barostat andjor g-switch interrupt the running of the time switch until the conditions are satisfactory (seat below

1-2 Page 2 RESTRICTED

BAROSTAT TIME RELEASE

TIME RELEASE STATIC ROD ---------~

ARMREST ADJUSTMENT LEVERS ----~

LUMBAR PAD

STARBOARD LAP STRAP

PERSONAL EQUIPMENT CONNECTOR

SEAT PAN F'"IRING HANDLE

STARBOARD SIDE

'D' RINGS ON lAP STRAPS

EMERGENCY OXYGEN CONTROL

LEG RESTRAINT STRAPS

SEAT I ADJUST

1-2 Fig. 1 Type 3KS2Sea._

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lAPS

PS

AP 101 8 -1902-1 5 Aircrew Equipment Assemblies and Oxygen System

DROGUE CONTAINER

RESTRAINING

WAISTBELT

RIPCORD

PORT QUICK -RELEASE F" ITTING F"OR SURVIVAL PACK

PORT LAP STRAP

SURVIVAL PACK LANYARD

MANUAL SEPARATION LEVER

SEAT HEIGHT 'LEAN F"ORWARD ' ADJUSTMENT LEVER LEV ER

PORT SIDE

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r1 ( (

GUILLOTINE

SYSTEM

,I jl (

Ill le-1-+...J~L---+--- EMERGENCY OXYGEN 11 CYLINDER ,, )(

PORT RESTR

i'

SAFETY PI N

DEMAND EMERGENCY OXYGEN REGULATOR ---H,H.,L..IIIUJ

EMERGENCY OXYGEN

I 2

GUARD PLATE

YELLOW TAB ON BUCKLE OF NEGATIVE-G RESTRAINT STRAP (PULL DOWN TO LOOSEN)

RUNNING END OF NEGATIVE-G RESTRAINT STRAP (PULL DOWN TO TIGHTEN)

I 2 Fig. 2 Seat Details

I

~ Now refers to Mk. 4 seat onl

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.TE

GUILLOTINE

SYSTEM

EQUIPMENT R

YELLOW TAB ON BUCKLE OF NEGATIVE-G RESTRAINT STRAP (PULL DOWN TO LOOSEN)

RUNNING END OF NEGATIVE-G RESTRAINT STRAP (PULL DOWN TO TIGHTEN)

1- 2 Fig. 2 Seat Details w refers to Mk. 4 seal only ~

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LINK LINE

~--- SHORT STRAP

PARACHUTE WITHDRAWAL k.;ll-,l.#,~.JJ----- LINE/LINK LINE COUPLING

STRAPPING IN

GUILLOTINE PIN

ARACHUTE RIPCORD HANDLE

HARNESS QUICK­RELEASE FITTING

SURVIVAL PACK

LOWERING LINE

CONNECTED TO

LIFEJACKET OR PRESSURE JERKIN

RESTRICTED AP 101B-1902-15 Aircrew Equipment Assemblies and Oxygen System

10,000 feet, speed below 300 knots). When the time switch operates, the harness, leg restraint lines, per­sonal equipment connector, face blind and parachute pack are all released from the seat. Simultaneously, the drogues are detached from the seat but remain attached to the apex of the parachute canopy, with­drawing it upwards. The parachute subsequently develops and pulls the pilot out of the seat.

Note 1: When the seat pan handle is used to initiate ejection, the handle remains au·ached to the seat and cannot !be retained during separation. Conversely, the face-screen handle, if used, remains in the pilot's hands after separa·tion and should be discarded as soon as convenient.

Note 2: For flights over mountainous territory, ra 5000 metre capsule can be substituted for the normal 10,000 foot capsule.

Manual Separation

14. To allow the occupant to release himself from the seat, if the automatic devices fail to operate, means of manual separation are embodied.

15. A static-line operated guillotine is provided on the port side of the drogue box. A safety pin is provided for the gun sear and, when not in use, is stowed with the other safety pins on the cockpit coaming.

16. Separation is achieved by pulling out and up the manual separation lever to .the tear of the seat pan on the left side. This releases the harness locks, the parachute retaining straps, the leg restraint lines and the man portion of the PEC and, as the seat occupant falls clear of the seat, a static line attached to the rear of the parachute pack withdraws the sear from the guillotine and severs the drogue link line. After leaving the seat, the parachute is opened by pulling the D-handle on the waistbelt.

REAR CREW SEATS

Crew Seats, General

17. a. The navigator/radar and the AEO have seats which are capable of swivelling inwards, to face almost forward (aircraft sense). The navigator plotter's seat does not swivel.

b. Each seat has an assistor cushion to help the occupant rise from the seat in emergency. The backrests embody spring cHps to help retain the ·top of the parachute pack against the seat back ; they also embody clips for stowage of the shoulder straps.

A Mk 46 or Mk 49 parachute, including a demand and inflation emergency oxygen set, is used. All three seats can be slid fore-and-aft on rails.

18. Swivel Seats for Navigator J Radar and ABO

a. These seats have a lever to the left of the seat pan to control seat movement. In its spring-loaded central position, the seat is locked. When moved forward, the seat is unlocked for the rake of the seat back and swivelling. The rake is spring-loaded to the forward position to clear obstacles when swivelling; in normal use, the rear position is adopted. The control lever may be released once either motion has begun and automatically relocks when the full travel has been reached.

b. When the seat is locked in the fore-and-aft direction to face <the table, the same lever may be pulled aft against its spring, to free the seat feyr sliding. There are several finite positions in which the seat may be locked when the lever is released.

c. For use by persons not occupying the seat, a handle at the top of the back rest, when moved to <the right, duplicates the forward motion of the lever for swivelling and raking. Similarly, a toe-operated lever behind the base of the seat, when moved to the right, releases the seat for fore-and-aft travel (but, unlike the lever, is effective even if the seat is swivelled).

d. In order to clear the table, the seat must be at the forward (aircraft sense) end of its rails before swivelling.

e. Thigh supports at the front of these seats can be adjusted up or down as desired, by a central star wheel beneath them.

f. An S-type Mk 2 personal survival pack in the seat pan is attached to the outside of the parachute harness. It is provided with a lowering line and should be lowered from the harness during the parachute descent.

19. Sliding Seat for Navigator/Plotter

a. The lever permitting the seat to slide fore-and­aft is on the right of the seat pan in this seat.

b. An S-type Mk 2 personal survival pack is provided.

c. To prevent trapping the knees under the table, the seat must be at the forward (aircraft sense) end of its rails before operating the assistor cushion.

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Assistor Cushions

20. To assist the occupant to his feet, under condi­tions of positive g, each seat has an assistor cushion fitted in the seat pan which can be inflated by air at a pressure of 1200 PSI. A pressure gauge is fitted on the bottle. The air bottle is stowed on the back of the navigator/radar's and AEO's scats and in the back of the seat pan of the navigator/plotter's seat. Pins are provided for the assistor cushion bottles. To make the bottles operative, the navigator/plotter's pin must be removed and the other two must be inserted.

21. A knob on the right of the seat pan of the navigator/radar's and AEO's seats and a lever on the left of the navigator/plotter's seat pan, when pulled upwards to the full extent of their travel, release the air to inflate the cushion. T he initial inflation, in turn, releases the safety harness lap strap anchorages, freeing the occupant from the seat.

CLOTHING AND PERSONAL EQUIPMENT CONNECTORS

Personal Equipment Connectors (PEC)

22. Pilots

a. The pilots' PEC are in three portions, the air­craft portion, the seat portion and the man portion. The aircraft portion is attached to the underside of the seat portion, on the right-hand side of each seat. The man portion, an integral part of the clothing, is attached (during strapping-in) to the top of the seat portion. When not in use, the seat por·tion is protected by a dust cover, for which a stowage is provided on the back of the co-pilot's seat.

b. The PEC connects all the personal services to the man. The aircraft supplies (main system oxygen, ventilation air, micftel) are fed into the aircraft portion. As the seat ascends the guide rails on ejeccion, a static rod causes the seat portion to break away.

c. To prevent loss of emergency oxygen, the lower orifices of the seat portion are closed by valves when the aircraft portion is removed.

d. The man and seat portions are mated by sliding the nose of the man portion into the hooks at the front of the seat portion and then pressing the handle at the rear downwards. ~ ~

e. To release, press down on ·the thumb button in the handle and lift the handle. (This also releases the leg restraint lines.)

23. Rear Crew

a. The rear crew and spare seat positions are each provided willh an aircraft hose assembly consisting of an oxygen hose, micf,tel lead and a static line, all taped together in a protective sleeve. A switch and associated electrical wiring, connected to the static line, are also incorporated in each assembly to operate the 'crew gone' lights in front of the 1st pilot.

b. The assemblies are of sufficient length to enable the crew members to move about the cabin and to pass through the door in order to abandon the air­craft.

c. Separate A VS hoses are provided for the crew members, the spare seats and the sextant positions. Quick-release connectors are fitted.

Air Ventilated Suits (AVS)

~ 24. AVS Mk 2A (nylon) or Mk 2C (cotton) may be worn; the hose from the suit is passed through the~ clothing to connect with the aircraft supply. The pilots' supply is through their PEC while the rear crew have separate connectors. The supply system is described in Chapter 1.

Pressure Jerkins and Anti-G Suits (High Altitude)

25. Pressure jerkins (Mk 4 for pilots and Mk 3 for rear crew) are worn in conjunction with anti-g suits Mk 7B when flying at high altitudes. These two items form a pressure suit to protect the crew member if cabin pressure fails ; they inflate automatically if the cabin altitude reaches 40,000 feet. They are inflated from the oxygen system, both being connected to the pressure jerkin hose assembly. The jerkin connection is permanently made but the hose from the anti-g suit has to be threaded through the outer clothing and then attached. The pressure jerkin embodies a life jacket. The attached hose assemblies terminate in the man portion of the PEC for the pilots and a bayonet connector for the rear crew.

Masks and Helmets

~ 26. Either a G-type helmet with a Mk 1 protective helmet, or a Mk 2 or Mk 3 protective helmet, may be worn; in all cases either a P2C or Q2C oxygen mask is worn. These masks are of the chain-toggle,~ pressure-breathing type with a bayonet hose connection; they are identical apart from size. A toggle on the front of the harness is normally in the up position. The mask should be tested before flight and the knurled screws adjusted so that there is no leakage during operation of the press-to-test facility with the mask toggle in the down position. If cabin pressurisation failure occurs, the toggle is put to the down position, to clamp the mask more tightly on the

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• RESTRicrED AP 101B-1902-15 Aircrew Equipment Assemblies and Oxygen System

face for pressure breathing. When pressure clothing is worn, the mask is plugged into the top end of the jerkin hose assembly.

Low-Altitude Clothing Assemblies

27. For flights below 45,000 feet (cabin pressure), it is not essential to wear pressure clothing with the .Mk 21 regulator. In such cases and when using a Mk 17 regulator, normal flying clothing is worn, with a separate life jacket. To enable the aircraft connec­tions to be made, a special mask hose assembly is

~required, Mk 2 for pilots; Mk 7 for rear crew (Mk 2 post-Mod 2393). ~

Rear Crew Safety Equipment

28. The rear crew members wear back-type Mk 46 or ~ Mk 49 parachutes with type S .Mk 2 personal survival ~

packs (PSP). When the static line is used, a barostat control delays deployment of the parachute until below 13,000 feet. This can be overriden by a handle on a strap between the legs.

29. A demand and inflation emergency oxygen set is provided. The cylinder and opening head are in the top of the parachute pack and the regulator is on the right-hand waist-belt. The operating handle is on the strap between the legs.

30. Full details of the rear crew emergency oxygen system are given in para 41.

OXYGEN SYSTEM

Description of Oxygen System

31. Oxygen is carried in 12 X 2250 litre bottles. On early aircraft, all the oxygen bottles are housed in the power compartment, aft of the bomb bay. In later aircraft, four bottles are housed in the power compart­ment and the remainder in !the bomb bay. The bottles are all charged through a connection in the power compartment; the correct charging pressure is 1800 PSI. Two pressure gauges at the AEO's position show the pressure in each half of the system.

32. From the oxygen bottles, the high pressure supply lines pass along the sides of the bomb bay and into the pressure cabin. .Master valves, one for each side of the system, are below the crew's flooring ; these valves are normally wire-locked to the open position. From the master valves, the supply lines pass along the cabin and are interconnected by transverse lines at four points.

33. The transverse connections are protected by non­return valves so that, if there is a leak on one side of the system; oxygen is not lost from the other side.

From the transverse lines, the supply is fed to pressure­reducing valves, one for each regulator, which reduce the pressure to 400 PSI. The medium pressure lines pass from each pressure-reducing valve to the regu­lators. From the regulators, oxygen at breathing pressure is fed to the PEC on demand.

Oxygen Regulators, General

34. a. The oxygen regulators may be Mk 21A, 21B or 17F, one being supplied for each crew mem­ber. The 1st pilot's and co-pilot's regulators are at the forward end of the port and starboard consoles, while the rear crew regulators are at their respective stations.

b. The Mk 17F and 21 series regulators have the same characteristics up to 39,000 feet cabin altitude. Above this altitude, the Mk 21 series automatically delivers a higher pressure than the .Mk 17F. The Mk 2 emergency regulators have the same characteristics as the Mk 21 series.

c. In all cases of cabin pressure failure, an irnme-~ diate descent i:s to be made until cabin altitude is

below 40,000 feet. When flight safety and fuel con­siderations allow, the descent should be continued, at normal rate, to below 25,000 feet. ~

35. Regulator Chm·acteristics. The regulators provide :

a. A mixture of oxygen and air, or 100% oxygen; the flow and volume delivered is in direct relation to the breathing demands of the user.

b. The correct ratio of air and oxygen according to cabin altitude. Above 33,700 feet, 100% oxygen is provided automatically. 100% oxygen may be selected at any height.

c. A safety pressure ; the mask cavity pressure is slightly higher than cabin pressure when cabin altitude reaches 11,000 to 14,000 feet and pressure breathing when the cabin altitude exceeds 39,000 feet. Additionally, the Mk 21 series regulators inflate pressure clothing at the same altitude as they give pressure breathing.

d. Positive oxygen pressure by manual selection :

(1) Emergency use (100% and EMERGENCY), ie, for toxic fumes .

(2) Mask and regulator testing (Mk 17F).

(3) Mask, regulator and pressure clothing testing CMk 21 series).

36. Regulator Limitations. The protection given by the aircrew equipment, in terms of cabin altitude, is shown in Table 1.

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Page 7 (AL 9)

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Table 1 Regulator Limitations

Pressure

l Anti-g Max Cabin

Remarks Regulator 1erkin Suit Alt (ft)

Mk 17 F ... . .. No No 50,000 -Mk 21A orB ... Yes Yes 56,000 Oxygen contents more than three-

eighths of total

Mk 21A orB ... Yes No 52,000 -

Mk 21A orB ... No No 45,000 Above this altitude, oxygen pressure is Mk 2 or 2A above 30 mm Hg (max lungs can

stand in comfort)

.Mk 3 or 3A . . . ... No No 50,000 -Mk 2 or 2A .. . ... Yes Yes 56,000 -

Note: ,fu the worst case, eg following loss of the canopy, or if the entrance door opens, aerodynamic suction can cause the cabin altitude to exceed the aircraft a1titude by up to 5000 feet.

Oxygen Regulators, Controls and Indicators

37. Mk 21 Regulator Controls

a. OXYGEN SUPPLY, ON jOFF Lever (wire­locked at ON). This lever controls the supply of oxygen to the regulator and must be ON at all times in flight.

b. NORMAL OXYGEN /100% OXYGEN Levet·. When in the normal position, this lever allows air to mix with the oxygen in suitable proportions, up to an altitude of 33,700 feet. Above this altitude, the air inlet is closed and 100% oxygen is delivered to the mask. With the lever at 100% OXYGEN, the air inlet is closed regardless of the altitude ; this position should always be used if toxic fumes are present.

c. JERKIN TEST / MASK TEST / EMER­GENCY /NORMAL Lever. When this lever is set to NORMAL, oxygen or a mixture of oxygen and air as selected by b. is fed to the mask at the required pressure when the user breathes in. When the lever is set to EMERGENCY, the pressure of the oxygen to ·the mask is slightly increased above the delivery pressure appropriate to the altitude. This increase is constant at all altitudes. The EMERGENCY position should be used if toxic fumes are present in the cockpit. To reach the MASK TEST position, the knob 1n the end of the lever must be pulled out ; in the MASK TEST position, the mask can be tested under pressure for leaks. The JERKIN TEST position is similar in operation to the MASK TEST position but gives a higher pressure and is used to test mask, jerkin, and g-suit simultaneously for leaks.

38. Mk 21 Series Regulator Indicators

a. A gauge on the regulator shows the pressure of oxygen being delivered to the regulator, while two gauges at the ABO's station show the main pressures.

b. A magnetic indicator on each regulator shows white when ·the user is breathing in. The pilots' and three rear crew members' indicators are duplicated on their respe<:tive panels, and the bomb aimer's is also duplicated, on a rear central support for the pilots' floor, where it can be monitored by other crew members.

39. Mk 17F Regulator Controls and Indicators. The .Mk 17F regulator carries a pressure gauge, a magnetic flow indicator, a NORMAL/100% selector, an ON/ OFF control and an EMERGENCY lever. This last control is pressed in to test the mask for leaks and deflected to either side to obtain oxygen at higher pressure in an emergency.

Emergency Oxygen

40. Pilots' Emergency Oxygen

a. Pilots can obtain emergency oxygen automatically on ejection or by pulling up a yellow and black knob inboard of the ejection seat on the forward comer of the seat pan.

b. As the bottle is attached to the seat, emergency oxygen is not available after separation from the seat.

c. A pressure-demand emergency oxygen set is fitted to each ejection seat, the cylinder and operat­ing head on the back and the regulator Mk 2 or 3 behind the PEC. The operating head initiates the action and delivers medium pressure oxygen to '!!he regulator, which then delivers it to the user on

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•

NAVIRAOAR

6th SEAT

CO-PILOT

NAY/PLOTTER

BOMB AlMER

RESTRICTED

MASTER VALVE

AEO

7th SEAT

1st PILOT

AP 101B-1902-15 Airc.rew Equipment Assemblies and Oxygen System

--. HIGH PRESSURE FROM BOTTLES

MEDIUM PRESSURE TO R.EGULATORS

LOW PRESSURE TO MASKS

PRESSURE REDUCING VALVE 1,600-400 PSI

liON-RETURN VALVE

IQ] REGULATOR

(}) PRESSURE fiAUGE

1-2 Fig 3 Oxygen System

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demand. Mk 2 and Mk 3 regulators have similar characteristics to the Mk 21 and Mk 17 series regulatot\S respectively, except thart safety pressure is delivered from ground level upwards. The cylinder contents are indicated on a pressure gauge. The endurance of the set is approximately 10 minutes provided that the mask is fitted correctly and that there are no leaks in the hose assembly.

41. Rear Crew Emergency Oxygen

a. The rear crew are provided with Mk 2A and 3A regulators which are identical in performance to the pilots', only the fitting and method of operation being different. An emergency oxygen bottle in the top of each rear crew parachute pack is turned on automatically when the parachute static line is operated. It can also be operated manually by a control on a strap between the user's legs. ~ ~

b. The pressure demand regulator is on the right­hand waist-belt and a contents gauge is on the bottle in the pack. The bottle is filled to capacity when the gauge needle is on the white line between the sectors marked FULL and REFILL.

USE OF AIRCREW EQUIPMENT ASSEMBLIES

Clothing Assembly 42. When wearing pressure clothing, the air ventilated suit and anti-g suit are put on beneath the shirt and trousers and the jerkin is put on top of the flying overalls. The hoses from the anti-g suit and A VS are fed through the outer clothing; the A VS hose is connected to the PEC for the pilots and to an individual connector for the rear crew, the anti-g hose is connected to the main oxygen hose. The oxygen hose from the mask is subsequently attached to the bayonet socket at the top of the main oxygen hose.

43. When pressure clothing is not worn, the mask hose assembly is used.

Strapping-in Procedure

44. Pilots. Ensure that the seat has been made safe for parking and that harness straps have been fully

~ extended. Prior to entering the seat, carry out the safety equipment checks given in the FRC. ~

a. Before flight, ensure that the ejection seat safety pins are correctly stowed, and the canopy safety pin is repositioned, as detailed in FRC.

b. Connect the lanyard of the personal survival pack to the pressure jerkin or life jacket. The line should be across the left thigh. c. Remove the PEC cover and connect the man portion to the seat portion, checking that it is

correctly locked. With low-altitude clothing, con­nect the oxygen hose clip to the D-ring on the life jacket.

d. Pass the left leg-restraining line through the D-ring on the right garter and insert the plug on the end of the line into the housing on the left snubbing unit. Do up the right line in the same manner, ensuring that the lines are not interlaced. Adjust the lines by pulling out the toggles in the front of the units and moviug the lines as required, ensuring that full rudder can be applied.

Note: The sockets will not grip the line plugs unless the man portion of the PEC has first been locked in position.

e. Bring the harness waist-belt across the body and adjust the quick-release fastener so that it lies centrally with the waist-belt close to the body. Ensure that the quick-release fastener is in the locked position.

f. Bring the negative-g strap up between tl1e legs ensuring that it is to the rear of the seat pan firing handle and not passing through it. Thread the lugs of the lap straps through their respective loops in the blue Y-piece of the negative-g strap, and con­nect the lugs into ·their respective positions in the quick-release fastener on the harness waist-belt. Ensure that the hoses of the PEC pass under the right lap strap and that the quick-release fastener is as low as possible, consistent with comfort. Obtain the assistance of the crew chief to adjust the back pad and lumbar cushion to the most comfortable position.

g. When connecting the harness lugs to the inertia proof quick-release fastener, turn the disc knob anti­clockwise to the yellow dots on the body ; hold the disc knob in this position and insert the first lug. Repeat the operation as each of the other lugs is inserted.

h. Tighten the lap straps by pulling on the running end with one hand and pushing the standing end towards the buckle with the other. Tighten the negative-g strap by pulling downwards on the free end of the blue strap. Move the body about inside the harness and then retighten the lap straps and the negative-g strap. Repeat the operations until the straps are as tight as possible. It is most important that the lap straps and the negative-g straps are as tight as possible.

i. Thread the crutch loops through the D-rings on the lap straps, ensuring that the loops are so twisted that they lie Bat against the thighs. Pass the right shoulder strap through the right crutch loop and the left strap through the left loop and attach the

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• RESTRICTED AP 101B-1902-15 Aircrew Equipment Assemblies and Oxygen System

straps to the quick-release fastener, pulling the loops down to rest against the quick-release fastener.

j. Sit up straight and tighten the inner (blue) shoulder straps, followed by the outer (khaki) ones. These straps should not be tightened to such an extent that the back is bent, as this may cause a hazard on ejection.

k. Operate the lean-forward lever, lean forward, return the lever to the locked posi~ion and lean back. Check that the harness springs back and locks. This provides a check that the manual override lever has not been inadvertently operated. Retighten the khaki shoulder straps if necessary.

1. Obtain the assistance of the crew chief to pull back the lift webs through the metal runners on the shoulders and then stow the excess length neatly (by lengthening the loops in the lift webs) behind the back.

m. Ensure there is no slack in the harness straps of the PSP.

n. Check the movement of me seat through its full range and ensure that the PEC remains connected. Re-adjust seat height and reach upwards to check that the face screen firing handle is within reach.

o. Put on the helmet and connect the mask tube to the bayonet socket in the hose assembly. Plug in the micftel lead.

p. Check the intercom and oxygen regulator.

q. With assistance from the crew chief, ensure ,that the various safety pins are correctly repositioned and stowed, including the large pip-pin required for canopy sear withdrawal.

45. Rear Crew Members

a. Check the contents of the assistor cushion bottle. The pressure should be 1200 PSI. Navjplotter removes inflation bag safety pin at 'his discretion; AEO and nav /radar connect operating cable to bottle.

b. Check that the demand emergency oxygen pin has been removed.

c. Slacken all the parachute straps fully and stow them in their stowages before getting in the seat.

d. Check that the PSP is correctly connected to the parachute harness.

e. Connect the personal survival pack to the life­jacket.

f. Connect the oxygen mask hose assembly 1:0 the aircraft assembly and make the micjtel connection.

g. Connect the static line to the aircraft hose assembly. It should remain connected at all times except when the crew member moves about the cabin.

h. Connect the emergency oxygen supply to the oxygen mask hose assembly.

i. Adjust the parachute quick-release fastener in front of the body. Pass the leg straps through the crutch loops and connect them to the quick-release fastener, ensuring that the hose assembly is beneath the right leg strap.

j. Connect the shoulder straps to the quick-release fastener. Tighten the harness.

k. Fasten the seat lap strap.

I. Check the seat swivelling and sliding actions and make sure that the personal survival pack lowering line does not foul on any portion of the seat.

m. Connect the A VS to its separate supply point.

n. Put on the helmet and connect the mask tube to the mask hose assembly.

o. Connect the mic/tel lead. p. Check the intercom and oxygen.

Normal Exit Procedure

46. Pilots a. Before leaving the aircraft, the seats and canopy must 'be made safe for parking by inserting the seat

~ pins through the sear of the ejection gun, the canopy jettison and time delay trip lever, the guillotine sear~ and in the seat pan handle. The seat pan handle locking pin cannot be inserted until the handle is fully home. To make the canopy ejection gun safe, remove the pip-pin from the operating linkage and insert the attached safety pin in the gun-firing sear.

b. Remove helmet and mask, disconnecting mask tube and micftellead.

c. Release the harness.

d. Disconnect life raft lanyard.

c. Disconnect PEC; this also releases the leg restraint lines.

f. Move out of the seat, seeing that the leg restraint lines pass through the garter D-rings.

47. Rear Crew a. Remove the helmet and mask, disconnecting the mask tube and mic/tellead.

b. Disconnect the safety harness and parachute harness.

c. Disconnect the emergency oxygen supply.

d. Slacken the parachute straps and stow them in their stowages.

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Page 11 (AL 9)

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e. Disconnect the life raft lanyard.

I. Disconnect the static line.

g. Disconnect the mask hose assembly from the aircraft hose assembly, placing the aircraft assembly in its stowage.

h. Navjplotter replaces inflation bag safety pin if required; AEO and nav /radar disconnect operating cables.

Pressurisation Failure

48. If the cabin pressure fails, each crew member should immediately depress the toggle on the mask harness ; this clamps the mask tightly against the face for pressure breathing and prevents oxygen leaks. If the cabin altitude is above 40,000 feet, the instructions in paragraph 34c. must be followed immediately. At cabin altitudes above 39,000 feet, pressure clothing inflates and oxygen is delivered under pressure. Although when below 40,000 feet the immediate adverse physiological effects of pressure breathing are past, the risk of decompression sickness remains high until a descent through 25,000 feet has been made.

Oxygen Failure Indication

49. a. If both the regulator and remote flow indicators cease to operate, check that the main

oxygen tube is correctly connected and check or set the air inlet switch to 100% OXYGEN. If it is impossible to breathe in freely, the regulator is faulty.

b. If breathing is normal, check that the pressure on the regulator gauge is normal and that the main oxygen pressure gauges indicate that oxygen is still available.

c. If the above indications are satisfactory, momentarily select EMERGENCY. A supply of oxygen under increased pressure indicates that the regulator is serviceable but the indicator is defective.

Regulator Failure

SO. Pilots. If a pilot's regulator fails above 10,000 feet cabin altitude, descend to 8000 feet cabin altitude immediately. If 100% oxygen was selected, select NORMAL so that the cabin air can be breathed freely while maintaining R/T and intercom. If a descent is not practicable at that moment, the pilot must select emergency oxygen by pulling up the handle beside his seat and switching off the regulator ; the endurance of the emergency oxygen is only ten minutes.

51. Crew Members. If a crew member's regulator fails, he can transfer to a spare seat hose assembly, if the seat is unoccupied ; this is operationally preferable to using emergency oxygen, as it avoids the necessity of a descent.

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PARTl

CHAPTER 3-ARMAMENT AND OPERATIONAL EQUIPMENT (Completely revised by AL 10)

Contents

General

General

Bomb Release and Jettison

Window La1mchers

Camera Installation

1. There are two methods of bomb aiming in the aircraft, the NBS and the co-pilot's visual sight. The location of the NBS equipment is described in Chapter 14. Bomb bay conditioning is described in Chapter 1 and bomb door sealing in Chapter 13. Bomb door operation is described in Ghapter 10.

Bomb Release and Jettison

2. Release. Bombs are released automa·tically by the NBS release pulse or manually by a bomb release button. The first pilot ha's a bomb release button to the left rear of the throttle quadrant. The co-"pilot's bomb release button, stowed in a clip below the coaming, operates through a time delay unit at the navigator/plotter's position. The navigator/radar's bomb release button is stowed in a clip adjacent to panel 9P. There is a further bomb release button at the prone 1bomb aimer's position.

WARNING I: If one or more 1000 lb bombs have fallen on to or against the bomb doors, the doors must be closed (if open) and must not be opened again in flight, except in emergency. If they are opened, a bomb may be trapped against t'he <bomb bay wall, damaging rlle flying controls and/or the hydraulic lines.

WARNING 2: Failure to switch on the time delay unit at the navigator's/plotter's station when the co­pilot's release button is to be used results in a failure to release.

3. Jettison

a. The 'bombs may be jettisoned live by selection of the LIVE JETTISON switches on the navigator f radar's side panel and on the visual bomb aimer's oxygen panel.

b. An EMERGENCY BOMB JETTISON switch on the 1st pilot's console panel enables the bombs to be jettisoned safe throuSJb a wiring system separ­ate from the normal 'bomb release system. The switch is a guarded, double-pole switch labelled JETTISON/OVERRIDE, spring-loaded to the off

Para

1

2

6

7

position. When the switch is held momentarily to the JETTISON (rear) position, the bomb doors open, the <bombs fall and the doors close. Movement to OVERRIDE cancels jettison, provided that the bomb doors have not reached the fully open position. The selection can also be used to close the bomb doors before the time sequence unit does so. The bomb door.s dose when the time delay unit on the cabin wall at the AEO's position completes the manually preset 24-second cycle.

c. If hydraulic power has failed, the bomb doors must first be opened by the emergency sysrem before >attempting to jettison.

d. When nuclear stores and practice bombs are carried, the jettison circuits are rendered inoperative.

4. Bomb Release Safety Lock. A bomb release safety lock (BRSL), to prevent inadvertent weapon release, is controlled by oa guarded wire~locked switch on the port console marked LOCK/UNLOCK. An amber light comes on if the safety iock is released; a green light shows the <loCk is engaged. Press-to-test facilities on both lights are used to check a duplicate electrical circuit.

S. A SFOM bombsight is fitted. The sight is a collimated fixed-angle sighting head, on the coaming in front of tJhe co-pilot. A lighting control is adjacent to the other cockpit lighting controls. The sight line is fixed relative to the aircraft in azimuth but the depression angle can be adjusted to allow for speed, AUW, 'height and .relative air density. Tthere are two scales on the sighting head, one calibrated in degrees and the ot'her in milliradians. In flight, the milliradian setting knob should be used for adjustments to vhe sighting head. One complete revolution of the milli­radian scale adjusts the sighting head by 10°. The optical glass is stowed in a container on the starboard console and must 1be inserted in the sighting bead prior to bombing. Care must be taken to avoid scratching the glass while handling.

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Window Launchers

6. The gravity/ cartridge window installations are operated by the ABO. Three gravity window containers (two pott, one starboard) and one cartridge discharger are fitted.

Camera Installation

7. An F95 Mk 9 camera on a frame above tlhe bomb aimer's window is controlled from the navigator /radar's

~ position. The film can be marked to show bomb release automatically by the NBS release pulse or manually by the event marker button. (Command modi­fication 0291/Mod 2505 transfers the event maJ.'Iker supply to fuse 569 from fuse 970 in the weapons release circuit to eliminate the pos~~bility of the btter fuse rupturing due to a defective camera.) The event marker facility is not incorporated in t1be F95 Mk 4 camera fitted to MRR aircraft post-SEM 045 (Mod 2502). ~

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•

RESTRICTED AP IOIB-1902-15

PART 1

CHAPTER 4-ELECTRICAL SYSTEM

Contents

Introduction

200 VOLT AC SUPPLIES Alternators Alternator Controls and Operation CSDU Tolerances Ram Air Turbine (RAT) Airborne Auxiliary Power Plant (AAPP) AAPP Controls and Operation . .. Alternator Failure

SECONDARY POWER SUPPLIES Secondary Power Supplies, General 115 Volt, 3-Phase, 400 Hz 115 Volt, Single-Phase, 1600 Hz 28 Volt DC Failure of 28 Volt TRU 112 Vdlt DC

EXTERNAL POWER SUPPLIES 200 Volt AC Ground Power Unit 28 Volt DC Ground Supply

Tables

Para 1

7 12 18 19 23 28 32

38 40 43 45 53 56

57 61

Main Electrical Supplies-200 Volt, 3-Phase, 400 Hz 115 Volt Supplies

Table 1 2 3 4

DC Supplies Vital, Essential and Non-Essential Loads

Introduction

Illustrations Electrical System Main Control and Distribution Panels ...

1. The main electrical system is AC operated; 200 volt, 3-phase, 400 Hz power is supplied by four 40 kVA alternators, one on each engine. Each alternator is driven from its appropriate engine via a hydro­mechanical constant-speed unit and a frequency control circuit which maintains the a1ternator speed at 6000 RPM ± 1% under steady state input speed and load conditions.

2. From each alternator, current is fed to an individual busbar which can be connected to a syn­chronising ring main busbar for load sharing purposes.

3. From the individual busbars, a number of trans­former rectifier units (TRU), transformers and fre­quency changers provide the secondary power supplies.

Fig 1 2

4. Provision is made for standby supplies in the event of a major AC failure. A ram air turbine (RAT) supplies power primarily at high altitude and an air­borne auxiliary power plant (AAPP) supplies power at lower altitudes.

5. The distribution feeders are triplicated as a safety measure and all circuits except the ground supply are

~fused. No distribution fuses are accessible in flight. ~

6. The main controls for the electrical system are grouped on the alternator control panel (lOP), the secondary supplies panel (50P) and the AAPP panel (70P), all at the AEO's station. The controls for the frequency changers are on panel 12P at the navigator/ radar's station .

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200 VOLT AC SUPPLIES

Alternators

7. Each alternator provides 3-phase power at 200 vO'lts, 400 Hz. Automatic protection is provided against:

a. Over-voltage.

b. Under-voltage or incorrect phase sequence.

c. Line-to-line and line-to-earth faults up to but not including the busbars.

d. Underspeed.

e. Overspeed.

8. Alternator supplies are controlled by two circuit breakers. With the alternator (A) breaker closed, the alternator feeds its own busbar. With the synchronis­ing (S) breaker and the A breaker closed, the alternator supplies its own busbar and the synchronising busbar. The A and S breakers are so arranged that, if the A breaker is opened~ the S breaker is normally closed, thus ensuring an alternative power supply to the individual busbar via the synchronising busbar. The S breaker normally opens before the A breaker is closed ; S breakers can be closed in order to parallel two or more alternators on the synchronising busbar.

9. When two alternators or more are connected in parallel, real load sharing (distribution of load between alternators) is controlled by magnetic amplifiers, which adjust the torque of the constant speed unit. Reactive load sharing is controlled by automatic adjustment of the voltage regulator settings. The load sharing limitations are -+- 3 kW (real) ± 21 kVAR (reactive) of the mean values.

10. The alternators are cooled by ram air ducted into the front end of the alternator and exhausted to atmosphere at the ,rear. Provision is made for induced cooling using bleed air from each engine to augment the air flow when the undercarriage is selected DOWN. If the undercarriage is selected down using emergency air, this induced cooling is not available.

11. Each load busbar supplies approximately a quarter of the total loads and the PFC loads are divided between the four busbars. A list of loads supplied by each main busbar is given in Table 1.

Alternator Controls and Operation

WARNING : In flight, tthe alternators must not be run ~in parallel, except as directed by the Flight Reference

Cards. ~

12. The alternator control panel (lOP) is at the AEO's station and carries the following controls and indicators.

a. A voltmeter and frequency meter for the selected INCOMING alternator.

b. RAT and AAPP test push buttons, used to obtain the readings for these supplies on the meters in a. above.

c. An alternator selector switch, incorporating a pushbutton to facilitate synchronisation of alterna­tors.

d. An EXTRA SUPPLIES TRIP pushbutton, used to trip any extra supply (RAT, AAPP, 200 vo'lt ground supply) from the synchronising busbar.

e. A mimic diagram of the 200 volt system. The diagram incorporates a voltmeter and a frequency meter to show supplies at the synchronising busbar, magnetic indicators which show continuity when an S breaker is closed and amber lights to show when an alternator is not connected to its own busbar. Magnetic indicators for the RAT and AAPP show continuity when they are connected to the synchron­ising busbar. Placed centrally on lOP is a red alternator failure warning light (duplicated on the pilots' centre panel) which shines steadily if one alternator fails and flashes if two or more fail. Beside the AAPP magnetic indicator is an AAPP ON pushbutton, beside each S breaker indicator is an alternator ISOLATE button and beside each amber light is an alternator RESET button.

f. A NON-ESSENTIAL SUPPLIES, TRIP/ RESET switch, spring-loaded to the central (guarded) position, is on panel lOP. This switch can be used to trip non-essential supplies without releasing the RAT and to reset non-essential supplies once power has been restored.

g. Four kW /kVAR meters, one for each alternator, normally read kW; a button, placed centrally, is labelled PUSH FOR KV AR.

h. Four ON/OFF switches, one for each alternator.

13. CSDU oil temperature gauges are provided at the ABO's station. The maximum permitted load for each alternator is 32 kW, provided that the CSDU oil tem­perature remains below 120°C. This oil temperature must not be exceeded and, if necessary, height or load· ing must be reduced to keep the oil temperature within limits.

14. When an engine is running, its alternator is brought on line as follows :

a. Check, by means of the alternator selector switch, that the alternator voltage and frequency are within 115 -+- 5 volts, 400 -+- 4Hz.

b. When the alternator is running correctly, set its alternator switch ON. Check that the S breaker opens, the A breaker closes (amber light out) and the kW /kVAR meter registers the load, indicating that the alternator is feeding its busbar.

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28VGROUND SUPPLY PLUG REFUELLING

-o-_/ STANDBY A/H TRANSFORMER POST MOD.2182

GROUND I. SUPPLY RELAY

No.2 TRANS. DOPPLER

ALTERNATORS 200V -400Hz

81-8----

FORWARD FACE OF BOMB BAY r----., 28P

.-------- -l-14 ~ I TO SCANNER

TRU

TO PANEL liP FUSES

8 -81-

RESTRICTED

ALTERNATORS 2oov •ooHz

8 8

I No. 4 TRANSFORMER

ISTBD

__ j

)UPLE CONTACTOR

~ oov IISV I~NSFORMER

'PORT

I

I

-- J

8 8 No.I

RES1 RICTI:.D

8 \

,..----__...·

1 os.ll8 ELEVONS

8

MAIN AIRBRAKE MOTOR Not4lS ELEVONS MAIN RUDDER •

No I'A' ~ I Nos.J & 6 •eRE.AKER ELEVONS

AUX. AIR BRAKE

MOTOR (PORT) AUX. RUDDER

AP 1018·1902-15 Electrical System

GPU BREAKER

1 o.2 'A' BREAKER n t;\

2S V 200V ROUND UPPLY LUGS

G 5 p

- 200V 400Hz AC

- IISV AC

- 28V DC

I 4 Page 3 (AU)

PANEL

3P 4P 9P 11P 15P 16P 19P 24P 25P 26P 27P 28P 29P 48P 58P 59P 60P 61P 62P 69P 71P 72P 74P 75P

14J 15J 16J 17J

1- 4 Pag~: 4

11!1 V, 1e00 C.P.S. CONTROL PANEL

STANDBY A/H TRANSFORMER POST-MOD. 2182

FREQUENCY CHANGERS 200V . .ol00 -3 PHASE TO 11!1V 1e0o-SINGLE PHASE.

I N.B.S. TRANSFORMER 11!1'1.

FUNCTION

28V. FUSE AND RELAY PANEL. (PORT) 28V. FUSE AND RELAY PANEL. (STBDl BOMBING CONTROL PANEL. 115 V. 1600 C.P.S. & 400 C.P.S. RADAR SUPPLIES PANEL. 28 V.D.C. DISTRIBUTION PANEL. 28 V.D.C. DISTRIBUTION PANEL. 28 V.D. C. VITAL SUPPLIES DISTRIBUTION PANEL . 115V. 400 C. P. S. DISTRIBUTION PANEL. ( FED FROM 27 Pl 115V 400 C.P.S. DISTRIBUTION PANEL. (FED FROM 28P) 28 V. D.C. REAR SPAR DISTRIBUTION PANEL 200 V. & 115V. DISTRIBUTION PANEL (PORT) 200 V. & 115V. DISTRIBUTION PANEL (PORT) A .A .P.P. & RAT FUSE & RELAY PANEL. 28 V.D.C . DISTRIBUTION PANEL. No.1 ALTERNATOR 200V. DISTRIBUTION PANEL. No.2 ALTERNATOR 200V. DISTRIBUTION PANEL. No. 3 ALTERNATOR 200V. DISTRIBUTION PANEL. No. 4 ALTERNATOR 200V. DISTRIBUTION PANEL. A .A .P.P. FUSE & RELAY PANEL. VENTILATED SUITS FUSE PANEL . M .F.S. POWER PANEL (28 V.D.C. FED FROM 3P &. 115 V.A .C . FROM 24P) M.F.S. POWER PANEL (28V. D.C.FED FROM 4P &115V.A .C. FROM 25Pl A .A . P.P. & R.A.T. RELAY PANEL. 200V. SUB- DISTRIBUTION PANEL (FED FROM 28 Pl

No. 1 ALTERNATOR LOAD SHARING JUNCTION BOX. No. 2 ALTERNATOR LOAD SHARING JUNCTION BOX. No. 3 ALTERNATOR LOAD SHARING JUNCTION BOX. No. 4 ALTER NAT OR LOAD SHARING JUNCTION BOX.

1- 4 Fig. 2 ~

\ A.A.P.P. CONTROL SWITCH PANEL. --

SECONDARY SUPPLIES SWITCH PANEL. -

Main Control and Distribution Panels Motl/~e111~ ~

RfSTRICTED

AUXILIARY POWER PLANT

d Distl'ibuf ~ ~ 10" Panels

::o

RESTRICTED AP 101B-1902-15 Electrical System

15. Whenever there is no supply on the synchronising busbar, it is arranged that No 2 alternator S breaker closes automatically.

16. To connect an alternator to the synchronising busbar, select the appropriate alternator on the selector switch and then press in the switch until the magnetic

~indicator shows continuity. When available, No 2 alternator should be used as the medium for switching supplies to the synchronising busbar. Failure to do so can, under cert>ain circumstances, cause involuntary and simultaneous connection of No 2 a:lternator and the selected alternator and could lead to failure of both. ~ To take an alternator off the synchronising busbar, press its ISOLATE button.

17. Before flight, the synchronisation and isolation, together with the real and reactive load sharing, of all four alternators must be checked. For take-off, the AAPP is connected to the synchronising busbar and is normally closed down at 20,000 feet. If the climb is continued, No 4 alternator is connected to the syn­chronising busbar but, for low level flight or descent below 20,000 feet, No 3 alternator is substituted for No 4 on the synchronising busbar. On airfield recovery, the AAPP is connected to the synchronising busbar at 5000 feet.

CSDU Tolerances 18. Although, in steady flight conditions, alternator frequences must be within 400 + 4 Hz and the real load sharing of alternators in parallel must be within + 3 kW, transient deviations are permissible in the circumstances quoted below :

a. Change of Engine Speed. During engine acceleration and deceleration (including taxying and ground running conditions) frequency deviations of up to ± 30 Hz are pern1issiblc. With alternators paralleled, these may also cause real load swings outside the normal load sharing limits. Recovery to normal conditions should be accomplished within 2 seconds of completing the engine speed change. b. Load Switching. When load switching, fre­quency deviations of up to ± 30 Hz are acceptable. Recovery to normal should be within 2 seconds on completion of load switching. c. Rarzdom Load Oscillations. With a new CSDU, low system doad or low CSDU oil temperature, random load oscillatixms between paralleled aher­nators are permissible up to + 2 kW beyond the norm~! variation. These oscillations are erratic, of a low frequency and normally in opposition to each other (one alternator taking up the load shed by another); the oscillations should not occur at full system load conditions.

Ram Air Turbine 19. A ram air turbine (RAT) in 1lhe underside of the port air intake can be lowered into the airstream to prov:ide power for electrical services in an emergency.

The RAT drives a 22 •kV A alternator which supplies 3-phase, 200V 400 Hz to the synchronising busbar. No protective features are included. The electrical supplies should not be used below 20,000 feet.

20. When the pilot pulls the RAT release toggle, a Bowden cable operated bomb slip releases the RAT into the airstream and, at the same time, all the non­essential loads (see Table 4) are shed automatically. A function of the non-essential loads being shed is that a cartridge start is pre-selected for the AAPP. Also, a circuit is made to provide RAT field initiation from the vital busbar to ensure a rapid build-up of alternator voltage. The RAT should be ready to supply lt'he loads within 2 seconds; to maintain its output, speed must be maintained above 0 ·SSM or 250 knots (0 · 88M optin1un1) at higher altitudes and suitably adjusted during descent to lower altitudes. The out­put can be checked on the alternator control panel either by pressing the RAT TEST pushbutton or, if the RAT is on the synchronising busbar, by reading the voltage and frequency direct from the synchronising busbar meters.

21. The RAT output cannot be connected to the syn­chronising busbar unless No 2 alternator A breaker is open, the RAT voltage is above 180 volts and there is no supply on the synchronising bus bar. The RAT is automaticaUy disconnected from the synchronising busbar if No 2 alternator or the AAPP is brought on line.

22. Once the RAT has been released, it cannot be ~retracted again in flight. However, in the event of

over-voltage when the RAT is in use, a RAT FIELD switch, NORMAL (mid)/OFF (down), is fitted at the ABO's panel lOP by Mod 2503. Selection to OFF energises the coil of relay RL3 in the voltage control protection unit (VCPU), which connects the RAT field to earth and reduces the RAT voltage output to near zero. ~

Airborne Auxiliary Power Plant (AAPP)

23. The AAPP consists of a gas turbine driving a 40 kVA alternator in a bay immediately aft of the star~ 1board wheelbay. It can provide a 200V supply for use in emergency or for use on the ground when an external power unit is not available. On the ground, and provided that the electrical loading is as Jinlited in Part 2, Chapter 3, para 12, an air-bleed facility may be used for air ventilated suits.

24. The AAPP supplies 200V 400 Hz, 3--phase AC as a standby electrical supply and may be connected to any individual busbar via the synchronising <busbar. Attempts must not be made to S1'art the AAPP above 30,000 feet. During an electrica:l emergency involving t!he use of the RAT, the AAPP must be started and connected to the synchronising busbar above 20,000 feet.

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25. Automatic protection is provided against line-to­earth faults and field overload.

26. The AAPP may be started electrically or by a cartridge. Whenever the non-essential services are tripped, either by operation of micro-switches as the RAT is released or by selection of the NON­ESSENTIAL SUPPLIES TRIP /RESET switch to TRIP, a cartridge start is pre-selected for the AAPP. Following RAT release, the non-essential services may be RESET, in which case the electrical method of starting the AAPP is restored. Cartridge starting of the AAPP must not be used on the ground.

27. A full description of the gas turbine is given in Chapter 5, Engines and the Limitations are given in Part 2.

AAPP Controls and Operation 28.

a. Alternator Control Panel (lOP) (1) AAPP TEST pushbutton for checking incoming voltage and frequency. (2) AAPP ON pushbutton, for bringing the AAPP alternator onto the synchronising busbar (provided that the output voltage is satisfactory). Interlock circuits prevent the AAPP from being connected to the synchronising busbar when No 1, 3 or 4 alternators are connected. If No 2 alternator, the GPU or the RAT are supplying the busbar, pressing the AAPP ON button dis­connects them and allows the AAPP alternator to be connected, provided that its output is satisfactory.

b. AAPP Panel (70P) (1) JPT gauge. (2) Oil pressure gauge. (3) Fuel level magnetic indicator, which shows HIGH when the AAPP tank contains 8 gallons or more, LOW when the tank contains less than 2 gallons and black when the fuel is at an inter­mediate level, or when no electrical power is available. ( 4) A START pushbutton, embodying an amber light (to indicate that the oxygen valve is open). (5) An LP COCK OPEN/SHUT switch. (6) An HP COCK OVERRIDE switch, spring­loaded to OPEN. (7) An IGNITION ISOLATION switch, spring­loaded to ON. (8) A fire warning light and test pushbutton. (9) An OXYGEN & RELIGHT, ON/OFF switch to provide oxygen enrichment for an elec­trical start above 15,000 feet, to override a selected cartridge and achieve an electrical start, to provide a relight facility.

(10) A split 2-pole ON/OFF master switch. Each half of this switch can be operated inde­pendently but the circuit is not complete until both switches are ON. One half of the switch also controls the air intake scoop.

29. Ground and Practice Airborne Startitzg. The drills for ground and practice airborne starting are given in the Flight Reference Cards.

30. Starting Procedures

a. Cartridge Starting. When the RAT release toggle is pulled or the NON-ESSENTIAL SUPPLIES switch is selected to TRIP, the AAPP start circuit is selected to cartridge. Immediate actions must be to switch on the master swit<lh and press the START button, checking that the oxygen light comes on and the START button springs out. Check that JPT, oil pressure, voltage and frequency are within limits. If the first cartridge fails to start the engine, the second may be used (subject to current Command orders and instructions), provided that the light in the button is out.

b. Electrical Starting. If the non-essential busbars are live (indioated thy ,the load shedding indicators on SOP showing a horizontal line) or if tlhe non-essential busbars have 'been RESET following RAT release (when the indicators show a horizontal line super­imposed on cross.Jhatch indication), then electrical starting of the AAPP is selected. Starting is achieved by setting the master switch ON, the OXYGEN & RELIGHT switch ON (above 15,000 feet), checking me light is on, pressing tlhe button firmly and checking that it holds in. Check that the JPT, oil pressure, voltage and frequency are within limits and switch OFF the OXYGEN & RELIGHT switch when the frequency stabilises within limits. An electrical start will also be achieved when the non­essential loads have been shed, by selecting the OXYGEN & RELIGHT switch ON (at any altitude), switching the master switch ON and pressing the START button.

31. To stop the AAPP, put the master switch OFF. If an electrical malfunction occurs, the master switch should be put OFF in two stages, the positive half first and, when the JPT has dropped, the negative half.

Alternator Failure

WARNING : If the red alternator warning light flashes, pull the RAT handle immediately.

32. Failure Warning. If one alternator fails, its amber warning light comes on (indicating that the A breaker is open) and the pilots' and ABO's red alternator failure warning lights shine steadily. If two or more alterna­tors fail, the appropriate A breaker warning lights come on and the pilots' and AEO's warning lights flash to

1-4 Page 6 RESTRIGrED

RESTRICTED AP lOlB-1902-15 Electrical System

indicate a more urgent situation. The red lights are non-essential loads and go out as soon as the non­essential loads are shed ; they come on again when non-essential loads are reset. The amber A breaker warning lights are not affected by load shedding.

33. Single Alternator Failure. If a single alternator fails, its load is taken over by the alternator connected to the synchronising busbar, this being No 3 alternator for all low level flying and No 4 alternator when at altitude. If the alternator connected to the synchronis­ing busbar fails, the automatic take-over facilicy pro­vided by No 2 alternator takes over the synchronising busbar loads. As soon as possible, provided that all loads have transferred satisfactorily, switch the failed alternator OFF and, observing the incoming volts and frequency meters (selected to the failed alternator) press the RESET button. If the volts and frequency remain outside limits, the alternator must be left OFF ; if the volts and frequency are within limits, the alternator may be switched ON. If it again trips off line, switch it OFF.

34. Double Alternator Failure. If two alternators fail, the loads are taken either by the alternator on the synchronising busbar or by the automatic take-over of No 2 alternator. If there is no serviceable alternator on the synchronising busbar (ie if No 2 is one of the two failed alternators), the loads are taken by the RAT. Try to reset the fai1ed alternators; if this is successful the RAT will trip off the synchronising bus bar when No 2 'A' breaker closes. The non-essential supplies, if required, may then be reset, subject to FRC limitations. If neither alternator can be reset and the RAT is sus­taining loads on the synchronising busbar, then descend to below 30,000 feet, cartridge start the AAPP and, when stabilised, replace the RAT by the AAPP. It is then permissible to reset the non-essential busbars for selective use of non-essential loads.

35. Three or Four Alternator Failure. If three or four alternators fail, attempt a rapid reset immediately by switching OFF, pressing the RESET button and switching ON each alternator in the order No 4 to No 1. Simultaneous'ly with this resetting action, release the RAT and descend to below 30,000 feet, maintaining speed above 0 · 85 M/250 knots. If the RAT is sustaining loads on the synchronising busbar, replace it by the AAPP when below 30,000 feet. Only reset the non-essential supplies if absolutely necessary ; in any case, loads may only be restored within the limits of the remaining supply ; airbrakes must not be used.

36. Failure due to Flame-Out. If an engine flames out, descend to a suitable altitude for relighting. If the frequency of the affected alternator falls below 396Hz, switch it OFF. When an engine has been relit, its alternator may be switched on again and normal pro­cedures adopted. If four engines flame out, wind-

milling RPM may be enough to keep the alternators on line if :the aircraft is flown at 0 · 91M/300 knots. At lower speeds or at heights below 32,000 feet, the alternators may come off-line on underspeed. Follow the emergency drill in the Flight Reference Cards.

37. Failure of Protection Circuits andfor Indicators. If a protective circuit or indicator fails, the indications of failure may not be immediately apparent. In these circumstances, services may run down or even stop before the failure is detected. Certain basic principles can be laid down to cover these cases, as follows:

a. If it is not possible to restore power to a failed busbar before the loads begin to run down, switch off all services supplied by that busbar before attempting to restore power.

b. If an alternator has failed but has not tripped off line, switch off the services connected to its busbar before switching off the alternator.

c. If the RAT is on the synchronising bus bar and its output subsequently fails, do not press the extra supplies trip button unless all four alternators are on, serviceable and supplying their services satis­factorily.

~ d. Do not parallel any alterator with one which has been recovered following a failure. ~

Note: The drills for the actions in paras 32 to 37 are detailed in the Flight Reference Cards.

SECONDARY POWER SUPPLIES

Secondary Power Supplies, General

38. The secondary power supplies required to operate the aircraft equipment are provided by trans­former rectifier units, transformers and frequency changers, al'l fed from the primary 200 volt AC busbars.

39. The controls and indicators for the main secondary supplies are grouped on the secondary supplies panel (SOP), which embodies a mimic diagram.

115 Volt, 3-Phase, 400 Hz

~ 40. Four transformers supply 115 volt, 3-phase, 400 Hz power: an NBS transformer (lkVA); a standby~ artificial horizon transformer (40VA) and two main transformers (each 3kVA) which supply the remainder of the 115 volt, 400 Hz services. The two main trans­formers are above panels 27P and 28P on the front wall of the bomb bay; the NBS transformer is on the port side of the nosewheel bay ~ ~ and the standby artificial transformer is adjacent to panel 75P in the cabin.

41. Each main transformer normally supplies its own services but, if one transformer fails, its loads are auto­matically transferred to the other transformer. Two

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Page 7 (AL 6)

RESTRIOT'ED

ON/OFF/COUPLE switches on SOP control the transformers and indications are provided by two three­position magnetic indicators, adjacent to the switches. When the transformers are ON and serviceable, the indicators show line continuity. If one transformer fails and automatic coupling takes place, or if COUPLE is selected, its indicator shows discontinuity. When automatic coupling has taken place, select manual COUPLE in order to disconnect the 200 volt input from the failed transformer. When OFF is selected, the loads have no supply and the indicator shows cross­hatch.

42. The NBS transformer is fed from No 1 or No 4 200 volt busbar, depending on which 1600 Hz fre­quency changer is supplying the NBS.

115 Volt, Single-Phase, 1600Hz 43. Two rotary 3kVA frequency changers provide the 1600 Hz supplies. All loads are normaliy fed from No 2 frequency changer, with No 1 switched off. No 1 frequency changer is used if No 2 fails. The frequency changers are on the port side of the nosewheel bay.

44. Each frequency changer is controlled by START and STOP pushbuttons on 12P and individual mag­netic indicators show black normally and white if a failure occurs. A FREQUENCY CHANGER F AlLURE switch, mounted centrally on 12P, enables the loads of a failed machine to be transferred to the serviceable frequency changer. When No 1 FAIL is selected the 1600 Hz supply for the NBS is taken from No 2 frequency changer and the NBS trans­former input is transferred from No 4 busbar to No 1 busbar.

28 Volt DC 45. Four busbars, interconnected by fuses and circuit breakers, are used to distribute the 28 volt power from two 7 · 5 kW transformer rectifier units (TRU) whose outputs are paralleled, either being capable of supply­ing the total 28 volt system 'load. The TRU are one on each side of the nosewheel bay.

46. The 28 volt loads are supplied from battery, virral, essential and non-essential busbars. The battery bus­bar is connected directly to the positive tenninal of the battery, the only loads being the AAPP starter motor and the battery test socket. The vital busbar is supplied, via 30 amp fuses and diodes, from both the battery and essential busbars, and feeds those loads ~which must be guaranteed a supply at all times. The~

essential busbar obtains its supply from either the battery, via the battery iso!lation contactor, the normal 28 volt plug from the GPU or the 28 volt TRU. The loads connected to this busbar are those required for

the safe flying of the aircraft. The non-essential bus­bars are fed via the port and starboard load shedding contactors from the essential busbar. These loads are not essential to the safe flying of the aircraft but with­out them the aircraft cannot fulfil its operational role. The vital, essential and non-essential loads are listed in Table 4.

~ 47. An additional 0 · 75 kW TRU provides ripple free 28 volt DC supplies for the cartridge discharger. A~ POWER, ON/OFF switch is on the selector control unit at the ABO's station.

48. A 24 volt 40 ampjhour type K2 (nickel cad­mium) battery, located on the starboard side of the nosewheel bay, is connected to the battery busbar but supplies the vital busbar at all times via a fuse and diode. When the battery switch is put to ON (SOP), the battery busbar is connected to the essential busbar for limited general use and for battery charging.

49. The controls for the 28 volt system are on the secondary supplies panel (SOP) and consist of two guarded ON/OFF/RESET switches for the TRU and a guarded ON/OFF switch, which is spring-loaded to the centre position, for the battery. Ammeters are provided in the lines. Three-position indicators between the essential and non-essential busbars show continuity between these busbars normally, show discontinuity when automatic load shedding (either by release of RAT or operation of the TRIP /RESET switch) has taken place and show override (continuity on background of cross-hatch) when RESET is selected after RAT release. The non-essential supplies TRIP /RESET switch on lOP sheds and restores the non-essential loads independently of RAT operation.

~50. It must be remembered that 28 volt DC is required for the operation of all contactors and relays for AC services. The TRU and battery, once switched ON, should be left on for the duration of the flight. Only in the case of a double TRU failure should they be switched OFF, as directed by the Flight Reference Cards. ~

51. A list of the 28 volt loads is given in Table 3 and a list of battery, vital, essential and non-essential busbar loads is given in Table 4. Various AC loads are shown in the non-essential group ; these are con­trolled by 28 volt relays and therefore become inopera­tive when the non-essential busbars are disconnected.

52. When the RAT handle is pulled and the RAT lowered into the slipstream, micro-switch contacts trip the load shedding contactors. All non-essential loads (see Table 4) are then shed. To regain these loads, reference to the FRC must be made and al'l non­essential loads physically switched off. Following this procedure prevents shock loads to the air<:raft electrical system.

1-4 Page 8 RESTRICTED

RESTRICTED AP 101B-1902-15 Electrical System

Failure of 28 Volt TRU

53. If one of the TRU fails in flight, 1he other TRU supplies the 28 volt services satisfactorily.

54. If both TRU fail, all the aircraft DC loads are taken by the battery and, unless immediate action is taken to reduce them, the bauery only lasts for an extremely short period. With all possible DC loads shed, the battery should last for approximately 30 minutes. Therefore, if a landing is not possible within this period, the battery must be switched off.

55. The drill given in the Flight Reference Cards must be actioned immediately.

112 Volt DC

56. 112 volt DC for the H2S scanner is supplied through a TRU. The supply to the TRU is controHed automatically when the H2S is brought into operation. The TRU is in the roof of the scanner compar.tment.

EXTERNALPO~RSUPPLffiS

200 Volt AC Ground Power Unit

57. For ground servicing and aircraft generation pur­poses, a 60 kVA diesel driven or electrical ground power unit (G PU) may be plugged into the aircraft. The 6-pin snatch-disconnect plug is in the port wing root. During a scramble, the socket is snatched out by the initial forward movement of the aircraft.

58. With American type ground power units, the GPU contactor is closed by a pushbutton adjacent to the 6-pin ground supply plug on the aircraft.

59. The external supply from the GPU may be dis­connected from the aircraft system by pressing the extra supplies trip button on lOP, by pressing the AAPP ON button, or by pressing the stop button on the GPU. The 200 volt ground supply is automatically isolated from each main busbar when its alternator is switched on. Alternators cannot be synchronised until the ground supply is disconnected from the synchronis-

~ ing busbar. A phase sequence unit to protect against incorrect phase sequence of GPU supplies is fitted.~

60. If the GPU supply fails, it must not be recon­nected until all those services which were being supplied by the GPU have been physically switched off, other­wise fuses in the distribution network may rupture, necessitating a comprehensive check of this fuse system.

28 Volt DC Ground Supply

61. A 28 volt, quick-disconnect ground supply plug is in the port wing root, angled rearwards. This supply feeds all ground service lighting, all norma~ 28 volt services and, if the battery switch is ON, cllarges the battery. During a scramble, the socket is snatched out by the initial forward movement of the aircraft.

62. For ground refuelling only, a separate 28 volt supply may be plugged in on the starboard side of the nose. This isolates the ground refuelling circuits from the normal 28 volt system, which remains de-energised during refuelling.

63. An aircraft earthing point, for use during refuel­ling and ground servicing (but not during under­carriage retraction tests) is provided in the nosewheel bay and a snatch-disconnect earthing point is in the wing root.

RESTRICTED 1-4

Page 9 (AL 7)

RESTRICTED

Table 1 Main Electrical Supplies-200 Volt, 3-Phase, 400Hz

No 1 Busbar

Essential Port main transformer (US-volt, 3-phase, 400Hz)

Non­essential

1-4 Page 10

PFC motors: No 1 elevon No 8 elevon

Port No 1 fuel pump

RAT blower motor

NBS transforme.r normal supply

No 2 frequency changer

Fuel pumps: Port Nos 4, 5, 7 main

and secondary

Bomb bay tank E fwd stbd

Part ECM

Windscreen demister heater

No 2 Busbar

No 1 (port) 28 volt TRU (also vital busb'ar)

PFC motors: No 3 elevon No 6 elevon

Port No 2 fuel pump

AAPP sump and cartridge heaters

Gold film windscreen heating (low)

Port pressure~head heater

Fuel pumps: Port Nos 3, 6 main

and secondary

Port Nos 1, 7 transfer

Bomb bay tank E fwd port, aft port

Rear cylindrical tank, fwd and centre

Port sequence timers

Airbrakes port motor (standby)

Auxiliary rudder

Part ECM

Hydraulic power pack

Gold film (medium and high)

No 3 Busbar

No 2 (ostbd) 28 volt TRU (also vital busbar)

PFC motors: No 4 elevon No 5 elevon

Rudder main

Stbd No 2 fuel pump

Cartridge discharger TRU

Fuel pumps: Stbd Nos 3, 6 nmin

and secondary

Stbd Nos 1 and 7 transfer

Bomb bay tank A aft port, fwd stbd, aft stbd, fwd port

Fwd cylindrical tank, fwd, centre and rear

Stbd sequence timers

Airbrakes stbd motor (main)

Part EOM

Air ventilated suits

Tail warning

RESTRICTED

No 4 Busbar

Stbd main transformer (llS volt, 3-phase, 400 Hz)

PFC motors: No 2 elevon No 7 elevon

Stbd No 1 fuel pump

IFF/SSR RT2

Standby artificial horizon transformer

Stbd pressure-head heater

NBS transformer alternate supply

Doppler 112 volt TRU ~scanner) No 1 frequency changer

Fuel pumps: Stbd Nos 4, 5, 7 main

and secondary

Bomb bay tank E aft stbd

Rear cylindrical tank, rear

GPI6 and GSR cooling

H2S amplidyne

Airborne warning

TFR

• RESTRicrED AP 101B-1902-15 Electrical System

Table 2 US Volt Supplies

a. 115 Volt, 3-Phase 400 Hz

(1) Port Main Transformer

Bomb fuzing and release NBS (T AS unit) Autopilot Port MFS (port compass, port gyro vertical, port

comparator, variation setting unit) U /V lighting JPT limiters (Nos 1 and 4 engines) Oil pressure gauges (Nos 1 and 2 engines) Fire detection (port wing and fuselage tanks)

~~ Radio altimeter Mk 7B Autostabilisation (mach trimmer, pitch and yaw

dampers) ADF Vibration unit for Mk 19F altimeter IFF /SSR (blower motor only)

(2) Starboard Main Transformer

Bomb fuzing and release De-icing Fluorescent lighting Cabin heating (magnetic amplifier) ~

Starboard MFS (stbd compass, stbd gyro vertical, stbd comparator, variation setting unit, com­puter)

Tacan

JPT limiters (Nos 2 and 3 engines) Oil pressure gauges (Nos 3 and 4 engines) Fire detection (stbd wing and bomb bay tanks) Fuel fiowmeters IFF (fan supplies) Autostabilisation system (as for port transformer)

(3) NBS Transformer

H2S NBC HRS GPI6 GSR

(4) Standby Artificial Horizon Transformer

Standby artificial horizon

b. 115 Volt, Single-Phase, 1600 Hz

(1) No 1 Frequency Changer NBC H2S

(2) No 2 Frequency Changer

Radar altimeter

Periscope heater c. 115 Volt, Single-Phase, 400 Hz Mk 30A altimeter

Table 3 DC Supplies

a. 28 Volt DC NBS, calculator type 7, bomb gear and bomb

jettison Radio altimeters Mks 6A and 7B, radio compass,

ILS, Tacan IFF /SSR, HF, RTl, RT2, intercom Window launchers, tail warning, ECM

MFS, T ASU, VG recorder and fatigue meter. Pressure-head heater control

Artificial feel, trimmers, PFC warning, auto­stabilisers

Autopilot

General instruments, fuel contents, LP cocks, cross­feed cocks, fuel pressure warning

JPT gauges AAPP gauges

Bomb bay cold air valve, cabin heating, pressurisa-tion and warning

De-icing, de-misting, windscreen wipers Red flood and white lighting, external lighting Ration heaters Tail parachute, UC operation, indication and nose­

wheel steering Canopy and entrance door warning Relighting, AAPP starting, fire extinguishers, crash

switches Load shedding, relays, battery charging and isolation,

frequency changer excitation F95 camera Vibration unit for Mk 29B altimeter

b. 112 Volt DC H2S scanner

RESTRicrED 1-4

Page 11 (AL 9)

Vital Loads

(supplied direct from battery)

Engine fire extinguishers

Fuel tank fire extinguishers

LP fuel cocks

Battery isolation

Crash switches

Engine relighting

28 volt DC control

RAT excitation

AAPP control circuits

Pressurisation control

Emergency door opening

Cabin decompression

Abandon aircraft sign

Crew escape lights

Panel 9P illumination

Battery Busbar Loads

AAPP starter motor

Battery test socket

RESTRICTED

Table 4 Vital, Essential and Non-Essential Loads

Essential Loads

(remaining after RAT has been lowered)

PFC motors (except rudder auxiliary) ~

PFC warning

Artificial feel and trimmers

Autostabilisers

Nos 1 and 2 fuel pumps, port and starboard

Fuel pressure warning, crossfeed cocks

RTl, RT2

HF, IFF/SSR

Intercom

MFS and standby artificial horizon

General instruments

Pressure-head heating control

Windscreen wipers

Pressurisation warning hom and lights

Cabin lighting

Fire detection and warning

Battery charging

28 volt DC control

115 volt 400 Hz control

Bomb jettison, bomb doors

Undercarriage operation, indication and nosewheel steering

Tail parachute

Canopy and entrance door warning

AAPP indication, sump and cartridge heaters, cartridge start and 200 volt AC controls

Gold film, low heat

Engine fire-warning test

Air-to-air refuelling

RPM governor

Non-Essential L oads

(shed by operation of RAT or by ABO)

External lighting, except navigation and anti-collision lights. (Non-essential cabin~ lighting to be switched off if restoring loads).

Air temperature indication

Pressurisation underheat control, bomb bay conditioning and temperature indi­cation

Ventilated suits

Ground refuelling, fuel transfer pumps, flowmeter selection, fuel contents and nitrogen

*Bomb bay fuel pumps (if fitted). Fuel tank pressurisation, hold-off supply

*Wing tank fuel pumps and sequence timers

Autopilot

*ILS

VG recorder and fatigue meter

Airbrakes (switch corresponding to air­brakes position before restoring loads)

Hydraulic power pack

Rudder auxiliary motor (but see Chap 7, para 19)

Alternator failure warning

Frequency changers

Wing and engine anti-icing control

Normal engine starting

ADF, sextant heater, air mileage unit

Radar altimeter Mk 6A

*Bomb gear, NBS, HRS, Doppler, ECM, tail warning, calculator type 7, window launching

*Ration heaters

*Tacan

*TFR ~Navigation and anti-collision lights • *Radio altimeter Mk 7B

Vibration unit for Mk 29B altimeter Gold film, medium and high heat

*Auxiliary AVS and face blowers

Tail skid proximity warning lights

* These loads muse be selecced OFF mamtally before rescoring non-essential loads.

1--4 Page 12 RESTRICTED

•

RESTRICTED i\P IOIB-1902-15

PART 1

CHAPTER 5-ENGINES AND AAPP

Contents OLYMPUS ENGINES Para

General 1 Throttle and HP Cock Controls 7

Engine Fuel System 10

Engine Starting System 12

Oil Systems ... 18

Engine Instruments 21

~~ Throttle Detents 22

OLYMPUS ENGINES General 1. The aircraft may be powered by Olympus Mk 200 series or Mk 301 engines which, due to differences in mechanical details, thrust and limitations, are never installed together on the same aircraft. In the TAKE OFF setting, Mk 201 series and Mk 301 engines produce approximately 17,000 lb and 18,000 lb of thrust respectively; in the CRUISE setting they pro­duce approximately 16,000 lb and 17,000 lb of thrust respectively. A rapid stanting system is embodied.

2. The basic engine consists of eight main sub­assemblies (see Fig 1). The intermediate casing houses dle drives from both compressors for the mechanica:Hy driven auxiliaries:

a. LP Compressor Components (1) LP driven fuel pump (2) Tacho generator (RPM)

b. HP Compressor Components (1) HP driven fuel pump (2) Main oil pump (3) Malin oil scavenge and four auxiliary scavenge pumps (4) Constant speed drive unit (CSDU) (5) Hydraulic pump (6) Tacho generator (not used) (7) Rutax air starter drive

~ 3. Engine air is automatically supplied on start-up for the following systems :

a. Turbine cooling. b. Pressurisation of bulkhead seals, oil bearing seals and CSDU oil tank. c. Induced cooling of alternators, CSDU oil and Zone 2B.

Air is also available for engine anti-icing when selected.~

4. Engine control is provided by a combined throttle/ HP cock. Automatic control is effected by components

AIRBORNE AUXILIARY POWER PLANT Para AAPP Description ... 23 AAPP Oil System ... 27 AAPP Controls and Instruments 29

Dlustrations Fig Bngine Layout 1 Engine Air Supplies for Normal Star:cing,

Air Conditioning and Antli.-icing ... 2 Engine Fuel System 3 Engine Oil System ... 4 Engine Rapid Start System 5

in the fuel chassis which allow an optimum supply of fuel to the engine, dependent on ambient conditions, airspeed and acceleration stage in the engine. Limita­tion on engine output is controlled by the position of a Take-off/Cruise switch which determines both the RPM governing datums and the corresponding JPT limiting datums for all four engines. Also, a jet-pipe temperature limiter is installed which controls the engine speed below the selected governing speed, if necessary, so that the corresponding jet-pipe tempera­ture limit is not exceeded. When JPT limiting is over­ridden, manual control of the throttles may be necessary to maintain the engines within JPT limitations.

5. The Mk 200 series jet pipes are fitted with con­vergent/ divergent nozzles for improved cruise per­formance; the Mk 301 jet pipes are fitted with convergent nozzles.

6. To minimise damage to the outer wing and the bomb bay following structural fai'lure of an engine, containment shields are fitted outboard of the outer engines and inboard of the inner engines.

Throttle and HP Cock Controls 7. The four throttle levers, which also control the HP cocks, are forward of the retractable centre console, in a quadrant marked OPEN/IDLING. The quadrant is gated at the IDLING position and the part of the quadrant below this, which controls the HP cock position, is marked OPEN/HP COCKS/SHUT and has a further gate at the SHUT position.

8. To move the throttle levers forward from the HP cocks SHUT gate and to move them aft from the IDLING gate, .the sleeves on the levers must be raised.

9. The throttle friction lever is on the starboard side of the throttle quadrant ; forward movement of the lever increases the friction.

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Page 1 (AL 7)

RESTRICTED

Engine Fuel System

10. The engine fuel system controls the correct amount of fuel the engine requires in the varying conditions of flight. The factors which effect this are:

a. Throttle setting.

b. Engine air intake pressure (PI).

c. Compressor delivery air pressure (P3).

d. JPT limiter isolate switch position.

e. JPT /RPM (datum limits selected by the TAKE­OFF /CRUISE switch).

11. Fuel from the tank booster pun1ps flows through the LP cock and fuel filter into the engine fuel system (Fig 3) in the following order:

a. LP and HP D1-iven Fuel Pumps. Two engine driven pumps are inter-connected and driven by the

~ LP and HP compressors respectively. ~

(1) The LP pump incorporates a double datum hydro mechanical RPM governor, a:llowing it to control RPM to take-off or cruise limits with the throttle at OPEN. The two RPM datums are controlled by the TAKE-OFF/CRUISE switch to the left of the throttle levers. In addition to selecting the required RPM datum, operation of the switch also selects the corresponding tem­perature datum for the engine JPT limiter. Therefore, actual RPM on the engine may be Jess than the limit selected in order to keep the JPT within limitations. The hydro-mechanical governor maintains the selected RPM datum regardless of changes in temperature and density of fuel; RPM may vary very slightly with increase in altitude.

(2) The HP pump incorporates an overspeed governor set at 3% above maximwn HP com­pressor revolutions should the governor on the LP driven pump fail.

b. Full Range Flow Control (FRFG). Fuel is passed from the fuel pumps to the FRFC which meters fuel to the engine depending upon :

(1) Throttle-HP cock position.

(2) Engine air intake pressure (Pl).

(3) /PT Datum Limits. Control is accom­plished by the Electro Pressure Control (EPC) which is activated by the JPT amplifier when the JPT approaches the limit selected by the TAKE­OFF /CRUISE switch. In the TAKE-OFF setting a depressed datum is brought into opera­tion which prevents turbine temperature overswing during rapid throttle opening by slowing down the rate of temperature rise during the last 5°C. Above about 20,000 ft, the TAKE-OFF and CRUISE datums may exceed their respective limits by up to 5°C, therefore in the TAKE-OFF setting at altitude the JPT may have to be kept

within limits by use of the throttles. The output from the JPT amplifier may be isolated by means of a JPT limiter ON /OFF switch aft of the TAKE-OFF/CRUISE switch. The switch must always be ON before TAKE-OFF is selected, so that the depressed datum will prevent turbine overswing during rapid throttle opening. The OFF position of the switch is for use when ground testing the RPM governors, or for usc if a limiter runs away or fails in flight. In these circumstances, rthe throttles must be handled care­fully to prevent JPT rising to the limits. The limiters may be tested on the ground by isolating the governors; two RPM GOVERNOR ISOLA­TION switches are provided at the top of the port fuse panel behind the AEO's seat. When held ON they permit take-off RPM to be selected when the pilot's switch is at CRUISE, without altering the cruise setting of the JPT limiter. Nos 1 and 4 and Nos 2 and 3 engine JPT amplifiers are supplied by the port and starboard main transformers respectively. If automatic load­shedding occurs, the system is de-energised and JPT control must be maintained by use of the throttles.

c. Air Fuel Ratio Cont1·ol (AFRG). The AFRC is the engine acceleration control. During accelera­tion it senses compressor delivery pressure (P3) and thus determines how much fuel may be passed to the burners.

(1) In Mk 200 series engines the AFRC over­fuels the engine in the early stages of acceleration. As fuel delivery is proportioned to P3 pressure, it is merely necessary to restrict the P3 pressure fed to the AFRC. This is done by incorporating Pl P3 switch which measures the compression ratio of the compressors, and decides how much P3 pressure may be fed to the AFRC. Above 82/83% RPM the overfuelling condition ceases and the Pl P3 switch allows full P3 pressure to the AFRC.

(2) In Mk 301 engines overfuelling is not a problem, so a Pl P3 switch is not required. The internal bellows of the Pl P3 switch are removed, thus it becomes an air potentiometer and full P3 pressure is fed to the AFRC. The AFRC is now working at its designed limit, but engine accelera­tion is still too slow. To enhance acceleration, a Speed Term Switch is fitted. It senses LP fuel pump pressure, and after 65% progressively allows more fuel to bypass we AFRC to the burners, until max RPM is reached.

d. Flo·zo Distribution and Dump Valve. During start up, a fine fuel spray is required and fuel is initially only fed to the primary orifices of the duplex burners. Above about 15% RPM fuel is also allowed to the main orifices of the burners. The

1-5 Page 2 RESTRICTED

AIR INTAKE CASING DELIVERY CASING AP lOl B-1902-15

Engines and Airborne Auxiliary Power Plant

INTERMEDIATE CASING

TURBINE ASSEMBLY

-------- -......._ __ --./

------ ...... ..... ---~

HP COMPRESSOR

LP COMPRESSOR FLAME CAN ASSEMBLY

1-5 Fig. 1 Engine Layout

TO PNEUMATIC

FLAMES TAT

ELEMENT

1\ BUr

SENSiaG \ TEMP

~ ~J

- HOT AIR FROM ENGINE

- COLI>AIR

- 80M8 8AY HEATING

WING ANTI - ICING

fEMPER4TURE SENSING BULB

l'EMPERATURE SENSING ELEMENT NOTE:­

EXHAUST ANNULUS

COLD AIR INTAKE tNOUNG ELEMENT

Au=-. EXTRACTOR L ~& ACTU4TOR

!/- UCT TEMPEAATURE

IJUI.8

--ENGINE DIVIDER FLAMESTAT...,.. ~ WAL..L HOT AIR SUPPLY

WING LEADING EDGE TEM PERATURE SENSING ELEMENT

CIRCUIT SHOWN FULL COLD-TROPICAL-SEA LEVEL

1-5 Fig. 2 Engine Air Supplies for Normal Starting, Air Conditioning and Anti-icing

RESTRICTED 1-5

Page 3

I

RESTRICTED

JPT

xxx-~----~---l-~-·--0 ENGINE <(CONTROL

~ D/E

\

v p1

PRIMARY

MAIN

•

FUEL DISTRIBUTOR DUMP VALVE

FUEL FLOW

Tx

.....

t I

JPT AMP

UNIT

FUEL COOLED

OIL COOLER

-

28V __ _. ----1 TO/CRUISE SWITCH

DATUMr-------~-----~ JPT

UNIT .,_ ____ .,_ ____ 11-~ LIMITER

FUEL DIPPING

VALVE

r-----,

sw

I SPEED t ------===-I TERM I r I SWITCH J I L----1

T ' '

AIR­FUEL

RATIO CONTROL

~ I I

ELECTRO-PRESS CONT

FULL RANGE - FLOW

CONTROL

1-5 Fig 3 Engine Fuel System

RESTRICTED

1-

J ..

JPT AMP

UNIT

FUEL COOLED

OIL COOLER

..

••

RESTRICTED

ENGINE 9 CONTROL ~ D/E

I TO/CRUISE 28V-------t I SWITCH

THROTTLE

HP QUAD

DATUM 1-----4--~ JPT

t-----.. ---·1-~ LIMITER sw

UNIT

FU EL DI PPING

VALVE

r-----, I

1 sPEED r-------0:::: --c::: --, I TERM I l SWITCH L ____ j , 1 I ELECTRO-

4 PRESS CONT L I AIR- I FUU...

THR

- ( FUEL RANGE

HP RATIO FLOW VALVE ~ -CONTROL CONTROL

1 P1 / PJ swf

I a

•

1- 5 Fig 3 Engine Fuel System

RESTRICTED

AP I OIB-1902-15 Engines and Airborne Auxiliary Power Plant

FUEL PRESSURE D/E

,,

TO/CRUISE GOV

LP PUMP

HP PUMP

0/SGOV

PRES .,_......,...,__ 28v sw -

Flt;ER~ L p

COCK

1-5 Page 5 (AL5)

RESTRICTED

01 L TANK OI L T ANK CONTENTS GAUGE

PRESSUR I ZATIO~N NON - RETURN. VALVE VALVE

~~~=-==-=ll - -------- -- -----OIL TANK

.. , I I II II II ----

OI L TANK DRAIN

OIL TANK OVERFLOW AVERY HOSE CONNEX I ON

OI L TANK VENT TO ENGINE BREATHER SYSTEM VIA SEPARATOR

-- SUCTION

--

1-5 Page 6

PRESSURE FEED

SCAVENGE SUCTION

SCAVENGE RETURN TO TANK

TANK VENT

SUNDSTRAND AND STARTER DR IVE CASTING

0 1 L PRESSURE T RANSMIT T ER

O I L SUMP DRA I N AVERY HOSE CONN EX ION

NOTE:-

THE SUNDSTRAND OIL SYSTEM IS SELF-CONTAINED AND ENT I RELY INDEPENDENT OF THE ENG I NE OIL SYSTEM

1-5 Fig 4 Engine Oil System

RESTRICTED

DRIVE

-----, I I I I I I I I I I I

~SCAVENGE FILTERS I

-~=-= ;;;·;..----· EMPERATURE

•

• RESTRICTED AP 101B-1902-15 Engines and Airborne Auxiliary Power Plant

dump valve ensures that when the HP cock is shut, all fuel from the burner pipes is drained into a collector box on the engine doors. The box is drained by the ground crew. It will only partially drain in flight. e. Dipping Valve. It is only used during a rapid start of the engine and ensures that excess fuel returns to the suction side of the pumps. It is automatically de-energised at the end of the start cycle. However, it can be held open by fuel pressure. To ensure that it has closed, the throttle must be briefly returned to the idling gate at the end of the start cycle.

Engine Starting Systems 12. Each engine embodies its own air starter motor. Air can be supplied from a ground air starter unit (Fig 2), feeding through a connection on the underside of the starboard wing, or from the rapid start system (Fig 5). The ground air supply feeds into the main lines of the engine air system and thence to the starter motors, through electrically-actuated valves. Com­pressed air from a running engine can be used to start the others, singly or simultaneously, provided that the appropriate ENGINE AIR switches on the starboard console are at OPEN.

13. The rapid starting system is so arranged that the powered flying controls and ar·tificial feel are switched on automatically when the simultaneous RAPID START button is pressed; because of the peak loads involved, electrical power during starting must be supplied by a 60 kV A ground power unit and not from the AAPP.

14. A gyro hold-off system is embodied, whereby the MFS, JPT limiters, refuelling relays, contents gauges, fire warning, autostabilisers and artificial horizons are de-energised until the engine start master switch is selected ON. The power to the gyros is initially boosted to 200V for approximately 20 seconds, to obtain fast run up.

15. Controls

a. The starting control panel on the 1st pilot's port console carries the following controls:

Four buttons for individual engine starting, each embodying a light.

A GYRO HOLD OFF butron.

A RAPID START button, for starting all engines simultaneously.

A NORMAL/RAPID lock..,toggle switch, for select­ing palouste or high pressure air.

An IGNition ON/OFF switch

An AIR CROSSFEED three-position magnetic indicator.

An ON/OFF lock-toggle start master switch.

b. An AIR CROSSFEED indicator shows OPEN whenever the start master switch is ON. When the air selector is at NORMAL, engines can only be started by the individual buttons, using an external air supply or crossfeed. When the air selector is at RAPID and the master switch ON, all engines can be started simultaneously using the RAPID START button or separately, using the individua'l buttons. If, for any reason, an engine fails to start on this system, there is sufficient air for one further start on each side; the individual start button must be used for this attempt. With the air selector switch at RAPID and the master switch OFF, the gyro hold-off system is effective when the GYRO HOLD OFF button is pressed.

c. Relighting Controls. A relight button in the head of each throttle/HP cock provides a means of relighting the engines in Bight. When one of the butt'Ons is pressed, 28V vital busbar power energises the igniter plugs regardless of throttle position or switch selections on .the engine start panel. The igniter plugs remain energised until the relight button is released.

16. Operacion (NORMAL). With the IGNition switch ON, MASTER switch ON and air selector switch to NORMAL, pressing a start button energises three circuits :

a. Solenoid to open position on starter air control valve.

b. Engine igniter plugs.

c. Palouste air bleed valve.

The increase in palouste air opens the air control valve and a pressure switch lights the start button. The air now rotates the starter turbine which drives the HP compressor. Fuel from the HP pump is directed to the primary burners where the igniters initiate combustion. The engine accelerates to self-sustaining speed and disengages from the starter turbine. An overspeed switch on the starter turbine operates the start relay which now de-energises the engine igniter plugs and the palouste air bleed valve and operates the solenoid on the air control valve to the closed position. When the air control valve closes, the pressure switch breaks, cancelling the light in the starter button.

17. Operation (RAPID). The rapid start facility uses a mixture of bottled air and fuel from the booster pumps. When this mixture is ignited, the resulting hot gases tum the starter turbine.

a. Air at 3300 PSI is stored in four booties in each of the two engine bays. A charging point is on each wing between the jet-pipe tunnels. Sufficient air is available for three engine starts per side. Air passes from the bottles to the manifold and thence to air bottle solenoid valves on each engine.

REST RICTED 1-5

Page 7 CAL 9)

RESTRICTED

MULTI-

....__ _ __ C...,ONNECTOR ~ CHARGING VALVE -

PRESSURE RELIEF VALVE

... AIR SUPPLY _, TO INBOARD ENGINE

GAUGE

AIR SUPPLY STARTER MOTOR ,------- ...

I \ I []

TO OUTBOARD ENGINE I

PRESSURE \ I , ________ .., ... AIR CYLINDERS

REDUCING VALVE 3 300/300 PSI

EXHAUST

DRAIN PIPE

1-5 Fig 5 Engine Rapid Start System

b. A solenoid-operated valve opens when the starter button is pressed and allows high pressure air from the storage cylinder to pass to the reducing valve, which reduces the pressure to 300 PSI for delivery to the combuster. A safety disc, which bursts at 550 PSI, protects the combuster from excessive pressures if the valve fails. If the disc bursts, air exhausts into the same ducting as the normal exhausts from the starter.

c. Air from the bottles, reduced to 300 PSI, flows to the combuster unit and mixes with pressurised fuel in the chamber. The remaining air passes between the two walls of the combustion unit for cooling purposes and also to a non-return valve to prevent air escaping through to the air control valve. The fuel is then atomised, mixed with the air, ignited and ejected on to the starter turbine. A pressure switch then operates the light in the starter button and maintains the electrical circuit to the start components, while a 2-second timer switches off the combuster igniter. The combuster continues to operate until, after a successful start, the cycle is terminated by the starter overspeed switch or the 12-second delay in the time delay unit. The light in the starter button goes out, the air bottle solenoid is de-energised and the time delay units are reset.

d. If the igniter plug fails to ignite the mixture, the pressure switch does not operate. This breaks the electric supply to the time delay units after 2 seconds and the air bottle solenoid and the time delay units are reset.

e. If the combuster ignites but the engine fails to light up, the rapid start sequence is terminated by the 12-second delay and the time delay unit is reset.

Oil Systems

18. Each engine has its own integral oil system. In 200 series engines, the tank is in the nose-bullet, a float­operated contents gauge is on the port side 'Of the air­intake casing and the tank filling point is on the lower port side of the engine front bulkhead. In 301 engines, the tank is on the port side of the LP compressor casing and has a contents sight glass, the fiWng point being betow the tank. The oil capacities are as follows:

Total oil (tank Tank oil Airspace and engine)

Engines capacity (galls) (galls) (galls)

200 series 4! lt 6t 301 3i ll 6i

19. Each engine also has an independent oil system for the alternator constant speed drive unit. The 14! pint tank (3t pints air space) is attached to the starboard

~ side of the intermediate casing. A tank sight glass~ shows tank contents. The tank filling point is on the underside of the constant speed drive unit, at the lower starboard side of the engine. The tank must be replenished within 15 minutes of shutting-down the engine. Great care must be taken not to overfill the tank or oil will find its way into the cabin conditioning system.

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20. Magnetic chip detectors (five) are in all oil recovery lines, as follows:

1. No 1 bearing

2. Nos 2 and 3 bearings and auxiliary drives

3. Nos 4 and 5 bearings

4. No 7 bearing

5. No 8 bearing

Engine Instruments

21. The following engine instruments are grouped together on the pilot's centre panel:

~ a. Four JPT gauges which are DC operated. ~

b. Four fuel pressure magnetic indicators, which show white when there is insufficient pressure.

c. Four RPM indicators, of the percentage type. The main scale is calibrated in tens from 0 to 100%, while a small scale gives readings from 0 to 10%.

d. Four oil pressure gauges.

e. One ENGINE CONTROL magnetic indicator (inoperative).

Throttle Detents

22. a. To reduce the possibility of compressor instability leading to engine flame-out, a detent is provided on the inboard throttles of 301 engines to

,.,..,. increase flight i~bov"- 15,000 feet. It gives ~. an increase of -J~ "Rf>M over the outboard·

engines at 50,000 feet. The detent can be manually overridden in emergency, eg rapid descent or relighting.

b. A detent isolation switch at the forward end of the starboard console permits withdrawal of the detent under air-to-air refuelling conditions, ie at airspeeds not exceeding 250 knots at medium altitude.

c. For use on the ground, a detent test switch is provided on Panel 4P.

AIRBORNE AUXILIARY POWER PLANT

AAPP Description

23. The AAPP is a gas turbine engine, outboard of No 4 engine. It provides an emergency 200 volt supply below 30,000 feet. On the ground, the AAPP supplies eleotrica'l power and air for the ventilated suits.

24. The engine has a single-side centrifugal com­pressor, driven by a single-stage axial turbine on a common shaft. The air intake is on the underside of the engine and air passes through a single, reverse-flow combustion chamber. The jet efflux is directed down­wards from the wing. The air intake is held in the closed position by hydraulic pressure, electrically operated. If hydraulic or electrical failure occurs, the intake opens.

25. No throttle control is provided, as the engine is designed to run at a constant speed. The air-intake shutter is controlled automatically by operation of the master switch which, with all the other controls, is at the AEO's station. Chapter 4, Electrical System, describes the use of the AAPP as a standby electrical supply.

26. For starting at higher altitudes, oxygen enrich­ment is provided. The unit carries its own supply in two bottles at the rear of the engine.

AAPP Oil System

27. Engine 'lubrication is provided by a gear-type pressure pump, which draws its supply from an oil sump formed by the lower part of the compressor casing. The sump capacity is 4t pints. The sump must always be filled to capacity.

28. The oil pressure varies very rapidly with tempera­rure; the minimum pressure is 4 PSI.

AAPP Controls and Instruments

29. a. The AAPP controls and instruments are described in Chapter 4, para 28.

~ b. The JPT gauge is DC operated. ~

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PART 1

CHAPTER 6-FIRE PROTECTION SYSTEMS

Contents

ENGINE FIRE PROTECfiON

Olympus Power Plant Fire Extinguishing System

AAPP Fire Extinguishing System

Para

1

5

FUEL TANK FIRE PROTECTION

Fuel Tank External Fire Protection 8

HAND FIRE EXTINGUISHERS AND INERTIA SWITCHES

Hand Operated Fire Extinguishers 15 19 Inertia Switches

Dlustrations

Engine Fire Protection

Fuel Tank Fire Protection

ENGINE FIRE PROTECI'ION Olympus Power Plant Fire Extinguishing System

1. Each engine bay is divided into three zones: Nos 1, 2A and 2B. Zone 1, which holds the LP compressor, has no fire protection. Zone 2A, which holds the HP compressor, has six flame detectors and Zone 2B has five. Each detector consists of one shrouded and one unshrouded thermocouple ; a temperature differential of 250°C induces sufficient current to operate the control unit relays and give fire warning. Fire warning indication cancels when the temperature differential drops below 250°C.

2. There are two methyl-bromide extinguisher bottles on each side of the bomb bay. The forward one of each pair serves the inboard engine and the rear one the outboard engine. The bottles feed through spray pipes on the sides of the engine bays.

3. Controls and Indicators a. Four guarded fire-extinguisher pushbuttons, one for each engine, are on the coaming above the centre instrument panel ; each button incorporates a red warning light. The lights may be tested by gently pulling out the pushbuttons, when the light should come on. The buttons return by spring action but they must not be allowed to snap back and they must not be pushed back. The action of the buttons is very sensitive and they must be treated with care. The fire warning is inoperative while gyro hold-off is engaged.

b. A test switch is provided at the top of panel 4P for testing the whole fire-warning circuit. All four

Fig 1

2

circuits are tested simultaneously and all four warn­ing lights should come on. A ground test facility is provided in the nosewheel bay. The test facility is supplied from the essential busbar.

4. Operation a. If an engine fire is indicated by a fire warning light coming on, the drill in the FRC must be carried out without delay.

b. The warning light goes out when the fire is extinguished and the temperature differential drops below 250°C. On no account make an attempt to relight the engine, as no extinguisher is available if another fire occurs.

Note: The fire warning light could also go out as a result of a damaged thermocouple, as they are wired in series.

AAPP Fire Extinguishing System

5. Four flame detectors provide fire detection in the AAPP, three on the rear panel and one on the top rail. The flame detectors are set to operate at a temperature difference of 185° C. The methyl-bromide extinguisher bottle is on the outside of the AAPP fireproof bulkhead and operates through two spray rings.

6. Controls and Indicators. A guarded fire-extin­guisher button, embodying a warning light, is on panel 70P. Below it is a FIRE TEST pushbutton, which is used to test the warning system. When the extin­guisher button is pressed, the LP cock is closed automatically.

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7. Operation a. H the fire warning light comes on, take the following actions :

Extinguisher button Press Master switch OFF

Subsequently, set the LP cock switch to OFF.

b. The warning light goes out when the fire is extinguished and the temperature differential drops

~ below 185°C. The AAPP cannot be restarted once~ the extinguisher button has been pressed.

FUEL TANK FIRE PROTECTION

Fuel Tank External Fire Protection 8. The fuel tank and leading edge fire protection system is operated automatically by fire-wire sensing elements, installed around the tank bays. When any part of the sensing element is heated to approximately 250°C, the circuit is made to the appropriate extin­guisher bottles. Triple-F-D fire detector units are pro­vided, two for the fuselage tanks, four for the wing tanks and one for the bomb bay tanks. See also para 18 for inertia switch operation.

~ 9. Manually Operated System. Post SEM 012, the wings and fuselage automatic system is made inoper­ative, the indicator deleted and the bomb bay fire warn­ing indicator repositioned. Two manually operated flap-guarded push switches on the co ... pilot's panel give independent control of the port and starboard extin­guisher systems, the switches are labelled 'WING I FUSELAGE, FIRE EXTINGUISHERS, PORT­STARBOARD'. Operation of these switches will dis­charge all fuselage, wing and leading edge fire extin­guishers but the wing tank fire extinguishers will 'be discharged to ohe inboard fuel tank bay only. •

10. Fuselage Tanks. 11he No 1 and No 2 tanks are protected by four methyl-bromide fire-extinguisher bottles, two for each pair of tanks. The contents of the bottles are discharged into the appropriate tank bays through spray pipes.

11. Wing Tanks. Tanks Nos 3, 4 and 6, in the inboard tank bay, and tanks Nos 5 and 7, in the out­board bay, are protected by six dual-head fire-extin­guisher bottles in each wing. The bottles can supply either the outboard or the inboard bay but cannot supply both bays simultaneously. The firing of either operating head discharges the total contents of the bottle.

12. Leading Edge Installation. Three extinguisher bottles in each wing, forward of the main wheel units, supply a spray pipe inside the anti-icing ducting in the leading edge. The extinguishers are actuated by the wing tank sensing units.

13. Bomb Bay Fuel Tanks. When provision is made for bomb bay fuel tanks, eight methyl-bromide extin­guishers are installed, four on each side of the bomb

bay. The extinguishers are automatically operated by fire-wire sensing elements, or by the inertia switches. 14. Indicators. Two red warning lights, on the co­pilot's instrument panel, one for the wings and fuselage and one for the bomb bay, indicate if a fire in the fuel tanks has caused the extinguishers to operate. The lights go out when the temperature falls below 250°C. Fuse-type indicators on the rear face of the front spar, on the port side of the bomb bay, show if the bomb bay extinguishers have operated. A test switch light is on a panel on the starboard side of the nosewheel bay. A relay in the circuit prevents inadvertent operation of the extinguishers when the test switch is operated.

~Post SEM 012, only one red warning light for the bomb bay fire warning is fitted to the co-pilot's panel. ~

HAND FIRE EXTINGUISHERS AND INERTIA SWITCHES

Hand Operated Fire Extinguishers 15. Five bromoohlorodifiuoromethane (BCF) hand­operated fire extinguishers are in stowages in the crew's compartment:

a. Behind each pilot's seat (2) b. Outboard of the nav /radar's seat (1) c. Behind the AEO's seat on the forward and rear wall of the HRS crate (2).

16. A BCF fire extinguisher is provided for external use and is stowed in the external emergency equipment compartment.

17. BCF extinguishers are a universal rype of extin­guisher which may be used on any type of fire without restriction.

18. Fire in the Cabin. The majority of fires in the cabin are electrical in origin and can usually be con­trolled by switching off all equipment in the vicinity. If the fire still persists, the hand extinguishers may be used. As circumstances permit and once the fire is extinguished, reduce atbin pressure by selecting COMBAT or NO PRESSURE (depending on altitude), to flush out the fumes more rapidly. Oxygen regulators should be selected to 100% and emergency and care taken not to get any BCF in the eyes.

Inertia Switches 19. Inertia switches, on the lower forward face of the rear pressure bulkhead, automatically discharge the engine, AAPP, fuel tank extinguishers (wing extin­guishers to inboard tank groups) and bomb bay extinguishers in a crash landing, when deceleration exceeds 4-tg. At the same time, the alternators are automatically switched off, the AAPP (if in use) is automatically stopped and the 24 volt battery is dis­connected from the essential busbar. By automatically switching off all the electrical supplies in the aircraft (except for the vital busbar supplies), risk of fire is minimised.

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•

--

DETECTOR HEAD

INERTI A SWITCH PANEL

LOCAT ION OF CONTROL UNITS AND INERTIA SWITCH E S

LOCATION

AP 101B-1902-15 Fire Protection Systems

. l lfi OF FIRE DETECTORS (ENGINE BAYS IDENTICAL)

KEY

r::::::==::=J SUPPLY PIPE

c=::==;;;;;; SPRAY PIPE

--- FIREWIRE DETECTOR

1--6 Fig. 1 Engine Fire Protection

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1-6 Page 4

+- FWD.

DETAIL A

FUSELAGE TANKS

WING TANKs

I~ Fig. 2 Fuel Tank Fire Protection

RESTRICfED

TO OUTBOARD WING TANKS

TO INBOARD WING TANKS

KEY

~FEED PIPE

C::::.:. SPRAY PIPE

- FIREWIRE DETECTOR ~ FLEXIBLE PIPE

AP 101B-1902-15

PART 1

CHAPTER 7-FLYING CONTROLS

Contents

Cockpit Controls and Indicators ...

Powered Flying Controls Elevons Rudder Powered Flying Control Units (PFCU) Electrical Supplies Controls Starting the PFC's PFC Failure ...

Artificial Feel Artificial Feel Units ... Rudder and Elevator Feel Aileron Feel ... Feel Relief Electrical Supplies Controls Starting Feel Locking ... Artificial Feel Failure

Trimmers Description ... Controls Trim Failures Electrical Supplies

Auto Stabilisers General Controls Electrical Supplies Yaw Dampers Yaw Damper Failures Pitch Dampers Pitch Damper Failures Auto Mach Trimmer Auto Mach Trimmer Failures

Airbrakes General Control and Operation

~ Failures

Brake Parachute ...

COCKPIT CONTROLS AND INDICATORS

1. The flying controls in the cockpit are conventional in operation. Dual interconnected control columns and pendulum-type rudder pedals are provided ; these operate powered controls through a series of linkages.

2. The rudder pedals can be adjusted for reach by a starwheel at the lower inboard edge of each pilot's

Para. 1

7 10 11 13 14 16 17

24 25 26 27 28 29 31 32 37

42 44 46 47

48 49 50 51 53 55 58 61 63

68 70 76 ~

77

instrument panel. Toe-buttons are provided on the pedals for brake operation.

3. The controls for the powered flying controls, artificial feel, artificial feel lock, auto-stabilisers and mach trimmer are grouped together on a panel on the port console. The elevator and aileron trim and feel relief switches are duplicated on the two control

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columns, while those for the rudder are on the fuel contents panel. The emergency trim control is on the forward end of the retractable console.

4. At the top of the pilots' centre panel is a bank of warning lights and magnetic indicators. The three left-hand magnetic indicators are for the PFC units, the artificial feel and the auto-stabilisers respectively. The amber MAIN WARNING lights at either end of the group come on if a fault develops in any of the systems (except the yaw damper) ; this warning is cancelled by pushing in the button of the channel concerned. The appropriate magnetic indicator shows white, as a reminder that a channel is unserviceable. The main warning lights are then available for any subsequent failure.

5. Below this group is the control surfaces position indicator, representing a view of the aircraft from the rear. There is a separate indicator for each of the control surfaces, with datum lines to show the surface position relative to the take-off position.

6. All button warning lights can be dimmed by rotating the bezel. With the exception of the mach trim reset buttons, which are blue, all lights are amber.

POWERED FLYING CONTROLS Elevons 7. Control of the aircraft about the longitudinal and lateral axes is achieved by eight elevons hinged into the wing trailing edge, four on each side. Each group of four is divided into two outboard and two inboard elevons. For identification purposes they are numbered 1 to 8 from port to starboard. Each surface is operated by a separate electro-hydraulically powered flying control unit (PFCU). Therefore, if a single unit fails,

• only one of the eight elevons is affected. Mter failure of a unit, the pilot is unable rto move its elevon which will slowly assume the trail position. ~

8. The dual interconnected control columns are operated in the conventional sense to control the aircraft in pitch and roll, any movement of the control column causing all eight elevons to move. To simplify subse­quent references, the term 'aileron' is used when referring to lateral control and 'elevator' when referring to longitudinal control.

9. Control column movement is transmitted to a mixer unit which, in turn, transmits the appropriate signals to the elcvons. Full elevator and full aileron travel cannot be obtained at the same time. In all cases, the movement of the outboard devons to the inboard elcvons is in the ratio of 5: 4.

Rudder 10. The single rudder is controlled by two powered flying control units, one main and one auxiliary. Normal control is by the main unit with the auxiliary unit idling ; changeover occurs automatically if the main unit fails. During ground checks, when the main PFCU is stopped a delay of up to 15 seconds occurs before the

auxiliary unit takes over. Since the control inputs to both rudder PFCU's are mechanically interconnected, a restriction of the input to one unit would prevent the pilot's demand being felt at the other. To prevent this, a trip mechanism is incorporated which, when activated, disconnects the faulty unit from the input linkage, gives failure warnings and, if the main unit is at fault operates the changeover mechanism.

Powered Flying Control Units (PFCU) • 11. Each PFCU consists, basically, of an electric motor

driving main and servo hydraulic pumps rutd a~

<hydraulic jack to move 1!he control surface. Movement of the cockpit control operates the assembly to supply fluid to the appropriate side of the jack, thus moving the control surface. When the control surface position coincides with the new position of the cockpit control, jack movement ceases and the control surface remains in the selected position until further control movements arc made. A stroke limiter prevents excessively harsh movements of the control surfaces.

12. Incorporated in the assembly is a surface lock valve. As long as servo pressure is available, the valve is held open to allow fluid to pass to either side of the jack. If, for any reason, this pressure is not available, the valve closes under a spring load and no further fluid can pass to or from either side of the jack. This prevents the surface from flapping in flight and acts as a ground lock. A bleed in each valve allows pressure on both sides of the jack to equalise slowly, thus allowing the surface to trail to the no-load position. When the valve doses it operates a micro-switch and a warning light on the pilot's panel comes on. There is only one control surface lock valve for the rudder assembly and it is housed in the auxiliary unit. An interconnection between the main and auxiliary units holds the lock voalve open until no servo pressure is avai]1l;ble from either unit. Indications of an individual rudder PFCU failure is given by a pressure switch in each unit.

Electrical Supplies 13. a. A 200-volt, 400 HZ AC supply is required to

operate the PFC motors. This is supplied from the main busbars, the distribution of loads being as follows:

(1) Elevons:

(2) Rudder:

Nos. 1 and 8, No. 1 busbar Nos. 3 and 6, No. 2 busbar Nos. 4 and 5, No.3 busbar Nos. 2 and 7, No. 4 busbar Main, No. 3 busbar Auxiliary, No. 2 busbar

b. PFC failure warning is operated by 28-volt DC.

Controls 14. The 10 push (off) spring-loaded stop buttons for the individual PFC units are arranged along the inboard edge of a panel on the port console, those for the elevons being grouped in pairs of devons. The inboard button of the rudder pair controls the main unit. Each

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• RESTRICTED AP IOlB-1902-15

Flying Controls

button incorporates a warning light which comes on if the unit malfunctions. Three PFC START pushbuttons, which also engage the feel systems, are at the rear of the panel and are marked A, R and E ; A signifies the outboard elevons and aileron feel, R the rudder and rudder feel and E the inboard elevons and elevator feel.

15. The PFC MI shows white if any PFC button is pressed when the servo pressure has fallen below 10 to 15% of normal and a 28 volt supply is avrulahle. If the system pressure is normal, the PFC MI shows white only while any PFC button is being pressed.

Starting the PFC 16. Willh electrical power available and with the PFC and artificial feel buttons out, the PFC motors and the artificial feel are started by pressing the START push­buttons in the order A, E, R. The PFC snart up in sequence (A: 1, 2, 7, 8. E: 3, 4, 5, 6. R: main, auxiliary.) Check that the lights in the buttons go out in aocordance with the table below and that tihe two left­hand MI on the centre instrument panel show black. Prior to take-off, set all the PFC units at the neutral position. After flight and before turning off the secon­dary power supplies, push in all .the control buttons. Those for the PFC units are spring-loaded to the out position.

PFC 1, 3, 2, 4 No main R aux R 7, 5 8, 6

Time Sees. 4 6 8 10

PFC Failure 17. If any elevon PFC unit fails, the light in its stop button comes on, the main warning (amber) lights come on and that section of the control surface indicator fails to respond to control movements. Press in the appro­priate button to switch off the 200 volt motor; the main warning lights then go out and the magnetic indicator shows white as a reminder that the unit has failed ; the light in the button remains on. Attempt to start the unit again by pressing the appropriate START button but, if unsuccessful, press the stop button again. Do not make any further attempts to re-start. The elevon cannot be moved and assumes a trail posicion.

18. If the main rudder PFC fails, its button light comes on and the main warning lig:hts come on. No in­dication is given on the control surfaces indicator, however, as the standby PPC takes over automatically. Take ·action as in para 17.

19. When the non-essential services are shed, the auxiliary rudder PFCU stops, provided that the main unit is serviceable ; only the appropriate button light comes on. When the non-essential services are restored,

the main warning lights come on and the auxiliary unit can be restarted, if power supplies permit, by pressing the rudder START button. If the main unit fails while the non-essential services are shed, the auxiliary unit can be restarted immediately by pressing the START button without having to restore the non­essential services. The main rudder stop button must then be pressed.

20. If a PFC unit fails or is stopped in flight, reduce speed to below 0 · 90M. The control surface should normally trail in the neutral position but, at high airspeeds or mach numbers, it may move to a down position ; as speed is reduced, the control surface should return slowly to the neutral position.

21. a. No undue difficulty should be experienced during general flying or landing with up to two surfaces stopped. Stick forces are heavier and the asymmetry of the remaining surfaces affects control harmonisation.

h. If two surfaces fail, on one side, Hmit tihe angle of bank to 20° in -turns and mainnain tihe circuit speed for weight until lined up on the final approach.

c. If a control surface does not trail in the neutral position, relieve the feel before starting the final approach.

d. In all cases of failure keep, the CG in the normal range.

22. Normally only one elevon PFC unit is stopped at a time for practice purposes. If two units are stopped they must never be adjacent (ftu~r 1Nmitation). It is recommended that No 1 or No 3 is stopped, as those can be restarted with nhe minimum delay. lt is further recommended that, when stopping or starting a PFC for practice puvposes, ~this should be done in 'level flight, so 'that all ·surfaces are at approximately the same position. If an inboard PFC is stopped, its associated pinch damper should also ibe stopped.

23. No unit may be stopped for more than 15 minutes for practice purposes. An overlap of 15 seconds must be allowed between starting one unit and stopping another.

ARTIFICIAL FEEL

Artificial Feel Units 24. As the flying control system is irreversible, aero­dynamic loads are not transmitted to the pilots' controls. To compensate for this lack of feel, artificial feel units are provided in the elevator, aileron and rudder control runs. Each unit is designed to give a suitable degree of feel for its particular control, the load on the pilots' controls varying with airspeed and/or control surface displacement. The feel units are positioned alongside the three control runs in the bomb bay and are basically electric actuators which move under the influence of

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speed change. The starboard pitot/static system provides the feel actuators with the airspeed signal. Movement of the actuator moves a follow-up potentio­meter which is compared with a warning potentiometer supplied with port pitot/static information. If the guard system senses a discrepancy equivalent to 30 knots of airspeed, rhe pilot is warned of a failure in that feel channel.

Rudder and Elevator Feel

25. The feel for each of these circuits is operated by a combination of spring-loading, related to control surface displacement, and electrical actuation, governed by airspeed, which varies the pilots' mechanical advantage over the feel spring. For the elevator circuit, the airspeed factor varies as the square of the airspeed, starting at 90 knots, with a maximum value of 460 knots. The rudder circuit varies as the cube of the air­speed, starting at 130 knots with a maximum value of 460 knots. In addition, a pre-loaded centring spring is fitted in the elevator and rudder circuits, to overcome the effects of friction in the control runs and the feel spring. The centring spring is on the input side of the feel unit and therefore provides a loading which is constant at all airspeeds but varies with the amount of control movement made by the pilot.

Aileron Feel

26. Aileron feel is not airspeed controlled, although the range of control travel is reduced with increased rurspeed. Stick loading is produced by resistance from torsion bars and, therefore, increases with increase of control displacement, the load being constant for any given angle. To prevent the pilot applying too great a control angle, variable stops in the feel unit decrease me range of movement as the wrspeed increases above 150 knots with a maximum value of 365 knots. This circuit is also pre-loaded to overcome the effects of friction.

Feel Relief

27. a. Elevator and rudder feel can be reduced to the minimum speed position by use of feel relief buttons. The aileron variable stops can also be withdrawn, thus allowing the full range of control movement.

b. If any part of the artificial feel system malfunc­tions or if, in an emergency, it is desired to regain low speed feel conditions, the artificial feel may be relieved using the appropriate button. The button on the control column relieves both the aileron and the elevator systems. The system in which relief is not required may be restarted by pressing the A start button for aileron feel or the E start button for elevator feel. Feel relief on the rudder is achieved by pressing the button on the fuel contents panel. To regain normal feel, the appropriate start button

should be pressed. Relieving the feel also removes power from the system and gives failure warning. When any part of the artificial feel system is relieved, care in handling the aircraft must be exercised if over-stressing is to be avoided.

c. Initial application of 28V to the aircraft causes feel relief in all three senses. d. ~ ~ The fiap covering the feel relief button on the stick is spring-loaded to closed.

Electrical Supplies 28. The actuators for the artificial feel are operated by 28 volt DC.

Controls 29. The three push (off) pull (on) buttons for the artificial feel warning systems are at the forward end of the panel and are marked FEEL A, R, E. In this case the letters are for aileron, rudder and elevator feel. Each button embodies a warning light; when the button is pushed in, the main warning on the pilots' centre panel is cancelled or inhibited for that channel but feel system operation is not affected.

30. The artificial feel indicator is either a 3-position indicator which shows black, white or ILS, or a 2-position indicator which shows black or white. It shows black during normal flight conditions and white if any artificial feel channel fails or is relieved. (Chapter 12 para 64 describes the indications when the autopilot is in use).

Starting 31. For starting, see para 16.

Feel Locking ~ 32. To prevent possible feel unit runaway after a

feel unit failure a locking facility is provided for all. three channels. It is controlled by a single guarded switch. When LOCK is selected, a green light comes on to show that no further movement of normal or relief actuators can take place until NORMAL has been selected, ahhougih failure warning is given if clle speed is altered by 30 knots from that 'at which LOCK was selected.

33. Not used.

34. If the speed is changed by more than approxi­mately 30 knots from the locking speed the main warning lights and the lights in the feel indicator buttons come on and the magnetic indicator goes white. If speed has been reduced, out-of-trim and manoeuvring forces are higher than usual. To prevent the main warnings coming on, push in the feel indicator buttons. The lights in the indicator buttons come on and the magnetic indicator shows white.

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RESTRICI'ED AP lOlB-1902-15 Flying Controls

35. Before the feel is unlocked, reduce speed to below 250 knots, trim out the control forces and then raise the feel indicator buttons. The main warning lights come on.

36. After unlocking the feel, ensure that all failure warnings disappear, the green light goes out and the feel forces are at their appropriate level before making large control movements.

Artificial Feel Failure

37. If a feel channel malfunctions, the main warning lights come on, the feel MI shows white (but see Chapter 12 para 64) and the appropriate feel indicator button lights up. Great care must be taken not to over-stress the aircraft. Disengage the autopilot, reduce speed to below 250 kt/0 · 90M and attempt to reset by pressing the appropriate start button.

38. If the warnings remain, cancel the main warning by pressing in the feel indicator button and engage the feel lock ; then, by careful use of the controls, determine whether the failure is to the maximum or minimum speed value and adjust the sortie profile accordingly. It is not practicable to detennine the position of the aileron variable stops.

39. If the failure is to the high speed condition, continue flight fur as long as possible without relieving me feel. If the failure is to the low speed condition, maintaining a forward CG assists in avoiding over­stressing. In either <:ase, when in the circuit and at pattern speed, unlock and relieve the feel and restart the serviceable Channel if applicable. Adjust me CG to minus 2 for landing if the elevator feel is relieved The autopilot must not be used with any feel malfunction.

40. After unlocking the feel, if a failure has occurred, the main warning lights and the appropriate channel indicator light remain on and the magnetic indicator remains white. Maintain speed below 250 knots or 0·90M and proceed as in paras 37 to 39.

41. Certain electrical faults can occur without any failure indications appearing. H at any time the feel system is thought to be faulty proceed as in para 37 to 39.

Description

42. Control forces felt by the pilot in flight are produced by compression or extension of the feel mechanism, in response to control movement or change of airspeed. Trim adjustment is made by varying the length of the control run between the pilots' controls and the feel unit using an electrically operated actuator which removes the load from the feel spring. Double­pole wiring and switching is used to prevent runaways.

43. The trinuning systems are duplicated as a precaution against failure and no warning indicator is provided. If the main systems become inoperative, however, the emergency system can be used. Trimming in the elevators sense is not permitted above 0 · 90M.

Controls 44. a. Eaoh pilot's control colUIDJi carries a double­

pole 4-way aileron and elevator trim button. The button--cover covers the two switches controlling the paired ttim systems. A catd:l on the forward edge of the button, when operated, opens the cover and allows each switch ro be opemted independently, to test the .systems. If 't!he 1st pilot and co-pilot attempt to trim in opposite directions, 1ihe circuit first selected operates and the other is ineffective.

b. Twin rudder trim switches are on the fuel contents panel, spring-loaded to the centre (oft) position. They are marked RUDDER TRIM, PORTjSTBD; both switches must be moved for the system to operate.

45. 11he emergency trimmer control on the retractable console is moved fure~nd~ for longitudinal trim, sideways for lateral trim and rotated for rudder trim;

~the button in the top of the control must be depressed during trim selections. ~

Trim Failures 46. If either of the relays in the main trim circuit sticks in the made position, the trim motor always runs in the same direction, regardless of selection. If a trim motor runs in the opposite direction to the trim selection made, centralise the trim switch and make all further trim selections on the emergency trim switch.

Electrical Supplies 47. Electrical supplies for the system are 28V DC.

AUTOST ABILISERS General 48. Pitch and yaw dampers in the elevon and rudder circuits improve the natural damping of aircraft oscil­lations. A mach trimmer counteracts the nose-down trim change at high mach numbers.

Controls 49. At the outboard side of the port console panel are controls for the autostabilisers and automadl trimmer.

a. A YAW DAMPERS NO ljoffjNO 2 switch. b. A RESET COMPARATOR spring-loaded button (for the pitch dampers and auto mach trim). c. Four PITCH DAMPERS push (ofi)/pull (on) buttons, each embodying a warning light (amber).

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d. Two AUTOTRIM RESET spring-loaded buttons, each embodying an extension indicator light (blue).

e. An AUTOTRIM ON/OFF pullfpush button, embodying a warning light (amber).

The autostabiliser magnetic indicator shows white if any system is switohed off or fails (but see para 52 for yaw damper failure indkations).

Electrical Supplies 50. The gyros and amplifiers in the system are operated by 115V, 3-phase, 400 Hz AC, while the servos, motOrs ~nd relays are operated by 28V DC.

Yaw Dampers

~51. The yaw damper system is duplicated, the YAW DAMPERS No 1/No 2 switch being selected as required. From each detector rate gyro, signals are passed to the actuator between the feel w1it and the rudder PFCU. Yaw damping is in operation at all heights when either channel is selected but is airspeed monitored; rudder displacement is constant up to 200 knots and then decreases as airspeed increases. The monitoring supplies are from both the port and stavboord pitot/s'tatic systems.

52. If a system malfunctions, no warning indication is given. The magnetic indicator, however, shows white when the switch is off or following certain power fuilures. The switch must be put to ·the off (centre) position after flight. ..

Yaw Damper Failures

53. .. ~ If aircraft yaw damping appears inefficient select ·the orher channel.

54. Runaway of the selected yaw damper may cause a single change of directional trim or a directional forced oscillation which is quite vigorous at high speeds. ~A runaway may be detected during pre-flight system

checks by unequal rudder trim deflection. Complete~ failure of the yaw damper results in a dutch-rolling tendency after a rudder or aileron disturbance.

Pitch Dampers 55. a. Pitch dampers improve longitudinal stability

at high altitudes and high mach numbers (above 0·90M).

b. There are four channels in the system, each one feeding to one of the inboard elevon PFC. A com­parator links the four channels and if any one cllannel differs in operation from the other three, the warning system operates.

c. The system is height-monitored and is inoperative below 20,000 feet. Above this altitude, the amplitude

of control movement increases with increase of altitude.

d. The pitch damper servos are electrically heated whenever the undercarriage is retracted, the heating current being connected by a microswitch in the port undercarriage.

56. The pitch dampers are energised by pulling out the selector buttons on the port console. The buttons may be pulled out at any stage in flight but the dampers are inoperative until the height switch permits their operation. The buttons must be pushed in after flight.

57. It may be necessary to use the RESET COMPARATOR button before all or some of the channel lights can be extinguished before take-off.

Pitch Damper Failures 58. If a fault develops on any one channel, the comparator detects the difference between that channel and the others ; the warning light in its button comes on, the magnetic indicator shows white and the main warning ligihts come on. Press the RESET COMPARATOR button. If rhe warnings disappear, disturb the aircraft sharply in pitch. If the warnings re-appear, or if they fuil to cancel after pressing the RESET COMPARATOR button, switch the pitch damper off.

59. A pitch damper straight runaway gives a barely noticeable change of longitudinal and lateral trim. An oscillatory runaway may produce a small forced pitching oscillation. Damping characteristics are satisfactory with two channels failed on one side.

60. Without adequate pitch damping, divergent pitching oscillations can occur above 0·90M. These short-period oscillations cannot be controlled by the pilot ; consequently if two or more pitch dampers should fail when speed is above 0·90M, the throttles must be dosed immed.ia~tely and a smooth pull-<>ut exe>CUted in order to reduce speed as quickly as possible. The airbrakes must not be used to reduce speed, as t!his may aggravate the oscillations.

Auto Mach Trimmer 61. a. The randem mach trimmer system operates

on the elevator control run, ·thus controlling all eight elevons. Signals are passed from two transmitting machmeters, separately fed by eaoh pi!tot-static system, through a follow-up system and an amplifier to the servos. The system is brought into operation by a height switch at 20,000 feet.

b. The mach trimmer applies up-elevon as the mach number increases above 0 · 88 ± 0 · 01 (200 series ·engines) or 0·87 ± 0·01 (301 engines). The amount of up-elevon applied is always the sum of the move­ment of the .two actuators. They shoUld be extended

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RESTRICTED AP lOlB-1902-15 Flying Controls

by the same amount. Full extensi~ of t:?otb act'l;lawrs represents three-quarters of the totai up-elevator movement available but this is only achieved at a mach number of approximately 0·96M, which is outside aircraft limits. A malfunction could result in full travel of the actuators at lower mach numbers.

c. An accelerometer control in each half of the system is set <to limits of 1 · 5 and 0 · 7 g. When g values exceed 1 · 5 the servos are prevented from e.'Ctending but arc still able to retract; when g values are below 0 · 7 the servos cannot retract further but are able to extend. The servos return to their normal

· function when g is between the limits, provided that there is no misalignment.

d. A comparator monitors the system and indicates to the pilot if there is any misalignment of the two servos. Independently wired switches are provided to reset the servos to the minimum position.

62. Bath macll trim channels axe energised by pulling out the single ON /OFF button on the port console. T'he blue lights in the RESET buttons are on whenever the servos are extended.

Automach Trimmer Failures 63. Substantial misalignment of the servos is indicated by the main warning lights, the magnetic indicator showing white and the amber light in the POWER ON button, (either one or both the blue lights also

~come on). If the misalignment and g occur simul­t"<lneously (eg, a runaway actuator) the actuators are 1ocked out and the mach signal held off so that no further movement is allowed. When the autopilot disengages due to spring strut collapse, Iock-<>ut of the AMT always occurs regardless of mach number. (S~e Chapter 12, para 85).

64. When failure warnings appear, reduce speed to 0 · 85 M or less and check whether the actuators retract (blue lights out). If the warning indications self-cancel, switch off the AMT and leave it off. If the warnings remain, swi-tch off the AMT (ro cancel main warning) and retract the actuators by intermittently blipping the RESET buttons. After any AMT failure, the system should not be used again. ~

65. A mach trimmer failure may cause a nose-up or nose-down runaway which can continue for as long as 6 seconds. Stick force during the runaway can be fairly heavy if it is not relieved by use of the normal trimmer.

66. If one channel fails to extend during acceleration to high mach number, push-forces are lower t:han usual ; failure to retract during deceleration may require a push-force to prevent the nose rising.

67. If either or both actuators extend when below 20,000 feet, the nmgnetic indicator shows whlte and either one or bovh blue ligbit:s come on.

AIRBRAKES General 68. The slat-type airbrakes in the main-plane, above and below the engine air intakes, are electrically operated by two motors, using 200V, 3-phase AC; the emergency motor is supplied by No 2 busbar, and the normal motor by No 3 busbar. The supplies to the airbrakes are disconnected if load shedding occurs.

(f). The airbrakes have three extended positions : a. Medium drag 35°.

~ b. High drag (UC up) 55° . c. High drag (UC down) 80° (STI/410 77° + 3°).

The transition from 5 5'0 to 80° (77'0 + 3 °) is automatic~ when the undercarriage is lowered but raising the undercarriage does not retract the airbrakes to 55" .

Control and Operation 70. The airbrakes are controlled by a ganged switch on the rear face of the throttle quadrant. The switch has three positions: IN, MEDIUM DRAG and HIGH DRAG ; the button in the centre of the switch must be pressed in before the switch can be moved to the HIGH DRAG position.

71. a. The airbrakes are operated by either of two electric motors, connected to torque tubes through differential gearing.

b. The normal motor is isolated when the NORMAL/EMERGENCY switch on the throtcle quadrant is selected to EMERGENCY; the emer­gency motor only is then used for all selections. The en1ergency motor is isolalted when the AIR­BRAKE ISOLATE EMERGENCY switch on top of panel 3P is selected to ISOLATE; the normal motor only is then used for all selections.

c. Without isolation, both motors drive any selection from IN, only the normal motor being used for subsequent selections. However, current operating practice is to use one motor only, usually the normal motor for all selections.

~ d. Pre-flight, the airbrakes are to be checked initially using the emergency motor. After retraction, providing the airbrake slats are flush or slightly proud of the aircraft skin, a further check is to be carried out using the normal motor.

WARNING. If, using the emergency motor, the air­brakes retract further than me Bush position, remain in EMERGENCY. A NORMAL selection may result in damage to the wing or airbrake mechanism. ~

72. The 3-position magnetic indicator for the air­brakes at the top of the pilots' centre panel shows blaok when power is on and me airbrakes are in; white when no power is available, when airbrakes are selected out or if the airJbrakes extend without selection

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or fail to retract completely. It shows cross-hatdled when me airbra.kes are in, -selected in, and a main con­tactor has welded (a momentary cross-hatched indication is given when the airbl"llkes are selected from IN to MEDIUM DRAG wirth the emergency motor isolated ; as soon as the airbrakes stallt to move 1!be indication reverts to white).

73. The airbrakes must not be operated on the ground when the bomb door access panels and the bomb doors are open.

74. The airbrakes must not be selected from HIGH to MEDIUM DRAG, or MEDIUM DRAG to IN, if iced up as this may cause damage to the drive medhanism or 'aillframe.

75. Sustained flight with airbrakes extended against engine power is not recommended.

Failures 76. a. L oad Shedding. After load shedding has taken

plare the motor whose electrical supplies are affected must be isolated before any other selection is made. When non-essential supplies are restored the air­brakes are then driven by the selected motor only.

~ b. All Other Failures. Select the NORMAL/ EMERGENCY switch to EMERGENCY, at the same time instructing the AEO to check/select the

AIRBRAK.E ISOLATE EMERGENCY switch to ISOLATE. Do not make any further airbrake selec- ~ tion unless flight safety considerations make- it imperative that the airbrakes are 1:etracted. (If this is the case and the airbrakes do not move when selected it may be possible to retract them by changing the control fuses in panel 3P).

BRAKE PARACHUTE

77. A brake parachute in the tail cone aft of t!be rudder provides additional braking during the landing run. Operation is electrically controlled (28V DC) by a split, 2-pole, JETTISON/STREAM switch on the centre instrument panel. Both halves should normally be used but either half of the switch can stream and jettison the parachute (at a slower rate). If both are used for streaming, both must be used for jettisoning.

78. A magnetic indicator, beside the external inter­com point, is visible when a small access panel on the starboard side of the rear fuselage is raised. The indicator shows black when the parachute door is locked and all switches and relays in the circuit are at their correct setting for streaming.

79. If unselected streaming occurs, the action of the door opening without electrical selection causes a supply to be fed to the jettison unit and the parachute is jettisoned automatically.

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• AP 101B-1902-15

PART I

CHAPTER 8-FUEL SYSTEM

Contents

General

FUEL TANKS AND RECUPERATORS Main Tanks Bomb Bay Tanks AAP·P Tank Tank Pressurisation and Venting Recuperators

REFUELLING AND DEFUELLING Air-to-Air Refuelling Refuelling on the Ground Defuelling

Para 1

8 12 16 17 23

24 27 32

FUEL SYSTEM CONTROLS AND INDICATORS Fuel Control Panels 35

38 43 47 51 54 59

Fuel Cocks Fuel Pumps Sequence Timers Fuel Contents Gauges and Flowmeters CG Indicator and Slide Rule Air-to-Air Refuelling Controls ...

FUEL SYSTEM MAN·AGEMENT Normal Use Air-to-Air Refuelling, CG Control Air-to-Air Refuelling, Transfer Procedure

MALFUNCTIONS Suspect Booster Pump ...

~ Non-Return Valve (NRV) Failure Sequence Timer Failure

64 70 74

78 83 ~ 84

Booster Pump Running Continuously at Full Speed ... Leaking Tanks ...

85 86

General

Leaking Fuel Line Amplifier Faults Electrical Failures

Dlustrations Main Fuel System Bomb Bay Fuel Systems Air-to-Air Refuelling System

1. Fuel is carried in 14 pressurised tanks, five in each wing and four in the fuselage, a:bove and to the rear of the nosewheel bay. The tanks are of the flexible bag type and each tank is enclosed by a metal casing which is pan of the aircraft structure. The tanks are not self-sealing but are crash-proof.

2. The tanks are divided into four groups, each group normally feeding its own engi-ne. A crossfeed system enables the various groups to be interconnected. Auto-

87 88 89

Fig 1 2 3

maoi:c fuel proportioning \is normally used to maintain the fuel CG position.

3. Provision is made for carrying two fuel tanks in the bomb bay, either saddle-shaped or cylindrical. Fuel from .these tanks passes into the main system through two delivery ·lines, one each side of the fuselage, to each side of the centre crossfeed cock.

4. An air-to-air refuelling probe is in the nose and pipes from it join the normal refuelling lines. ~ ~

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5. A pressure refuelling system is provided for ground use.

6. The majority of the controls and indicators for the fuel system are grouped in the form of a mimic diagram on the retractable console. The air-to-air refuelling controls are on the starboard console.

7. Electrical Supplies. All fuel system components operated from the aircraft electrical supplies are listed in Tables 1) 2 and 4 of Chapter 4. All fuel pumps utilise 200V AC but the fuselage tank pumps are controlled by 28V essential DC while the bomb bay and wing tanks pumps are controlled by 28V DC non­essential supply.

FUEL TANKS AND RECUPERATORS Main Tanks

8. The tanks on each side of the aircraft are num­bered from 1 to 7) No 1 and 2 being the fuselage tanks and the remainder the wing tanks. The tank numbers correspond to the CG position of each tank) No 1 having the furthest forward CG and No 7 the furthest aft. No 1, 4) 5 and 7 tanks comprise the outboard tank group (No 1 group port) No 4 group stariboard). Similarly) No 2) 3 and 6 tanks comprise the inner tank groups (No 2 port and No 3 starboard). Ea'Oh group normally feeds its associ'ated engine.

9. The maximum fuel capacities are given in the Leading Particulars (preliminary page 9).

10. The total contents) in gallons, varies between different fuels because the point at which the refuelling valves close is affected by the permittivity, specific gravity and temperature of the fuel used.

11. Each wing tank contains a reservoir, with clack valves at its base. These valves are normally open but close when the head of fuel in the reservoir is 'built up by the auxiliary pumps.

Bomb Bay Tanks 12. Two fuel tanks may be carried in the bomb bay, one 'at the forward and one at 1he rear end. The tanks can be of the saddle or cylindrical type ; the former being referred to as the 'A' when fitted forward or the ' E ' when fitted aft. If only one bomb bay tank is oawied it must be fitted in the forward position. Note: See Part 2) Chapter 1, para 20 for allowable bomb bay tank configumtions.

13. Saddle Tanks. Each saddle tank has four pumps) two on each side ; each pair of pumps feeds into the delivery line on that side.

14. Cylindrical Tanks. Each cylindrical tank has three pumps) feeding into a common line) which feeds both delivery lines. The tanks are referred to as the forward and aft tanks.

15. Tank Capacities. The approximate tank capaci­ties are as follows:

Tank Gallons lb Avtur lb Avtag

A 718 5744 5529 E 721 5768 5552

Cylindrical (each, usable) 995 7960 7662

AAPP Tank 16. The fuel tank for the airborne auxiliary power plant, in the Stal'board wing to the rear of the AAPP,

1-8 Page 2 RESTRICTED

•

- MAIN FUEL SUPPLY

-- REFUEl./ OEFUE~

e ELECTRICALLY-OPERATED COCK

® OEFUELLING COCK

...... TRANSFER PUMP ........ BOOSTER PUMP

• REFUELUNG VALV.E

+ NON-RETURN VALVE

-co REFUELLING CONNECTION

cp LP COCK - PRESSURE SWITCH

SEE FIG. 3

ENG. ENG.

1-8 Fig. 1 Main Fuel System

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AP lOlB-1902-15 Fuel System

1-8 Page 3 (AL 2)

KEY

FUEL SUPPLY 'TO ENGINES - - REF\II LLING

lfjJ llOOSTER PUMP

5J REF\IELLING VALVE

El NOH-RETURN VALVl

wfj ELECTI\ICALLY OPERATED COCK

~ MANUALLY OPERATED COCK

-o ~:~m"'

1-8

FROM PORT REFUELLING

POI NT-._. ~

......-E

t:::-

CENTRE CROSSFEED COCK

w

TANK A

T ANKE

CENTRE CROSSFEED COCK

di>A

rj ) ~ i

~ :.-

-·- """"' -.. - -FORWAAD TANK

• --1

) D c bJJ_ ·- --·- ....,_.

.. - liMQ

REAR TANK

1-8 Fig. 2 Bomb-bay Fuel Systems

Page 4 RESTRICTED

rt o-

UJ~ ~

SADDLE TANK INSTALLATION

CYLINDRICAL TANK INSTALLATION •

• RESTRICTED AP IOIB-1902-15 Fuel System

has a capacity of 10 gallons. The tank is filled from the main fuel system via a pipe line from the wing tanks of No 4 group delivery line, whenever the No 4 group wing booster pumps are running. In addition, the tank is supplied from the refuelling line. A float valve in the tank shuts off the supply when the tank is full.

Tank Pressurisation and Venting 17. Each tank group can be pressurised with air from its associated engine. The system maintains a pressure of 1·82 to 2·3 PSI above ambient in ~the tanks through­out the altitude range of the aircraft, so as to prevent loss by boiling off fuel at high altitudes and high fuel temperatures. Pressurisation also prevents negative differential pressure in the tanks, thus preventing risk of tank collapse. The bomb bay ranks are not pressurised.

18. Pressurisation of the main tanks is controlled by a switch on the air-to-air refuelling panel ; below the switch are four magnetic indicators, one for each tank group, which show black when the 'tanks are pres­surised. A switch on the centre console for bomb bay tank pressurisation is inoperative.

19. Combined inwardfoutward relief valves are fitted in all tanks and float valves in the wing tanks. The float valve closes if the fuel level rises, to prevent fuel flowing back into the pressurisation lines. The inward relief valve is set at t PSI and the outward valve at ! to 1 PSI. Two vent valves for each tank group, one in the bomb bay roof and one under the mainplane, give atmospheric venting when tank pressurisation is not in operation; when pressurisation is switched on, the valve is adjusted by a master control valve which, in conjunc­tion with an air valve, maintains a pressure of 1·82 to 2·3 PSI above ambient. In the event of overpressurisa­tion an outward relief valve in the vent valve relieves at 2·65 to 3·0 PSI.

20. The AAPP tank is vented to the maio system in No 4 tank, No 4 group.

21. The bomb bay tanks ~re vented into the fuselage tank lines.

22. a. When tank pressurisation is selected on initially, the pressure indicators may remain white until RPM reach approximately 80%; they may similarly revert to white after the landing run. b. Normally, with power on and pressure off, the pressure indicators show white. If electrical load shedding takes place, the tanks are pressurised regardless of selection and the indicators show black. c. During descent, with pressurisation on, and after landing when it is selected off, rumbling noises may be heard ; this is normal.

Recuperators 23. A 6-gallon recuperator for each tank group supplies fuel to the engines if there is a pressure drop in the fuel supply. Each recuperator is supplied with fuel from its own tank group and with air from its own engine. If the fuel pressure drops, the air pressure (at 6 to 10 PSI) forces the fuel into the engine feed lines. Sufficient fuel is available to supply the engines for approximately 10 seconds at full power at sea level, and up to 2 minutes at a·ltitude at cruising RPM. When negative g is removed, nonnal fuel flow is resumed and the boosrer pumps recharge the recuperator.

REFUELLING AND DEFUELLING

Air-to-Air Refuelling 24. Fuel from the probe flows aft on either side of the cabin, through non-return valves, to join the main refuelling lines. All tanks are refuelled at the same time, the rate of flow being approximately 4000 lbfmin.

~ 25. Not used.

26. Probe lighting is provided by two lamps in the nose supplied through separate circuits.

Refuelling on the Ground 27. Pressure refuelling of ~the aircraft is via two refuelling points in each main wheelbay. These two points supply a common refuelling line on each side, running to a refuelling valve in the sump plate of each tank. Refuelling instructions are detailed in the Flight Reference Cards.

28. During refuelling, each tank is filled to the same percentage of its capacity in order to maintain a central CG position. A selector in the port main wheelbay allows selection of quantities from 0 to 100% and operates through the electrical output of the contents gauge amplifiers.

29. A control panel at each refuelling point carries the switches for the system and indicator lights to show the progress of refuelling.

30. Only one tank in each group is .filled at a time, automatic changeover to the next tank taking place when the first tank is filled to the selected percentage. The order of filling is 1, 4, 5 and 7 in the outboard groups and 2, 3 and 6 in the inboard groups.

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Page 5 (AL 9)

PROBE

1-8

NON-RETURN VALVU

RESTRICTED

DETAILS OF Mk.B PROBE HEAD

PRESSURE REDUCING VALVE 1600/20 PSI

PRESSURE RE~IEF VALVE 60•70· PSI

Not<-RETURN VALVE

•

BLOW-<WF VALVE 2000 PSI

CHARGING POINT

TO PROBE

NITROGEN SYSTEM

1-8 Fig 3 Air-to-air Refuelling System

Page 6 RESTRICTED

•

RESTRICfED AP lOlB-1902-15 Fuel System

31. Bomb Bay Tanks a. The bomb bay tanks are filled from the port refuelling point. A double-level float switch is fitted in each tank. The bomb bay refuelling panel is in the pon wheelbay and carries a master switch and low and high level indicator lights for each tank. With the master switch ON, the refuelling valves open and the indicator lights come on. When the tank is almost full, the lower float switch partially closes the refuelling valve, reducing the rate of flow and extinguishing the green light. The high level switch completely closes the refuelling valve when the tank is full and extinguishes the red light.

b. The main tanks must be refuelled before the bomb bay tanks.

c. The bomb bay refuelling master switch must be selected off when refuelling is complete.

Defuelling

32. Main Tanks

a. When defuelling the main tanks, it is important to check that the rear tanks are defuelled before the forward tanks. The method of defuelling given below is designed to prevent ingress of air into the engine feed lines.

b. Connect the bowser to the ground refuelling points, open the defuelling cock and close the tank servicing cocks of Nos 1 to 4 tanks. With an elec­trical supply connected to the aircraft, set the AUTO/MANUAL switches to MANUAL. Set the bowser to suck, switch ON Nos 5, 6 and 7 booster

• pumps and defuel the tanks to 500 lb indicated contents. As each tank reaches 500 lb, close its ~ servicing cock and switch off its booster pump.

.. Defuel the Nos 1 to 4 tanks to 500 lb in the same ~ manner. Then open all tank servicing cocks and, with no suction from the bowser, switch ON all

.. booster pumps to remove the remaining 500 lb fuel~ from each tank.

33. Bomb Bay Tanks. The bomb bay tanks must be defuelled before the main tanks. With the main tank

.. servicing cocks shut, and the forward servicing cocks shut, open the servicing cocks of the rear tank, set the bowser to suck and switch on its booster pumps. When the fuel level in the rear tank falls to 500 lb, close its ~ servicing cocks and switch off its booster pumps. Then open the servicing cocks of the forward tank, switch

.. on its booster pumps and defuel to 500 lb. Then, with~ no suction on the bowser, open the servicing cocks of both tanks and switch on their booster pumps to remove the remainder of the fuel.

34. Defuelling with Unserviceable Booster Pumps. In either of the above drills, if the booster pumps are

.. unserviceable, the last 500 lb fuel must be drained off. ~

FUEL SYSTEM CONTROLS AND INDICATORS

Fuel Control Panels

35. Retractable Console. The fuel panel on the retractable console carries a mimic diagram of the system, including the bomb bay tanks. Forward of the diagram are three CG control switches, two FWD I AFT transfer pump switches, one for each side of the system, and one PORT JSTBD switch for use during air-to-air refuelling. On each side of the diagram are two AUTO/MANUAL switches, one for each group; these switches control sequence timing (para 4 7). In each tank on the diagram is an OFF jon pump switch (which controls both main and auxiliary pumps) and a CONT pushbutton for contents reading. The cross­feed cocks are represented in the diagram by three magnetic indicators, with OPEN/close cock switches to the rear of them. Four pushbuttons, marked NO-ENG, are provided for flowmeter selection. The bomb bay system diagram has two BOMB BAY/ MAIN switches, two ON/OFF pump switches for each tank and a pressurisation switch (inoperative).

36. Starboard Console. The air-to-air refuelling con­trols are grouped on a panel on the starboard console and consist of two probe lighting dimmer switches, a nitrogen switch, a main tanks pressurisation switch, four tank pressurisation magnetic indicators, an ON/ OFF split double-pole master switch, a refuelling indicator and a refuelling gallery pressure gauge.

37. Gauge Panels. The contents gauges for the main tanks, one per group, are on a panel forward of the throttles, while that for the bomb bay tanks is on a panel attached to the inboard guide rail of the 1st pilot's seat, with a pushbutton for individual tank selection.

Fuel Cocks 38. HP Cocks. The four HP cocks are opened by the initial movement of the throttle levers forward from the fully closed position. The sleeves on the levers must be held up to permit movement between the HP COCK SHUT position and the IDLING gate.

39. LP Cocks. The four LP cocks are electrically controlled by four guarded ON/OFF switches on the underside of the coaming above the pilots' centre panel. Each cock is fitted with a bypass through a non-return valve, which acts as a thermal relief for fuel trapped between the engine and the LP cock, when the cock is dosed. The LP cocks are supplied wi:th power from the vital busbar.

40. Crossfeed Cocks and Indicators

a. There are two wing crossfeed cocks, each con­necting the tank groups on that side, and a centre crossfeed cock between Nos 2 and 3 groups. The cocks are electrically operated by 28V essential power.

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Page 7 (AL 7)

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b. Three 3-position magnetic indicators show con­tinuity with the diagram lines when the cocks are open, discontinuity when the cocks are shut and cross-hatch when the cocks are at intermediate position or when no power is available.

41. Bomb Bay Tank Cocks. Shut-off cocks are provided in the delivery lines from the bomb bay tanks. These cocks are opened together if any bomb bay tank pump switch is selected on.

42. Ground Servicing Cocks. A manually operated cock in the delivery line from each tank is used during ground servicing. The cock-i>perating levers are so designed that the cover plates cannot be fitted into place unless the cocks are in the open position.

Fuel Pumps 43. Booster Pumps and Auxiliary Pumps

a. Each wing tank contains both a booster pump and an auxiliary pump ; the fuselage tanks have a single booster pump. The booster pumps are attached to the sump plate at the bottom of each tank and, in the case of the wing tanks, are inside the reservoir.

b. Starvation of the booster pumps with change of aircraft attitude is prevented by the auxiliary pumps at the inboard end of each wing tank. These pumps run whenever the booster pumps are switched on and supply fuel to the reservoir, thus maintaining a head of fuel for the booster pumps.

44. Bomb Bay Tank Pumps a. Each saddle-type (A and E) bomb bay tank has four booster pumps, one pair supplying each feed from the tank. The pumps run in parallel and each pump switch controls one port and one starboard pump in its tank.

b. Each cylindrical tank has .three booster pumps, feeding into a common line. The same controls are provided, the right-hand pump switch for each tank controlling the forward pump and the left-hand switch controlling the other two pumps. The tanks are labelled FWD and AFf instead of A and E.

c. When l'he BOMB BAY /MAIN switches are set to BOMB BAY, the main tank sequence timers are isolated and the main tank pumps run at reduced speed, provided also that the main tanks are selected to AUTO.

45. Transfer Pumps a. vransfer pumps in Nos 1 and 7 tanks on each side allow the fuel to be transferred in either direc­tion between these tanks, if it is necessary to adjust the fuel CG position. As both tanks are in the same group, transfer does not affect the group contents. Transfer is via the refuelling lines.

b. With a transfer pump switch at FWD, the refuelling valve of No 1 tank opens and No 7 tank pump starts and transfers fuel to No 1 tank;

placing a switch to AFT opens the refuelling valve of No 7 tank and starts the No 1 tank pump. The rate of fuel transfer is approximately 100 lbjmin when transferring from No 1 to No 7 and approxi­mately 50 lbjmin when transferring from No 7 to No 1; approximately 300 lb of fuel must be transferred to alter the slide rule index by 1 · 2. Mter transfer, check that the desired amount of fuel has been transferred. c. If, when transferring fuel, the receiving tank is full when the transfer pump is still running, a float switch closes the refuelling valve. d. The function of these switches is altered during air-to-air refuelling.

46. Fuel Pressure Warning Indicators a. Four magnetic indicators on the pilots' centre panel, below the JPT gauges, show black when the fuel delivery pressure to the engine is satisfactory, white when the pressure downstream of the filter falls below 5 PSI and black when there are no power supplies. b. The two bomb bay fuel indicators show black when the fuel pressure is sufficient and the cocks are open. They show white if the fuel pressure falls below 10 PSI. c. If it is necessary to close 'the LP cock after shutting down an engine in flight, the magnetic indicators do not always turn white immediately. The time taken for fuel pressure beyond the LP cock to fall below 5 PSI depends on the rate of decay of pressure through the LP cock bypass valve.

Sequence Timers 47. Because of the configuration of the aircraft, the fuel tanks are disposed forward and aft of the aircraft centre of gravity. It is therefore essential that fuel should be used at approximately the same rate from all tanks, in order to maintain the fuel CG position.

48. An electrically operated sequence timer on each side of the aircraft ensures even fuel distribution, by causing the main pumps in each tank to run alternately at full speed and reduced speed (the auxiliary pumps run continuously). The quantity of fuel pumped from any one tank during one cycle of the sequence timer (five minutes) is proportional to the tank capacity ; the distribution of fuel is thus maintained throughout the tanks. The sequence timer motors use 200 volt AC.

49. The sequence of tank feeding is as follows:

Outboard engine Inboard engine Period No tanks tanks

1 1 6 2 7 2 or 6 3 4 3 4 5 2

It will be seen that No 2 tank feeds twice in each cycle, as it is the largest tank.

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50. With all booster pumps ON and the AUTO/ MANUAL switches at AUTO, sequence timing is in operation ; to interrupt the sequence timing in any group, put the appropriate switch to MANUAL and, if it is desired to use fuel from any particular tank in the group, switch OFF all booster pumps which are not required. The AUTO/MANUAL switches should be put to MANUAL after flight, to de-energise the relays.

Fuel Contents Gauges and Flowmeters

51. Main Tanks Contents Gauges a. A capacitor-type system provides indication of fuel contents. The four gauges, one for each tank group, are on a panel forward of the throttle levers. Each gauge is calibrated with two concentric scales, reading in pounds X 1000. Normally, each gauge reads the contents of its appropriate group, on the inner scale. An individual tank reading is obtained on the outer scale by pressing the pushbutton in the appropriate tank position on the mimic diagram.

b. In no circumstances should two gauge push­buttons in the same group be pressed simultaneously, as this damages the instrument.

c. Four group contents gauges are at the navigator /plotter's position; there is no means of reading individual tank contents on these gauges.

d. The contents magnetic indicator for the AAPP is at the AEO's station.

e. An accurate 28 volt supply is needed to ensure correct fuel readings.

52. Bomb Bay Tanks Contents Gauge. The fuel gauge of the bomb bay tanks is on the 1st pilot's seat guide rail, facing towards the co-pilot. The total contents of both tanks are normally shown ; individual tank contents can be obtained by pressing the appro­priate pushbutton below the gauge.

53. Flowmeters a. A Mk 3 flowmeter system is installed. This system is designed to give the following approximate indications:

(1) Fuel consumption by individual engines (Jbjmin) (2) Total fuel consumption by all four engines (lb/min) (3) Total amount of fuel gone (lb)

b. Two indicators, one giving total flow and pounds gone and the other giving instantaneous flow for individual engines, are on the co-pilot's instru­ment panel, together with a FUEL FLOW /RESET/ NORMAL switch for resetting the total flow indica­tor. Selection of an individual engine flow is

obtained by pressing the appropriate engine push­button on the fuel system mimic diagram. The instrument continues to indicate the flow to that engine until another engine is selected.

CG Indicator and Slide Rule 54. A fuel CG position indicator, on the pilots' centre panel, indicates the CG of the fuel system (not of the aircraft). The instrument registers automatically when air-to-air refuelling is in operation but readings can be taken in other flight conditions by pressing the CG CHECK button. This button can also be used, when air-to-air refuelling is in progress, to regain contents gauge readings.

55. The instrument face has two arcs, one for each side of the fuel system ; each arc is divided into three sectors, a central green sector to indicate the safe range and red outer sectors marked NOSE HEAVY and TAIL HEAVY. The needles should be on or near the zero position if the fuel is correctly proportioned. The green sector covers a range of 60,000 lb ft, 30,000 lb ft forward of zero and 30,000 lb ft aft. If, for example> both needles were on the forward limit of the green sector, the fuel CG would be 60,000 lb ft forward of the zero position with equally proportioned fuel.

56. The instrument can be checked before Bight by pushing the CG pushbutton and observing any slight movement of the needles. If no movement is observed, it may be because of exact fuel proportioning. This can be checked by transferring fuel from Nos 1 or 7 tanks and checking the indicator for movement while pressing the CHECK button.

57. The bomb bay tank fuel is not inc'luded in the CG indication.

58. CG Slide Rttle. A slide rule for calculating air­craft CG is provided and stowed below the starboard console.

Note: The CG limitations take into account the shift caused by undercarriage retraction.

Air-to-air Refuelling Controls 59. The air-to-air refuelling indicator consists of the outline of the aircraft, with numbered lights in the approximate position of each tank. The lights in the indicator can be adjusted for day or night use by revolving the ring round the indicator. The lights come on when the valves open and go out individually as the tanks fill.

60. The master switch must be put ON before drogue engagement and must not be put off until con­tact has been broken. With either half of the switch

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ON, the refuelling valves (in all tanks which are not full) are opened, tank pressurisation is switched off, the fuel contents gauges are isolated and read zero, all lights in the indicator come on (except for full tanks) and the CG indicator registers automatically. As each tank is filled, a double float switch closes the refuelling valve and the appropriate light in the indicator goes out. ~ ~

61. 100% refuelling may be carried out, provided that the fuel system is depressurised. If pressure remains in the tanks, refuelling may be carried out provided that the tanker main pump is switched off. Contact must be broken immediately if the fuel gallery pressure exceeds 10 PSI.

62. The master switch must not be set OFF until contact is broken. The nitrogen purge switch is then set ON ; this action opens the nitrogen cock and the No 2 tanks refuelling valves and nitrogen pressure at 20 PSI forces fuel from the probe lines into the No 2 tanks. Contents gauging is regained when the master switch is OFF.

63. CG Control a. The aircraft CG can be controlled during air-to­air refuelling by three switches on the fuel control panel: the two switches which normally control the transfer pumps, for fore-and-aft control, and the PORT JST ARBOARD switch at the top of the panel, marked FR RECEIVER CG CONTROL, for lateral control.

b. When the refuelling master switch is ON, the transfer pump switches are disconnected from the transfer pumps ; setting them to FWD closes the refuelling valves in tanks 6 and 7, while setting them to AFT closes the refuelling valves in tanks 1 and 2. In each case, the refuelling valves remain open in all other tanks. If the lateral control switch is moved to PORT or STBD, the refuelling valves in the Nos 6 and 7 tanks on the opposite side are closed.

FUEL SYSTEM MANAGEMENT

Normal Use 64. General. Before starting engines, ensure that the LP cocks are open, check that the AUTO/ MANUAL switches are at MANUAL and check the fuel contents, booster pump operation and cross­feed cock operation. Leave the crossfeed cocks closed and switch on one booster pump in each group (magnetic indicators black). After starting the engines, the fuel pumps and AUTO/MANUAL switrches may be used as required to ensure the C of G remains within the 1imi~ for take-off. Before take-off, ensure all booster pumps on, transfer switches central, all crossfeed cocks closed, AUTO/MANUAL switches at AUTO and

BOMB BAY / MAIN switches to MAIN with bomb bay pumps OFF. Select tank pressurisation ON before take­off and leave it on until after landing (see also para 22). After landing, set all AUTO/MANUAL switches to MANUAL and switch off the booster pumps, except that a rear tank pump should be left ON for each run­ning engine (No 6 for inboards, No 7 for outboards). After shutdown switch all booster pumps off. The LP cocks are normaHy left open.

65. In Flight a. Normally the fuel system is set up as follows: all 14 booster ptunps ON, all four AUTO/MANUAL switches to AUTO and crossfeed cocks shut. Fuel balancing is maintained automatically thus ensuring that the fuel CG remains approximately constant.

b. The fuel system should be in a balanced state, namely all four group comenrs should read the same and individual tanks should correspond with their opposite tank to within 100 lb. If necessary, fuel should be adjusted to within these limits as soon as practicable.

c. Make systematic checks of all tank contents at frequent intervals. When checking contents, the buttons for corresponding tanks on opposite sides should be pressed simultaneously ; both readings should be approximately the same. Not more than one button in any one group may be pressed at the same time as damage to the mechanism will result. Fuel checks should be made on the following occasions at least (bur see d. below):

Before take-off At the top of clirnb/s At 30 minute intervals thereafter Before let-down/s Before joining the circuit pattern.

d. Events such as a stores hang-up, booster pump failure or tank leak, may necessitate a departure from automatic fuel balancing, in which case the CG position can be maintained or altered by use of the transfer switches or selective switching of the pumps. Before switching off any pump select that group AUTO/MANUAL switch to MANUAL. CG checks should be made on the occasions listed above, except that once per hour should be sufficient at altitude.

e. When the contents of any tank have fallen to 650 lb set the AUTO/MANUAL switch for that group to MANUAL and when the tank fuel level reaches 400 lb switch off .that tank booster pump.

f. Before switching off any pump in a group, that group AUTO/MANUAL switch should be set to MANUAL to ensure the remaining pumps run at full speed.

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66. Crossfeeding a. If an engine fails or is shut down for a long period, crossfeeding is necessary to maintain a lateral balance. The appropriate crossfeed cocks must be opened; the system can be left in AUTO but, because of pressure loss in the lines, the side with the engine not running feeds more slowly than the other. Select MANUAL and switch off the booster pwnps on the side containing less fuel until lateral balance has been restored.

b. If fuel feeds unevenly for a known reason, such as during a practice asymmetric circuit after take-off, or low-level flight where throttle misalignment can be the cause, the ensuing state of lateral imbalance must be removed as soon as possible. Select all the crossfeed cocks open and set the AUTO/MANUAL switches in the low groups to MANUAL and the booster pumps in those groups off. When the group contents all agree, select all 14 booster pumps on, all AUTO/MANUAL switches to AUTO and close the crossfeed cocks. Check individual tanks in corresponding pairs and, if necessary, balance indi­vidual tanks to within 100 lb by switching off the one with the lower contents with its group selected to MANUAL. When all corresponding tanks are equal, switch on all 14 booster pumps and all four groups to AUTO.

c. The crossfeed cocks can be used as required so long as there are no faults in the system. However, it should be remembered that no more than three engines should be fed from a single pump source, as in the case of draining a leaking tank.

WARNING: Do not crossfeed if the cause of the imbalance is not known. If the cause is a leaking pipe line, more fuel will be lost ; if it is due to a fault, balancing the fuel masks the indications for a further period.

67. CG Adjustments a. There are recommended index positions for certain phases of flight. These are :

Landing between minus 2·0 and +2·0 UC down

Low-level minus 1 · S and minus 2 · S UC up Flying for range between minus 2 · 0 and + 2 · 0

UC up Flying for astrofbombing between minus 2 · 0 and

minus 5 ·O UC up Landing with E feel failure (low speed) minus 2 · 0

UC down

b. These positions can normally be achieved by using fuel from individual tanks by selective switch­ing or by transferring fuel between Nos 1 and 7 tanks. If a large adjustment is required a combination of both these methods can be used. Where possible, a ratio of 2: 3 between .the Nos 1 and 2 tank contents should be maintained.

68. Minimum Fuel a. For Landing. Although all the fuel indicated on the contents gauges is usable, because of the wide distribution of fuel between tanks and the possibility of uncovering pumps while manoeuvring the aircraft, the recommended minimum fuel for landing is 10,000 lb.

b. Fuselage Tanks. To allow for load shedding when at low fuel states, the contents of each No 1 and No 2 tank must not fall below 800 lb and 1000 lb respectively. If for some reason a No 1 or No 2 tank reaches its miillimum figure, illhe appropriate group must be switched to MANUAL, me wing crossfeed opened and the booster pump of the low tank switched off.

69. Use of the Bomb Bay Fuel Tanks

a. Use the fuel in the bomb bay tanks as soon as possible after take-off. Before selecting the bomb bay tanks, check that all 14 booster pumps are on and that all four AUTO/MANUAL switches are at AUTO.

b. When established on the climb, open the wing crossfecd cocks, switch on both bomb bay pwnp switches for the tank(s) in use and then set the BOMB BAY/MAIN switches to BOMB BAY. (Check that the bomb bay magnetic indicators change to white when the first pump switch is put on and almost immediately revert to black as the pressure reaches 10 PSI, thus showing that the pmnps are working and the cocks have opened).

c. When the bomb bay magnetic indicators show white, or the tank contents indicate 400 lb whichever is earlier, set the BOMB BAY/MAIN switches to MAIN, switch off the bomb bay tank booster pwnps and close the wing crossfeed cocks.

Air-to-Air Refuelling CG Control

70. The fuel CG indicator must be monitored through­out the refuelling sequence, especially in the initial stages. If a tank refuelling valve fails to open, especially in a tank with a large moment arm, the appropria~ needle moves quite rapidly forward or aft ; this is the only indication of a valve failing to open. As movement of the needle towards the limit of the green sector becomes apparent, move the appropriate FWD j AFT switch on the fuel panel in the opposite direction from needle movement. Refuelling may be continued for as long as it is possible to keep the needles in the green sector; if it becomes impossible, contact must be broken immediately.

71. Normally, if all tanks are accepting fuel, the needles move to and fro within the green sector and no action need be taken apart from monitoring. If contact

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is broken before all tanks are full, the subsequent tank contents check, with ·the master switch OFF, may show considerable variations in tank percentages. MANUAL use of the fuel system then becomes necessary to balance the tanks.

72. Even if it is known that the tanks are only partially to be filled, monitoring of the tanks-full indica tor is still necessary; in practice, Nos 7, 5 and 1 tanks are the first to fill, in that order, and the other tanks may lag behind by as much as 25%. Any method of adjusting the fuel in individual tanks, in an attempt to make all tanks fill simultaneously, reduces the safety margin provided by this lag and increases the possibility of rupturing a tank if its valve remains open when all the others have closed. Such a failure could be catastrophic.

73. If the bomb bay tanks refuelling master switch in the port wheelbay is inadvertently left on or an electrical fault produces the same effect, the bomb bay tanks refuelling valves open in flight as the fuel contents decrease. This results, during air-to-air refuelling, in fuel entering the bomb bay tanks ; a careful watch must be kept on their contents gauges in order to keep the CG within limits.

Air-to-Air Refuelling, Transfer Procedure 74. General. ~ ~ The following description refers to refuelling from the Victor tanker.

75. Before Contact a. Check and note the fuel tank contents and the amount of fuel to be received and ensure that all tanks are in proportion.

b. Switch off the fuel tank pressurisation and ensure that the tanks depressurise. If the tanks do not depressurise, refuelling may be carried out but the tanker main pump must be switched off. If the fuel gallery pressure exceeds 10 PSI, contact must be broken immediately.

c. It is recommended that, when necessary, the airbrakes are used and power maintained between 80 and 90% RPM, where the best throttle response occurs.

d. Switch the H2S off (leave the NBC on), HF to standby or receive and active ECM to standby.

e. Switch on the master switch, checking that the fuel contents gauges zero, the CG indicator is functioning and the tank refuelling indicator lights come on. If a tank light fails to come on it may be because the tank is full, the float switches have failed to operate or a light filament has failed ; the defective bulb may be replaced by one of the spares provided in the instrument.

f. Refuelling valve failure is not shown on the indicator. Failure of a tank to fill can be checked on

the CG indicator or, after breaking contact, by checking the fuel contents. Alternatively, the button by the CG indicator can be pressed and, while it is held in, the gauges indicate.

g. Auto-stabilisers should be on but contacts may be made with any or all of the stability aids inoperative. This is not recommended except in emergency.

76. In Contact a. When the green light in the tanker bay comes on, the fuel gallery pressure gauge starts to register. With the tanker main pump running, the pressure should be 33 to 37 PSI; if the pump is inoperative, the pressure is 7 to 8 PSI and refuelling takes con­siderably longer than normal. These pressures apply only when refuelling from the Victor Tanker.

b. As the tanks reach maximum capacity, the associated lights on the ·tanks indicator go out momentarily but flicker on and off at irregular intervals as the proportioning system feeds fuel from those tanks to the engines.

WARNING : In no circumstances may the refuelling master switch be set OFF while in contact, whether fuel is flowing or not. If the switch is set OFF while fuel is entering the receiver at maximum rate, the flick pressure •resulting from the closure of all fourteen refuelling valves may result in serious damage to the pipe lines.

77. After Reftlelling a. When refuelling is complete, break contact, switch off the master switch, check that the fuel gauges are functioning normally and check that the tank indicator lights are out.

b. Operate the probe nitrogen purge system for three minutes, then switch on the fuel tank pressuri­sation system. During purging, the No 2 tanks indicators lights should be on and the fuel pressure gauge should indicate 2 to 4 PSI; an irregular muffled thumping may be heard.

c. If necessary, use fuel selectively to achieve correct proportioning.

d. Do not switch on the H2S for 10 minutes after breaking contact.

MALFUNCTIONS

Suspect Booster Pump

78. Malfunction of a booster pump is indicated by the contents reading of its tank reading higher than the corresponding tank in the opposite group, with all group contents readings equal. If malfunction is suspected switch the affected tank group to MANUAL and switch off all pumps except the suspect one; pump failure is confirmed if the fuel pressure indicator shows

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white, after allowing for recuperator discharge time. Switch another pump in that group on and the suspect one off and check its control fuse. If blown, replace and try again (change fuse once only). If the MI stays white, the pump must be left off and the fuel remaining in that tank must normally be considered unusable and should not be included in the recommended minimum total fuel for landing. Leave the group in MANUAL and adjust the CG by selective use of the individual booster pumps.

79. If, after waiting for the recuperator discharge time, the fuel pressure indicator remains black, it can be reasoned that the pump is still running but at a reduced speed. This could be due to a number of causes and at this point it must be decided whether the fuel is required or nol. If it is, leave the pump on and drain the tank, then switch off the suspect pump and remove its control fuse. If the fuel is not required, switch the pump off and adjust the sortie length accordingly, using selective switching of the other booster pumps for CG control.

80. If the booster pump fails in a No 1 or No 7 tank, the fuel from the failed tank may be transferred to the serviceable tank. Adjust the CG as necessary and, as far as possible, keep a similar percentage of fuel in all tanks, by selective use of the booster pumps.

81. Only in exceptional circumstances may fuel supply of an engine by suction feed only (no main booster pumps running) be used in an attempt to use the contents of a tank with a failed booster pump. If the engine speed is above 80% RPM and the aircraft above 29,000 feet using Avtag or 37,000 feet using Avtur, the engine may flame out; these heights are reduced to 18,000 feet and 23,000 feet if tank pressuri­sation is not operating. Even if an engine docs not flame out, operation with suction feed only at any altitude may result in damage to certain components of the engine fuel system and, therefore, must be reported after landing.

82. If a bomb bay tank pump fails, the remaining pumps should be able to supply the engines. If both pumps fail on one side (saddle tanks), when only one tank is in use, the pressure indicator on that side shows white. Open the centre crossfced cock and feed all engines from the remaining bomb bay pumps. If it is desired to have more than these two pumps in operation, leave the centre crossfeed cock closed and place the appropriate BOMB BAY/MAIN switch to MAIN, thus allowing normal fuel sequencing to take place for the groups selected to MAIN. Fuel balancing can be adjusted once the bomb bay fuel bas been used.

Non-Return Valve (NRV) Failure ~ 83. If a tank NRV in the fuel supply <line malfunc­

tions, fuel may enter the affected tank from the main

supply line; this is most likely to be noticed follow­ing a booster pump failure in that tank. Except in the unlikely event of a double NRV failure, a wing rank will only be filled by the other wing tanks of its own group, but a fuselage tank could be filled from any other tank, depending on booster pump and cross­feed cock selection. If following a booster pump failure the fuel contents of the affected tank are found to be increasing, it is likely that the NRV is faulty. Monitor the rate of increase of the tank contents to determine whether the tank will fill before a landing can be made. If the tank is allowed to fill to maximum capacity, venting occurs and fuel is lost overboard. In addition the position of the CG must be monitored. If the affected tank is a wing tank, select the group to manual and feed the engine from the fuselage tank, or open the crossfeed cocks as required. If the affected tank is a fuselage tank, switch off all the booster pumps in the group, close the c.Tossfeed cocks to that group and gravity feed observing the provisions of para 81. If the affected tank is No 1 or No 7, transfer fuel as required.~

Sequence Timer Failure

84. Failure of a sequence timer is indicated by the contents of one tank per group on one side falling while the remaining tank contents on that side remain constant. A possibility exists, due to timer failure during overlap periods, that three tanks could be affected on the failed side. Select MANUAL for both groups on the failed side and maintain an even fuel distribution by selective use of the booster pumps, using the serviceable side as a guide.

Booster Pump Running Continuously at Full Speed

85. If, due to a relay failure, a booster pump runs continuously at full speed when the group is selected to AUTO, the fuel level in that tank falls more rapidly than in the corresponding tanks on the other side of the aircraft. The group contents, however, remain similar to the other groups. Select MANUAL for the affected group and maintain an even fuel distribution by selective use of fuel pumps.

Leaking Tanks

86. If a tank is leaking, group contents will normally be lower than the other groups and the tank fuel level falls continuously while the level in the other tanks falls more slowly. Attempt to confirm the leak by scanning the under-surface of the aircraft through the periscope. To minimise fuel loss, use a leaking tank to feed up to three engines. Open the appropriate cross­feed cocks, select MANUAL for all tank groups and switch off all pumps for the groups, except that in the leaking tank. Monitor the fuel contents and, when the tank is nearly empty (approx 650 lb), switch on all the other pumps, select AUTO for all tank groups except

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the leaking group, which should be left in MANUAL, and then switch off the booster pump in the leaking tank when it is empty. If the leaking tank is No 1 or No 7, fuel may also be transferred to increase the drainage rate. During this operation, make any necessary adjust­ments to the CG by using the transfer pumps. Land as soon as possible.

Leaking Fuel Line 87. If a leak occurs in the fuel line to the engine, that group's contents will be lower than the other groups and all the tanks in that group will be lower than their corresponding tanks. To remove the risk of fire, shut down the engine in that group and, if practicable, shut down the adjacent engine on that side. To minimise fuel loss, switch the group(s) to MANUAL and switch all booster pumps off in the group(s) ; select tank pressurisation off. Land as soon as possible.

Amplifier Faults

88. a. If a fuel contents amplifier fails, a sudden indication is given. The individual gauge reading could start fluctuating or could indica1:e zero or maximum, but the group contents would differ from the corresponding group on the other side.

b. These indications could be caused by water in the undercarriage bays.

Electrical Failures

89. a. The following fuel system controls and indicators are de-energised if load shedding takes place (see Chapter 4, Electrical System):

(1) Transfer pumps.

(2) Flowmeter selection (the meter continues to indicate for the last selected engine).

(3) Fuel contents gauges and CG indicator.

( 4) Wing tank and bomb bay fuel pumps.

(5) Sequence timers.

b. Before reset action is taken, the wing and bomb bay tank pumps must be switched off (so as not to overload the electrical system) and the AUTO/ MANUAL switches put to MANUAL.

c. If bomb bay tanks are in use and load shedding takes place before the fuel has been used from these tanks, there is considerable rearward CG shift as fuel is used from the fuselage tanks. If it is essential to remain at an altitude where loads cannot be reset, it is recommended that fuel be used from No 2 fuselage tank only; in this way the CG shift is not marked and up to half an hour's flight (approx 6000 lb fuel) is possible before the CG reaches the

~ aft limit. (See Part 3, Chapter 3, para 35 to 37 for landing with the CG outside the recommended range).

WARNING: If either of the sequence timer control fuses fail, the related main group fuel pumps run at full speed. With the crossfeed cocks open, the affected group tends to feed more than one engine, especially when the BOMB BAY MAIN switches are selected to· BOMB BAY (as aU other main fuel pumps are then running at low speed). The contents of the affected group show a ·reduction compared to the other groups; the fault could be mistaken for a fuel leak. Identical symptoms occur if a AUTO/MANUAL switch is selected to MANUAL. ~

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PART 1

CHAPTER 9-GENERAL AND EMERGENCY EQUIPMENT AND CONTROLS

Contents

ENTRANCE DOOR, LADDERS AND CANOPY Para Entrance to the Aircraft 1 Ladders ... 5 Canopy, Description 7 Canopy Jettison 9

WINDSCREEN AND ASSOCIATED EQUIPMENT

Windscreen and DV Panels Windscreen Wipen Sun Visors and Anti-Flash Screens

LIGHTING

Internal Lighting External Lighting

MISCELLANEOUS EQUIPMENT

Sextant ... Ration Heaters ...

~ Periscope Rapid Start External Connections

EMERGENCY EXITS AND EQUIPMENT

Entrance Door and Static Lines Canopy Jettison First-Aid Kit Crash Axe and Asbestos Gloves Life Raft Signal Pistol and Cartridges External Emergency Equipment

ENTRANCE DOOR, LADDERS AND CANOPY

Entrance to the Aircraft 1. The aircraft is entered by the door on the under­side of the fuselage, below the crew compartment. The door is hinged at the forward end and opens downwards. Door opening can be operated either mechanically or pneumatically ; door dosing is pneumatically operated.

2. Door Opening Mechanism a. The door can be opened from outside by a handle near the rear edge of the door. Operation of the handle deflates the door seal and withdraws the door bolts; the door then opens under gravity.

b. Door opening from the inside is by a lever in a gated quadrant on the port side of the door, at the

13 16 19

21 25

31 32 33 34

39 42 43 44 45 46 47 ..

forward end. Movement of this lever to the gate deflates the door seals and withdraws the locking bolts. If the aircraft is on the ground, the door then opens under gravity. To select the EMER­GENCY door opening position it is necessary to move the lever to the fully forward position, in the aircraft sense, via the quadrant gate. The gate is negotiated by moving the lever outwards towards the port side of the aircraft, then continuing for­ward. In an emergency, the door may also be opened by a switch at the navigator/plotter's position.

c. An additional door-opening switch is provided, recessed into the edge of the tab1e between the navi­gator/plotter and the AEO and covered by a flap.

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The switch has three positions ON/OFF/ON, and either ON position opens the door through the same circuits as the navigator's switch.

d. Emergency operation of the door is described in Chapter 13.

e. The main door may jam closed if the seal deflation valve plunger is out of alignment. The seal can be deflated by pulling the door inflation pipe from its rubber connector on the starboard side of the seal, or ·by cutting the rubber connection.

3. Door Closing Mechanism a. The door may be closed from the outside by operation of a pushbutton to the rear of the door.

b. The door is closed from inside by a toggle mechanism, the handle of which is below a cover on the crew's floor, between the navigators' seats. The door is closed by pnewnatic pressure at 400 PSI and, when the door is closed and locked, the door seal is automatically inflated by nitrogen at a reduced pressure of 25 PSI.

4. Door Indicators. On the door operating quadrant there are two lights, one red and one green. The green light, marked DOOR SAFE, comes on when the door is locked; the red light, marked DOOR NOT SAFE, comes on whenever the door is unlocked. In addition, there is a magnetic indicator above the pilots' centre instrument pane~ which shows white when the door is unlocked. In the nosewheel bay is a green light which comes on when the door is locked.

Ladders 5. A folding ladder is attached to the entrance door. Before flight it is to be removed and strapped to the stowage on top of the bomb aimcr's window cover. On the occasions when this position is covered by extra equipment the ladder should be carried in the pannier.

WARNING: Due to the hazard it presents to evacuating the cabin in an emergency, the ladder must not be left in position on the entrance door.

6. A ladder, which can be stowed by lifting and sliding to port, provides access to the pilots' cockpit.

Canopy, Description 7. The canopy, of double skin construction, is attached to the fuselage nose section by six attachment points. A seal, inflated by pneumatic pressure, ensures an airtight fit. A schrader valve allows the seal to be inflated by handpump, for weather proofing, when the pneumatic system is being serviced.

8. Canopy Locking Indicators a. Two pointers, one on either side of the cockpit, indicate against an arc whether the canopy is locked or unlocked. The locked position is indicated by a small white segment, while the UNLOCKED range is indicated by a larger, red segment

b. A magnetic indicator, at the top of the pilotS' centre panel, shows white when the canopy is un­locked.

Canopy Jettison 9. Two yellow/black wire-locked canopy jettison levers are provided, one on each side of the cockpit, above the consoles. A pip-pin is provided to lock each lever in the safe position and, when on the ground, the pins arc inserted through the hole in the coaming rail to prevent inadvcnent operation. The pins must always be removed before flight.

10. A yellow external release handle is on the port side of the nose. When this handle or either of the jettison handles is pulled, mechanical linkages open the canopy attachment jaws and operate a torque ·tube to fire the jettison gun.

11. The canopy is also jettisoned when any of the seat handles are pulled. In this case, pulling the handle operates the following sequence:

a. A pneumatic valve is opened, which allows air pressure at 1200 PSI to pass to a jack. b. The jack operates a torque tube to open the canopy attachments and fire the jettison gun. c. As the canopy clears the aircraft, it operates a time delay unit which fires the ejection seats 1 second later.

12. Canopy Jettison Gun Safety-Pin a. A safety-pin, with pip-pin attached, is provided for the jettison gun sear at the rear of the canopy, in the cabin. The pin must be inserted in the sear after flight. Before flight, the pin must be removed from the sear and the pip-pin must be inserted in the adjacent jettison lever mechanism, to link the manual mechanism to the gun. b. Instructions on the use of these pins is given on two tablets on the perspex cover of the gun, together with a diagrammatic arrangement of the devices. The pins are insened through sliding panels in the perspex.

WARNING : Canopy jettison lever pip-pins must be removed and a check made that the canopy indicator remains black, before the jettison gun is made live.

WINDSCREEN AND ASSOCIATED EQUIPMENT

Windscreen and DV Panels 13. The laminated windscreen, embodying gold film heating, is divided into three sections.

14. Thermal demisting, which is part of the air­conditioning system, is dealt with in Chapter 1. Wind­screen de-icing is dealt with in Chapter 11.

15. DV Panels. A triangular DV panel at each end of the windscreen is hinged at the lower edge and opens inwards. The panel is released by pressing a catch in the handle at the top, then depressing the handle, pulling it inwards and sliding it back. No stowages

1-9 Page 2 RESTRICTED

RESTRICTED AP 101B-1902-15 General and Emergency Equipment and Controls

are provided. When replacing the panel, care must be taken to ensure that the balls at the base are fitted into the guide tube correctly, otherwise a serious cabin pressure leak can occur.

Windscreen Wipers 16. Windscreen wipers are provided for each pilot's windscreen panel and for the centre panel. The wipers are electro-hydraulically operated and the blades are of the parallel-motion type and Gre self-parking. The wipers for the 1st pilot's and centre windscreen share a common hydraulic system and electrical motor, while the co-pilot's has an independent system.

17. The wipers are controlled by two double-pole, 3-position OFF/FAST/SLOW switches, one for the 1st pilot's and centre windscreen panel wipers and one for the co-pilot's panel wiper. The 1St pilot's switch is on the left of his panel. The co-pilot's switch is on <the right of his panel. The wipers must never be used on a dry windscreen.

18. The wipers are operated by 28V DC.

Sun Visors and Anti-Flash Screens 19. Sun Visors. Sun visors are provided for the windscreen and side panels ; the side visors are sliding and the front ones are hinged at the top. They are attached to the lower edge of the canopy.

20. Anti-Flash Screens. Anti-Bash screens are pro­vided for the windscreen, side screens and crew windows. Those for the windscreens have slide fasteners; when not in use, they are rolled down and stowed. The screens for the side windows are sliding shutters.

LIGHTING Internal Lighting

21. Cockpit Lighting a. Panel Lighting. The pilots' instrument panels and the consoles are lit by a mixture of red flood, white flood (fluorescent) and U /V lighting, while the E2B compasses have red lamps. Pillar lamps are built into the front instrument panels. The red lamps are controlled by a series of dimmer switches on the cockpit walls, forward of the port and star­board consoles. One U /V dimmer switch is on each side console. The white flood lighting is controlled by two switches, one on the outboard side of each pilot's instrument panel.

b. Anti-Dazzle Lighting. Anti-dazzle lighting is provided and is controlled by a BRIGHT joff./DIM switch on the port of the fuel contents panel ; an additional OFF /BRIGHT switch is provided at the nav /plotter's position.

c. Wander Lamp. A wander lamp, with its own integral switch, is attached to the canopy, above the pilots' position.

22. Crew's Lighting. Both panel lighting and angle­poise lamps are provided for the crew members. This lighting is controlled by a series of ON/OFF switches and dimmer switches at the crew positions.

23. Servicing Lamps. Servicing lamps are provided in the bomb bay, wheelbays, power compartment and rear fuselage. The master switch for these lamps is in the starboard side of the nosewheel bay. In addition, there are sockets for inspection lamps, one on the front spar bulkhead and one on the rear spar bulkhead, in the bomb bay. The ·lamps are only operative while an external 28V DC supply is connected.

24. Electrical Supplies. The U /V and fluorescent lighting uses llSV 3-phase AC from the main trans­formers; the red flood and anti-dazzle lighting is 28V DC operated.

External Lighting 25. Master Switch. Before any of the external light­ing (except the servicing lights) can be used, the EXTERNAL LIGHT master switch must be put ON. This switch is on the inboard side of the starboard console.

26. Navigation and Anti-Collision Lights. Steady navigation lights are provided together with red rotating lights, one on the upper fuselage and two on the underside, below the engine air intakes. The control switch is marked NAV LIGHTS-STDY /FLASH. When FLASH is selected, the navigation lights are steady and the 1:otating lights operate.

27. Identification Light. The downward identification light is controlled by a single-pole, STEADY /off/ MORSE switch, on the inboard side of the starboard console. The switch is spring-loaded from MORSE to off..

28. Landing and Taxying Lamps. There is a com­bined landing/taxying lamp under each wing, the lamp being extended further for the taxying position than for landing. The lamps are individually controlled by two double-pole, 3-position, RETRACT /LANDING/ TAXI switches on the inboard side of the starboard console. The landing lamps incorporate a slipping clutch mechanism and blow in if the airspeed exceeds 180 knots. Once the lamps have blown in, <the control switches must be reselected to RETRACT and then to LANDING (with the airspeed below 180 knots) before the lamps will re-extend.

29. Probe Lighting. Two lamps in the nose of the aircraft light the probe for air-to-air refuelling. They are controlled by individual dimmer switches on the starboard console.

RESTRICTED 1-9

Page 3 (AL 10)

RESTRICI"ED

30. Electrical Supplies. The external lighting uses ~ 28V DC and the navigation and anti-collision lights

are essential loads, the remainder of the external lighting is non~ssential. ~

MISCELLANEOUS EQUIPMENT

Sextant

31. A pressure-tight sextant mounting is provided on each side of the canopy coaming. The sextant is held in either the retracted or the operating position without loss of pressure. A lever in the mounting allows the sealing plate to be opened only when the sextant has been inserted in the carrier tube. The sextant must not be removed until the sealing plate is closed.

Ration Heaters

~ 32. There are five ration heaters, one at each crew position. The 28V DC non-essential supplies are controlled by two switches at the AEO's position on panel SOP. One switch controls rhe pilots' heaters, the other the rear crew's heaters. Sealed food tins are not to be inserted in the heaters; no tin is to be heated continuously fur more rhan two hours.

Periscope

33. The rearward-facing periscopes are controlled by a handle below the ABO table, raised to select rhe upper periscope, lowered to select the lower one, and moved sideways to rotate l!he selected periscope. The peris­copes are heated by a llSV 1600 Hz, single-phase supply from No 2 frequency changer, controlled by an ON /OFF switch on the ABO's upper panel. ~

Rapid Start f:xtemal Connections

34. The aircraft is fivted with external pull-off con­nections, so that the air conditioning hoses, the electrical supply cables, the telescramble line and the static vent plugs are all removed automatically· as the aircraft moves forward.

35. Electrical Supplies. The 28V, 200V and true earth plugs are on a sloping bracket on the port side of the power compartment, accessible through a spring­loaded access panel in the underside of the port mainplane.

36. Telescramble. A telescramble and micftel socket are on the starboard side of the power compartment, accessible through a spring-loaded access panel. Only the telescramble socket is used for rapid take-off, a further socket being introduced to this line for use of the crew chief. The telescramble system feeds into the intercom station boxes.

37. Air Conditioning. The cabin conditioning hose connection is on the starboard side of the cockpit and the ventilated suit connection is below the engine intakes on the port side.

38. Static Vent Plugs. The static vent plugs are pulled out by a cable.

EMERGENCY EXITS AND EQUIPMENT

Entrance Door and Static Lines

39. The entrance door below rhe fuselage is the escape exit for 11h.e rear crew except in a crash -landing. It can be operated by either the EMERGENCY position of the lever on the port side of the door, or by a switch at the navigator I plotter position marked EMERGENCY DOOR OPEN, or by the switch at the edge of 11he table, or by a combination of these controls. See also Ohapter 13 para 6. The door can be safely opened in flight at speeds up to 220 knots.

WARNING: If the door is opened by either switch, the normal lever may not pass through the emergency gate. If it does not and there is a failure of the 28V supply, the door will close again under the action of the slipstream. The first man at the door must check that the lever is gated in the EMERGENCY position.

40. After decompression, when the door has been opened, the crew swing themselves out of the opening, using the handle on the back of the nav /plotter's seat.

41. Static lines for the rear crew and 6t!h and 7th crew members' parachutes are fitted in the aircraft oxygen hose assemblies. One end of each line is connected to a 'Strong point on <the floor beside the oxygen point, while the other end carries the parachute avtachment link. Wirhin each hose assembly, connected to the static line, is a switch and associated electrical wiring. As each crew member abandons the aircraft in an emergency, the pull on rhe static line operates the switch, illuminating one of the five blue 'crew gone' lights on the bo~om of the 1st pilot's instrument panel.

Canopy Jettison

42. The canopy can be jettisoned by pulling back either of the jettison levers (one beside each pilot). If possible, pull both levers simultaneously. The canopy is automatically jettisoned when any ejection seat firing handle is pulled. It can also be jettisoned on the ground from outside, by pulling the yellow painted handle on the port side of the nose.

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RESTRICTED AP lOlB-1902-15 General and Emergency Equipment and Controls

First-Aid Kit 43. A first-aid kit is srowed at the rear of the co-pilot's seat.

Crash Axe and Asbestos Gloves

44. A crash axe and asbestos gloves are stowed on the cover of t'he bomb aimer's window.

Life Raft

45. In addition ro the individual life rafts carried by each crew member, a life raft Type MSS complete with 'Survival equipment, is stowed at the rear of tlbe canopy, below the fairing, outside the pressure ca-bin. The DINGHY RELEASE handle on the forward face of the ·stowage container is inaccessible unless the canopy has been jettisoned. Pulling the handle releases and inflates

the life raft, which remains attached to the aircraft by a painter.

Signal Pistol and Cartridges

46. A pressure-tight mounting for the signal pistol on the cabin port wall, above and forward of the AEO seat has a stowage beside it for 12 cartridges. ~~

External Emergency Equipment 47. A compartment on the port side of the nose, opened from 11he outside, carries a first-aid kit, a crash axe, a pair of asbestos gloves and a BCF hand fire­extinguisher. Break-in markings, for access to the cabin and to the emergency equipment, are painted in ycllow on the outside of the fuselage.

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Page 5 (AL 10)

RESTRICTED AP 101B-1902-15

PART I

CHAPTER to-HYDRAULIC SYSTEM AND UNDERCARRIAGE EMERGENCY LOWERING SYSTEM

Contents

General

General ...

MAIN SYSTEM SUPPLIES

Reservoir Engine-Driven Pumps ... Hydraulic Pressure Gauge Operation of the Main System Hydraulic Power Pack ... Hydraulic Power Pack Controls ... Hydraulic System Faults

UNDERCARRIAGE SYSTEM

General ... Undercarriage Emergency Lowering System Undercarriage Malfunctions Nosewheel Centring and Steering

WHEELBRAKES

General ... Operation Brakes Accumulator Charging

BOMB DOORS

Bomb Door Opemtion

~AAPP SCOOP

Operation

ntustrations Main Hydraulic System ... Undercarriage System Brakes System ... Bomb Doors System

1. The main hydraulic system provides pressure for:

a. Undercarriage raising and lowering and bogie trim.

b. Nosewheel centring and steering.

c. Wheclbrakes.

d. Bomb doors opening and closing.

e. AAPP air scoop closing.

Para

1

5 7

10 11 14 19 21

27 34 37

46

48 52 54

57

60 ~

Fig 1 2 3 4

2. An electrically-operated hydraulic power pack (EHPP) may be used for operation of the bomb doors and for rechar&ing the brake accumulators.

3. A nitrogen system is provided for emergency undercarriage lowering.

4. Separate self-contained electro-hydraulic systems operate the powered flying control units (Chapter 7) and the windscreen wipers (Chapter 9).

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Page 1 (AL 9)

RESTRICTED

MAIN SYSTEM SUPPLffiS

Reservoir

5. The main system contains 12 gallons of fluid of which 2!- gallons are contained in a spherical tank set in the roof of the bomb bay, at mid-position on the port side. The combined filling of the main and EHPP reservoirs is through a combined charging point on the inboard wall of the starboard undercarriage bay. For replenishment the undercarriage must be down, bomb doors open, parking brake off, and the accumu­lators charged with nitrogen and hydraulic fluid. Indica­tions of a full system are given by excess fluid spilling from the overflow adjacent to the charging point and correct level indications on the sight glasses of the main and EHPP reservoirs.

6. To ensure that a positive head of pressure is main­tained at all altitudes, the reservoir is pressurised with air from Nos 1, 2 and 3 engines. A pressure reducing valve reduces the engine air pressure from 105 PSI to 15 to 18 PSI, while a blow-off valve opens at 22 to 27 PSI and closes at 16 PSI. Mod 2321 replaces the engine air pressure with nitrogen pressure from the

~radio pressurisation system (see Chapter 13 para 21 and~ Fig 4).

Engine-Driven Pumps

7. Three engine-driven pumps, one on each of Nos 1, 2 and 3 engines, draw fluid from the reservoir via filters. The pumps incorporate an automatic cut-out and, when idling, circulate fluid back to the reservoir through the main return line. The pumps are low pressure spur gear, high pressure radial piston pumps.

8. From the pumps, fluid is delivered via non-return valves to the main gallery, at a pressure of 3600 to 4250 PSI. In addition to supplying the various ser­vices, this pressure is used to charge the wheelbrakes accumulators.

9. Test points for the suction, delivery and return lines are in the port main undercarriage wheelbay.

Hydraulic Pressure Gauge

10. A triple pressure gauge is on the pilots' centre instrument panel. The left-hand arc shows the pressure in the main gallery, while the two right-hand arcs show the pressure in the two brake accumulators. Each needle is separately wired and fused and has its own pressure transmitter.

Operation of the Main System

11. Before flight, exhaust the brake accumulators resi­dual pressure ; then, while making the external inspec­tion, check that the accumulator inflation pressure is 2550 .! sg PSI.

12. When the engines are running, check that the main and accumulator pressures are between 3600 and

4250 PSI. When a hydraulic service is operated, the main pressure drops ; check that it does not exceed 4400 PSI after the operation and that it finally stabilises within the normal range of 3600 to 4250 PSI. In flight, disregard fluctuations within the normal operating range. If hydraulic main line pressure indicates divergent increasing fluctuations, return to base and land as soon as possible. Select undercarriage DOWN as soon as practicable and leave it DOWN, bearing in mind the increased fuel consumption with lowered undercarriage.

13. To check that all pumps are working, it is neces­sary, during ground running, to time the operation of the bomb doors using the normal selector. With all three pumps working, this should be not more than 8 seconds between selection of OPEN and the doors reaching the open position.

Hydraulic Power Pack 14. An electrically operated hydraulic power pack (EHPP), on the starboard side of the bomb bay, pro­vides emergency pressure for operating the bomb doors and for recharging the brake accumulators.

15. The unit consists of a 3-phase electric motor driving a pump in an 11 pint reservoir. The reservoir is prcssurised to 15 to 18 PSI by engine air. A bypass valve and combined delay unit prevent the motor from starting on Joad, giving 2 seconds off-load running before the pump starts to build up pressure. Post-mod 2321, the pack is pressurised by nitrogen at 15 PSI.

16. a. The power pack is filled from the main reservoir, via a non-return valve. If a fractured pipe causes a loss of all the main fluid supply, the power pack can be used to recharge the brake accumulators until the EHPP reservoir is empty (normally approxi­mately one charge). Bomb door operation is not limited as, with the emergency system in use, the

< shuttle valve (Fig 4) provides an alternative supply._ and return, via separate lines, to the EHPP reservoir.

b. A pressure switch prevents the power pack motor from stalling and overheating and a 28V hold-off device ensures that the hydraulic circuit cannot be operated under load until an external 200V supply is connected, or the AAPP is supplying the busbar, or unless the pressure-head heater is ON.

17. The pwnp delivers fluid at 2 · 5 galls/min at a pressure of 3500 :to 3900 PSI; the pump is cut off by a pressure switch in the electrical circuit to the motor, when the pressure reaches 3500 to 3900 PSI. The

~delay unit is automatically reset. ~

18. The power pack motor uses 200 volt AC from No 2 busbar.

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•

SHUT-OFF VALVE

1-10 Fig. 1 Main Hydraulic System

RESTRICfED

AP 101B-1902-15 Hydraulic System and Emergency Air System

ENGINE - AIR

PRESSURE

- MAIN PRESSURE

POWER PACK

- AIR PRESSURE

SUCTION

c:=J RETURN

POST-MOD 2321, THE SYSTEM IS \ I PRESSURISED WITH

BLANK I NITROGEN INSTEAD ._ OF ENGINE AIR.

80H8 DOORS

1-10 Page 3 (AL 2)

PORT SIDE ONLY SHOWN. STARBOARD SIDE SIMILAR. MAIN WHEEL

DOOR

MAIN SUPPLY

RESTRICTOR

MAIN WHEEL DOORS JACKS

MAIN WHEEL. UNIT

-- --1 I I

MAIN WHEEL DOWN LOCK

f

I I I I I I

RESTRICT OR

NOSE WHEEL DOORS

SELECTOR VALVE

I I I I

NOSEWHEEL UNIT

NOSEWHEEL STEERING AND CENTRING MAIN

PRESSURE REGULATOR

NOSEWHEEL CENTRING

,~,~

SUPPLY

VALVE ~r--,""43-..--.

TO STARBOARD MAIN WHEELS --- I I _.J. ___ ..J.. _____ _ I I

~~M~~~~R--------~-1 ~ t' •

NOSEWHEEL

1UNIT JACK

---/ STEERING

JACK

- HAIH SUPPLY

- UPLINES

- DOliN LINES

~ .. -----~~~~--llfa~j---i VALVES I-J ETTISON ~ B

VALVE 9 y JETTISON

VALVE SHUTnE VALVE

= RETURN EK£RGEHCY MI.

c }-iPRESSUR~-{ ) EMERGE.NCY t11f!.. 1.:...1 GAUGES I.:.) EMERGENCY .ttR 7'1<fltoc..bi'-f 3,000 PSI ,... ',;<Q~er1 3,000 PSI

NON· RETURN VALVE

MOSEWIIEEL STEERING AND CENTRING

MANUALLY-OPERATED

PRESSURE RELEASE VALVE

TO TRIPLE PRESSURE GAUGE

ACCI.MULATOR PRESSURE GAUGE

1-10 Page 4

PRESSURE REDUCING VALVE

4,000/2,500 PSI

PORT BRAKE UNITS

1-10 Fig. 2 Undercarriage System

SHUTTLE VALVE

MAIN SUPPLY

FROM EMERGENCY POWER PACK ..,

RETURN TO RESERVOIR

DUAL PRESSURE GAUGES IN NOSEWHEEL BAY

- MAIN SYSTEM PRESSURE

EMERGENCY POWER PACK PRESSURE

- BRAKE SYSTEM PRESSURE

- PRESSURE FROM PILOTS' PEDALS

RETURN TO RESERVOIR

EE NON-RETURN VALVE

1- 10 Fig. 3 Brakes System RESTRICTED

RETURN

t MANUALLY- OPERATED PRESSURE RELEASE VALVE

TO TRIPlE. PRESSURE GAUGE

ACCUMULATOR PRESSURE GAUGE

PRESSURE

<E- FROM CO-PILOTS ~ FOOT PEDALS

STARBOARD BRAKE UNITS

RESTRICTED AP lOlB-1902-15 Hydraulic System and Emergency Air System

Hydraulic Power Pack Controls 19. Bomb Door Operation. The hydraulic power pack is energised to supply the bomb doors when the bomb doors emergency switch is put to the OPEN or CLOSED position.

20. Brake Accumulator Charging. The brake accumulators may be charged from the power pack by operation of either of two START/ STOP switches, one on the pilots' centre panel and one in the nose­wheel bay. The switches are spring-loaded to the central position. With the switch moved to START and released, the power pack charges the accumulators in approximately 6 seconds and switches off auto­matically when the line pressure reaches 3500 to 3900 PSI, unless it has already been stopped by selecting STOP. While the brake accumulators are being charged, both normal and emergency bomb doors selection is inhibited by an automatic shut-off valve. To ensure that the delay unit is reset, select STOP after each operation.

Hydraulic System Faults ~Note: Reference must always be made to the Flight

Reference Cards on hydraulic faults. .~

21. If the main pressure gauge pointer moves to the top of the scale, it is likely that either the fuse has failed or the transmitter is faulty.

~ 22. If the main pressure gauge pointer reads lower than normal a hydraulic fault should be assumed, although a transmitter fault is also possible. Attempting to confirm a hydraulic fault by selecting the AAPP scoop open, then closed, is likely to be inconclusive.

23. If a brake accumulator pressure gauge pointer moves to the top of the scale, either a fuse failure/ transmitter fault has occurred or a hot air duct leak is increasing the pressure. If operation of the wheel brakes temporarily reduces the pressure, a hot air duct leak is probable.

24. If a single brake accumulator pressure gauge indicates less than normal, while the main pressure remains normal, a transmitter fault is most likely. If however, the main pressure gauge also indicates less than normal, a hydraulic leak is the most likely cause and, therefore, any remaining pressure in the main system and in the affected accumulator will be lost.

25. If all three needles indicate full scale deflection it is likely that the hydraulic system has over-pressurised. Power failure is only a remote possibility because of separate wiring and fusing.

26. If a hydraulic failure occurs leaving only one brake accumulator, symmetric but limited braking will still be available via the brake control valve. ~

UNDERCARRIAGE SYSTEM General 27. The undercarriage mainwheel units are four­wheel, eight-tyred bogies ; the nosewheel unit is twin­tyred and steerable. When undercarriage retraction is selected, the bogies pivot, to lie parallel to the main oleos.

28. Hydraulic pressure operates the undercarriage doors, extension mechanism, bogie trimmers and down locks, through electrically-controlled selector valves ; sequencing of the operation is controlled by micro­switches. Each main wheel is fitted w~th a hydrauli­cally operated down lock, the nosewheel has a mechanically operated down 1ock. All down locks are of the over-centre type and will remain in the locked position should hydraulic pressure be sub­sequently lost.

29. Undercarriage raising and lowering is controlled by an UP and a DOWN button on the pilots' centre panel.

WARNING: To ensure that the electrical contacts are made when the landing gear selector is operated, the UP or DOWN button must be pressed fully in.

30. When the weight of the aircraft is on its wheels, a micro-switch on each bogie is held open and an interruptor pin behind the UP button prevents it from being pushed in. When the weight is off the wheels, the bogies trail and both micro-switches (in series) close. Power is then applied to a solenoid which with­draws the pin, allowing the UP button to be depressed. This device may be overridden however, by rotating the flange of the UP button slowly and gently clock­wise, through approximately 60°, at the same time exerting positive forward pressure.

WARNING : In spite of •this safety device, the UP button must always be regarded as operative, as the protective devices may not function.

31. Do not use excessive force on the UP button when making a normal selection.

32. The undercarriage position indicator is on the pilots' centre panel and indicates as follows:

All wheels up and doors locked closed . . . No lights

Wheels unlocked

Wheels locked down

Three red lights

Three green lights

33. A flag indicator, marked U jC, is incorporated in the co-pilot's ASI and shows if the speed is reduced below 160 knots when the undercarriage is not locked down. The absence of the indicator must not be taken as proof that the undercarriage has locked down.

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Page 5 (AL 9)

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Undercarriage Emergency Nitrogen System

34. The emergency nitrogen supply for the main and nosewheels is contained in two separate bottles in the nosewheel bay. The bottles are charged to a pressure of 3000 PSI via charge points adjacent to two gauges on the starboard rear wall of the nosewheel bay. The two controlling valves are mechanically linl{ed and are operated by a single handle on the right of the throttle quadrant. The handle is guarded and wire-locked.

35. When the handle is pulled to its full exten:t, nitrogen from the bottles passes to shuttle valves and jettison valves, expelling hydraulic fluid from the lines and allowing nitrogen to pass to the jacks. The under­carriage then lowers, regardless of the position of the normal selector.

36. Selection of emergency nitrogen also isolates the normal selector solenoids, so that the undercarriage cannot again be retracted once the emergency selector has been operated. Nosewheel steering, however, is still available. The alternator CSDU oil system and engine bay (zone 2B) grotmd cooling system is in­operative when the undercarriage has been lowered by emergency nitrogen. In these circumstances after landing, a maximum individual alternator load of 20 kW may be maintained for a maximum period of 15 minutes subject to CSDU oil temperature, and engine speeds must be kept at idling so far as is practicable.

Undercarriage MaHunctions

Note: Reference must always be made to the Flight Reference Cards on undercarriage malftmctions.

37. If the UP button cannot be depressed normally, the emergency override facility may be used to raise the undercarriage in flight, on condition that the follow­ing requirements are met and it is essential to continue the sortie:

a. The bogies are in the trail position.

b. Excessive force must not be used to depress the UP button after rotating the override flange.

c. No attempt may be made to return the override flange to its normal position once it has been moved.

d. The safety dip must be fitted on the UP selector button before landing.

e. Operation of the emergency override must be reported in the F700.

Note: Should it be necessary to abandon the aircraft in flight, use the override facility if the UP button will not depress.

38. If, during pre-flight checks, it is found that the override has been operated, the aircraft must not be flown. No attempt may be made to reset the override when the electrical system is live.

39. Under normal operating conditions, if all com­ponents of the undercarriage do not retract completely after an UP selection has been made, make no further attempt to raise the tmdercarriage. Lower the tmder­carriage by use of the normal system, or on the emergency system, if damage or out of sequence operation is apparent and land at the normal weight.

~Note: The term 'out of sequence operation' refers to the operating sequence for each individual under­carriage tmit and its associated door, and not to the order of retraction or lowering of the undercarriage when taken as a whole. ~

40. If an undercarriage malftmction results in one leg remaining locked up after a DOWN selection, when the hydraulic pressure is normal, fuses are serviceable and there is no visual damage, re-cycle the under­carriage once only.

41. If, after normal DOWN selection, three green lights are not obtained, first check the main hydraulic pressure. If this is normal check the bulb change­over, the indicator fuse and the selector fuse.

42. If hydraulic pressure is low, or if no electrical fault can be fotmd, the tmdercarriage must be lowered by emergency nitrogen, after first re-selecting DOWN on the norma-l control. If practicable, operation of the emergency nitrogen should be left until the aircraft is over the terminal airfield, so that lowering may be visually monitored by experts and, if any part of the aircraft falls away, it can be recovered and examined. If a faulty sequence valve has closed a door before the wheel is fully up, use emergency nitrogen without operating the DOWN button; this bypasses the sequence valves. The DOWN button should not be used to lower the tmdercarriage following a double TRU failure.

43. If only one tmit remains up, mechanical failure may be indicated, in which case emergency nitrogen may not release the tmit.

44. If, after using emergency nitrogen, loud bangs and heavy vibration occur, it may be that the undercarriage structure has been damaged. The indication of throe greens may be suspect under these conditions.

4 5. If the tmdercarriage is successfully lowered using the emergency nitrogen system, a certain amount of hydraulic fluid will be lost from the hydraulic system. Stop the aircraft as soon as possible after clearing the runway and shut down the engines.

Nosewheel Centring and Steering 46. Nosewheel Steering

a. Nosewheel steering is hydraulically operated and electrically controlled. The nosewheel can move through 47t0 on either side of centre and movement is controlled by a pushbutton on the pilots' control column and by movement of the rudder pedals.

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•

• RESTRICTED AP 101B-1902-IS

Hydraulic System and Emergency Air System

-E~~[}t---- BOMB DOOR JACKS -----'0~~~,..

MAIN RETURN _ _.

NORMAL SELECTOR VALVES

MAIN -m;:::;:;;;;l FEED

NRV

OPEN

RETURN FEED EMERGENCY CLOSE

1-10 Fig 4 Bomb Doors System

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CLOSE

SHUTTLE VALVE

EMERGENCY SELECTOR VALVE (OFF)

EHPP -- FEED

EHPP .. _____ RETURN

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RESTRICTED AP 101B-1902-15 Hydraulic System and Emergency Air System

b. With the steering pushbutton depressed, rudder ~ pedal movement causes the nosewheel to move in the~

appropriate direction ; the operation of a drum switch cuts off the electrical supply to the selector valve when the selected angle is achieved.

c. As the nosewheel leaves the ground, a micro­switch de-energises the stop valve in the steering circuit. This allows a bypass valve to open, per­mitting flow from one side of the steering jack to the other. The centring jack is now the only unit exerting any force and the nosewheel is automatically centred.

d. A NOSEWHEEL STEERING EMERGENCY OVERRIDE- EMERGENCY/ NORMAL wire­locked switch is on panel 3P. If steering is not restored after landing, the switch can be put to EMERGENCY to override the microswitch referred to in sub-para c. If the switch is used on the ground, it must be set to NORMAL before the undercarriage is raised after take-off and retumed to EMERGENCY when the undercarriage is locked down for landing.

47. Nosewheel Centring. Nosewheel centring is hydraulically operated, the main delivery pressure passing through a pressure regulator valve direct to the centring jack. The centring system operates auto­matically when nosewheel steering is not in use, when the nosewheel is off the ground, or if the main hydraulic system fails.

WHEELBRAKES

General

48. A maxaret unit in the brake unit of each double­tyred wheel temporarily relieves pressure at the brake if a skid is detected on that wheel. The eight maxaret units operate independently of each other and only operate when the wheels are rotating. The following points must be remembered:

a. If brakes are applied before a wheel touches the ground, the wheel locks and the maxaret unit cannot operate.

b. To stop rotation of the wheels after take-off, it is necessary to apply brake pressure for at least 4 seconds.

c. When landing with a reduced amount of hydraulic fluid (after a line leak) maxaretting must be avoided, as fluid under pressure would be bled rapidly to the return lines.

49. The brake units are hydraulically operated, the main system pressure being reduced to 2500 PSI at the brakes. Two accumulators, charged to main line pressure provide instant response and a reserve of pressure for brake operation. These accumulators can be recharged by .the hydraulic power pack (see para 14). A failure of one accwnulator does not prevent

the other from supplying pressure to both sets of wheels. A drop in nitrogen pressure in one or both accumulators would be disguised as long as the main hydraulic pressure remains normal.

50. The pressure at the brakes is shown on two dual pressure gauges in the nosewheel bay and the pressure at the hydraulic accumulators is shown on the triple pressure gauge on roe pilots' centre panel. The accumu­lator nitrogen pressure gauges and charging points are also in the nosewheel bay, together with two manually-operated pressure release valves, for releasing any residual hydraulic pressure in the accumulators.

51. A parking brake is provided, which operates through a Bowden cable to open simultaneously all the hydraulic valves in the brakes control valve.

Operation 52. Brake selection is controlled by toe-buttons on the rudder pedals. The pressure delivered to the brakes is proportional to the force applied to the toe-buttons ; when this pressure is released, the relay in the brakes control valve closes and the fluid from the brakes is returned to the reservoir.

53. The parking brake is applied by rurning and pulling the lever on the left of the throttle quadrant.

Brakes Accumulator Charging 54. If it is necessary to recharge the accumulators in flight, or on the ground when the engines are not running, this may be done by the hydraulic power pack (see para 14).

55. To charge the accumulators, either of the START /STOP switches, on the pilots' centre panel or in the nosewheel bay, is put to START. This action opens the shut-off valve and admits fluid to the accumulator circuit. A pressure switch operates to cut off the supply when the pressure reaches 3500 to 3900 PSI, or the supply may be stopped at any time by moving the control switch to STOP.

56. Either the GPU or the AAPP must be feeding the synchronising busbar or the pressure-head heater switch must be on, to override the 28 volt hold-off.

BOMB DOORS Bomb Door Operation 57. The bomb doors are hydraulically operated. For normal operation, supplies from the main system are fed throught dual selector valves to the four door jacks, via a shuttle valve which forms part of the door lock­ing assembly. If the normal supply fails, the doors can be operated from the EHPP through the emergency selector valve, and the main system is isolated by move­ment of the shuttle valve in accordance with the power selection.

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Page 9 CAL 9)

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58. Controls and Operation a. The 1st pilot has two switches on the port console, labelled BOMB DOOR CONTROL NOR­MAL and EMERGENCY. The NORMAL switch is a rotary type with three positions: OPEN I AUTO/CLOSE. When OPEN or CLOSE are selected, the bomb doors operate at the time of selection. When AUTO is selected the bomb doors are under the control of the NBS and do not open until they receive the appropriate signal.

b. The EMERGENCY switch is a guarded double­pole three~position switch, labelled OPEN /NOR­MAL/CLOSE. When this switch is operated, the doors are opened or closed by supplies from the hydraulic power pack and the electrical supplies are cut off from the normal selector. This switch is inoperative if the power pack is being used to charge the brake accumulator.

c. The nav /radar has a single-pole BOMB DOORS OPEN /CLOSED switch. Operation of either tWs or the pilot's switch opens the doors but, to close them both switches must be operated. The bomb doors can also be operated by the emergency bomb jettison switch; see Chapter 3.

d. The shuttle valve shuttle is normally positioned centrally and emergency selection to OPEN or CLOSE moves it to one end or other of the valve; when an emergency selection is cancelled, the shuttle returns to its central position. lf the shuttle valve jams, the doors cannot be operated.

e. Whenever the EMERGENCY switch is used, make the same selection on the NORMAL switch before returning the EMERGENCY switch to NORMAL.

f. The following points are to be observed: (1) Before operating the EMERGENCY switch to any position (including NORMAL), the pilot's NORMAL switch should first be checked to ensure

that it agrees with the postnon of the doors ; the navigator's switch should then be checked CLOSED.

(2) If the doors fail to operate correctly, set the navigator's switch to correspond with the doors position. Set the pilot's NORMAL switch to correspond with the doors position. Set the EMERGENCY switch to NORMAL. Set the navigator's switch to CLOSED. If the system is serviceable, the doors should operate normally. (3) Random selections by different crew mem­bers must not be made.

~59. Indicator. The bomb door indicator is to the right of the alternator warning light. Either a round 2-position or a square 3-position magnetic indicator may be fitted. The 2-position magnetic indicator shows black when the bomb doors are closed and wWte at all other times. The 3-position magnetic indicator shows black when the bomb doors are dosed, white when they are fully open and cross-hatch when they are at an intermediate position or when there is no electrical supply.

AAPP SCOOP Operation 60. The AAPP scoop is controlled by the AAPP master switch and operated by a spring-loaded jack. When the master switch is OFF, main line pressure reduced to 1800 PSI closes the scoop and holds it closed against the spring pressure. When the master switch is selected to ON, a solenoid-operated valve is de-energised opening the hydraulic return line so that the spring pressure opens the scoop. If the 28V DC supply to the solenoid fails, the scoop opens. If the main hydraulic pressure reduces below 1800 PSI, pressure to the jack is maintained by an NRV. Eventually (approximately 90 minutes), the hydraulic pressure decays and the scoop opens. The circuits are so arranged that the scoop does not close until the AAPP has run down. ~

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•

EPP

2222 EPP SUPPLY

RETURN

1-10 Fig. 5 Blue Steel System

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AP lOlB-1902-15 Hydraulic System and Emergency Air System

FIN GAP DOORS

1-10 Page 11

RESTRICTED AP 101 B-1902-15

PART 1

CHAPTER 11-ICE PROTECTION (All text completely revised by AL10)

Contents

General

General

ANTI-ICING SYSTEMS

Airframe Anti-icing System

Fin Anti-icing System

Anti-icing Controls

DE-ICING SYSTEMS

Windscreen De-icing System

Refuelling Probe De-icing

ELECTRICAL SUPPLIES

Electrical Supplies

Dlustrations

Airframe Anti-icing System

Engine Anti-icing System

1. A rhermal anti-icing system provides protection for the leading edges of 'the wings and for the engine air intakes. Hot air from tlhe engines and cold air from individual int!akes are mixed and pass along the inside of the skin and are then exhausted to atmosphere.

2. In the engine, thermal anti-icing is provided for the compressor entry guide vanes, intake struts and bullet by means of hot air from the HP compressor through an onjoff electrically-operated valve. This anti­icing air exhausts into the LP compressor on 200 series engines and overboard on 301 engines.

3. Gold film heating is provided for the pilots' windscreen.

4. The air intake for the ECM air scoop has an elec­trically powered anti-icing system for the wpour cycling cooling pack (VOCP).

ANTI-ICING SYSTEMS

Airframe Anti icing System

5. Hot air, brnnched from each half of the supplies to me cabin heating system, is fed 1:0 a hot air valve at the root of each wing. An electrically-operated flap varies t!he area of the cold air intakes below each engine air intake to control the amount of cold air mixed with the hot air. The mixed air is passed along ducts inside

Para 1

5 8

9

13

14

15

Fig 1 2

the skin, round the engine air intake leading edges and 'Separator, and along the leading edge of the wing.

6. The heating air from the engine air inmkes is exhausted through an electrioaHy-<>perated extractor flap on the lower skin of the air intake, while the air from t!he wing duct flows rearwards through the wing, to exhaust through a louvre on rhe underside of the outer wing.

7. An overheat switcll operates at 165 + 5°C to dose the hot air wlve. When nhe ducting temperature drops to 160°C, the hot air valve returns to its original setting.

Fin Anti-icing System 8. Pin anti-icing is inoperntive, apart from the temperature gauge used to detect any hot air leak.

Anti-icing Controls

9. The anti-icing controls are grouped together on a panel at the rear of the starboard console. One group is for tlhe port wing <and engines and one for the star­'board wing and engines. The 'fin anti-icing' switch contrors the beating for ·the ECM ram air intake via two operative positions, ON (outboard) and OFF (centre). The ECM intake heater must not be used for more than 60 'SeCOnds on nhe ground and must always ·be selected ON when the airframe system is switched on. Anti-icing controls include:

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Page 1 (AL 10)

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a. An AUTO/OFF/MANUAL double pole switch.

b. A temperature gauge, reading from 0 to 200°C.

c. A manual heat control switch, labelled INC/ DEC, spring-loaded to t!he central neutral position.

I 0. The thermal anti-icing system is designed to pre­vent the formation of ice and must, therefore, be switched on before entering icing conditions, whenever icing conditions are anticipated or forecast, or if a rapid build-up of ice is seen on the windscreen or windscreen wipers. The system must be used for 301 engines before and during flight in cloud even if there is no visual evidence of icing. 'Jihe wing system must not 'be used on the-ground, except during take-off and on the landing run. The ENGINE AIR switches must be OPEN before hot air can be supplied to the air­frame systems.

a. With the AUTO/OFF/MANUAL switch at AUTO, the hot air valve is under the control of a wing leading edge temperature sensing element set to maintain a skin temperature of 10°C, and me cold Qir intake and extractor flaps move to the fully open position. After a 10 second thermal delay, me cold air intake flap is placed under tihe control of a duot sensor set to maintain a duct temperature of 150°C.

b. When MANUAL is selected, the oold intake and extractor flaps open fully. The manual heat control switch must first be held to INC until t!he temperature rea<:hes 140 ± 5°C, and subsequently adjusted to maintaln me correct temperature.

c. When AUTO or MANUAL is selected, engine anti-icing is also switched on, irrespective of ENGINE AIR switch settings. Engine anti-icing must be used on the ground in conditions of cold, damp weather (temperature + 3°C or below and the humidity is 90% or more or it reduces visibility to less than 1000 metres).

11. Two separate guarded ·ON/OFF switdles, one on either side of the manual beat control switches, enable engine anti-icing to be used wben the wing anti-icing is not in use.

12. Operating techniques for the airframe and engine anti-icing systems, wi'lfh the degree of protection avail­able, are derailed in Part 3, Otapter 6.

DE-ICING SYSTEMS

Windscreen De-icing System

13. The pilots have gold film windscreen heating designed to operate at 40°C; an overheat control cuts off the electrical supply when the windscreen tempera­ture reaches 55° -~- 5'°C, and switches it on again when the temperature drops by 15°C.

a. A 3-position switch on the co-pilot's panel allows selection of three levels of heat: LOW (250 wattsjsq ft), MEDIUM (500 wattsjsq ft) and HIGH (750 wattsjsq ft). A press-to-test warning light beside ·the switch gives warning of overheat. On the starboard console, three 3-position magnetic indi­cators, one for each windscreen, show NORMAL when the windscreen hearing is operating satis­factorily, 0 /H when a windscreen is overheating and cross-hatch when windscreen heating is off or isolated. Isolation switches for individual windscreens are on 't!he rear face of panel 4P (nav /radar position).

b. H the warning light comes on, switch off the appropriate isolation switch. Cleek that the magnetic indicator shows cross-hatch and that the light goes out.

c. If the windscreen is damaged (other rhan damage to the inner laminations) select COMBAT pressure immediately and LOW on the heat control. If overheat warning is also indicated, or if a higher setting is required on the serviceable screens, isolate the affected windscreen.

Note: When the ambient temperature exceeds + 30°C, IDGH should be selected during the Pre-Take-Off ohecks and MEDIUM at the top of the climb, to prevent windscreen cracking.

Refuelling Probe De-icing 14. Provision for de-icing the refuelling probe may ·be in some aircraft but the system is inoperative and the controlling switch must be off at all times.

ELECTRICAL SUPPLIES

Electrical Supplies

15. The anti-icing system uses llSV AC from port and starboard main transformers and 28V DC. The de-icing systems use 28V DC and 200V AC from No 2 busbar. 1f load shedding ocours, the supplies to the anti-icing are cut off and gold film reverts to iow heat.

1-11 Page 2 RESTRICTED

0 TEMPERATURE TO CABIN CONTROL VALVE

[I] EXHAUST AIR EXTRACTOR

0 ENGINE AIR CROSS FEED COCK

• TEMPERATURE SENSOR

1-11 Fig. 1 Airframe Anti-icing System

RESTRicrED

--AUTO SENSOR

rso·c

AP 101B-1902-15 Ice Protection

HOT AIR SUPPLY

COLD AIR SUPPLY

CONDITIONED AIR

OVERHEAT SENSOR 1bs!s·c lh:v

1-11 Page 3

ENGINE ANTI- ICING AIR EXHAUSTS INTO AIR INTAKE.

ENGINE AIR INTAKE VANES ANTI-ICED BY HOT AIR FROM NOSE FAIRING.

AIR INTAKE NOSE FAIRING AND OIL TANK HEATED BY HOT AIR FROM THE L. P. COMPRESSOR ENTRY GUIDE BLADES.

OIL TANK

CIRCUMFERENTIAL COLLECTOR RING THROUGH WHICH HOT AIR IS LED TO THE L.P. COMPRESSOR ENTRY GUIDE BLADES.

ELECTRICALLY OPERATED VALVE TO CONTROL HOT AIR BLED FROM THE H.P. COMPRESSOR DELIVERY CASING FOR ENGINE ANTI-ICING.

TEDDINGTON TYPE VALVE INTEGRAL WITH ENGINE.

1-11 Fig. 2 Engine Anti-icing System

1-11 Page 4

( :2.. ~ .s:cx:u~ .r; Jlo/ £./YY//f".c s V-?7/LA<)

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RESTRICTED AP 101B-1902-15 Ice Protection

3-position magnetic indicators, one for each wind­screen, show NORMAL when the windscreen heating is operating satisfactorily, 0/H when a windscreen is overheating and cross-hatch when windscreen heating is off or isolated. Isolation switches for the individual windscreens are provided and are on the rear face of panel 4P (navfradar position).

b. If the warning light comes on, switch off the appropriate isolation switch. Check that the magnetic indicator shows cross-hatch and that the light goes out.

c. If the windscreen is damaged (other than damage to the inner laminations) select COMBAT pressure immediately and LOW on the heat control. If overheat warning is also indicated, or if a higher setting is required on the serviceable screens, isolate the affected windscreen.

Note: When the ambient temperature exceeds + 30°C, HIGH should be selected during the Pre-Take-Off checks and MEDIUM at the top of the climb, to prevent windscreen cracking.

24. Bomb Aimer's Control. The bomb aimer has a NORMAL/OFF /EMERGENCY switch on his oxygen and intercom panel. With either NORMAL or EMERGENCY selected, a solenoid-operated valve opens and pressurised fluid is sprayed on tb.e window. The rate of flow is 4 pints per hour on NORMAL and 16 pints per hour on EMERGENCY.

Refuelling Probe De-icing

25. Provision for de-icing the refuelling probe may be in some aircraft but the system is inoperative and the controlling switch must be off at all times.

ELECTRICAL SUPPLmS Electrical Supplies 26. The anti-icing system uses 115 volt AC from port and starboard main transformers and also 28 volt DC. The de-icing systems use 28 volt DC and 200 volt AC from No 2 busbar.

27. If load shedding occurs, the supplies to <the anti­icing are cut off and gold film reverts to low heat.

RESTRicyEI) 1-11

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General

RESTRICTED AP 101B-1902-15

PART 1

CHAPTER 12-MILIT ARY FLIGHT SYSTEM, AUTOPILOT AND OTHER FLIGHT INSTRUMENTS

(All text completely revised by AL10)

MILITARY FLIGHT SYSTEM General ... Twin Vertical Gyros Director Horizons Twin Azimuth Gyro Unit Beam Compasses Annunciator Units MFS Selector ... Navigator's Controls Power Supplies ... Pre-Flight Checks Normal Flight Control Failures ...

AUTO-PILOT General ... Controls and Indicators ... Autopilot Manoeuvre Envelope Pre-Flight Checks Use in Flight Auto-ILS Approach

Contents

Use of the Autopilot after Engine Failure Malfunctions

OT HER FLIGHT INSTRUMENTS Pitot-Static System Pressure-Operated Instruments Standby Instruments Miscellaneous Instruments

FULL FUNCTIONAL CHECKS Military Flight System Autopilot

lllustrations Director Horizons Beam Compass .. . MFS System Diagram .. . Pitot-Static System

MILITARY FLIGHT SYSTEM

Para 1 7

10 17 20 29 30 34 37 41 44 48

54 60 66 69 70 78 82 84

88 93 99

101

107 132

Fig 1 2 3 4

1. The Military Flight System (MFS) consists of:

2. Basic information is displayed on the director horizon and the beam compass at each pilot's station. These two instruments, which replace the normal artificial horizon, gyro-magnetic compass, ILS indi­cator, PDI, Zero Reader indicator and selector and the autopilot heading selector, provide flight director signals for the pilots.

a. Twin aircraft attitude systems.

b. Twin compass systems.

c. Autopilot Mk lOB.

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Page 1 (AL 10)

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3. ILS signals can be fed into the system, the ILS localiser information being presented on the beam compass and the ILS glidepath information on the director horizon.

4. Flight director signals are fed into the director horizon, telling the pilot the attitude required to achieve the desired condition of flight. If the autopilot is in use, it is supplied with heading signals from the system.

5. Track control facilities enable the navigator to apply drift and variation signals to the heading signals. The beading signals can be supplied to the NBS. The heading reference system (HRS) signals can be fed to the director horizons.

6.. The two halves of the system function independ­ently but are monitored by comparator units.

Twin Vertical Gyros 7. A unit in the pressure cabin (behind the navigator's panel) consists of port and srarboard gravity-monitored vertical gyros, driven by the 115V AC supply. These supply pitch and roll signals to the port and starboard director horizons respectively. The gyros have free­dom in pitch of 85° and complete freedom in roll. The vertical gyro erects to within -!0 of the true vertical within 30 seconds of switching on. While each gyro is building up to its normal speed of 23,000 RPM, the flow of current is high to give fast erection. When the gyro reaches full speed, the erection rate decreases to the normal value of 2!'0 per minute in pitch and 5° per minute in roll. The lower erection rate in pitch counters acceleration errors ; turning errors are countered by introducing pitch/bank erection in turns.

8. Pitch and roll signals are electrically transmitted from the gyros to the director horizons, each gyro feeding its own director horizon.

9. Each system measures the aircraft attitude inde­pendently but the two are so monitored by a compara­tor unit that attitude warning flags show on both horizons if their signals differ by more than the equiva­lent of 10° in roll and 3t 0 in pitch. The flags may a.lso appear aiirer a power failure.

Director Horizons 10. Each director horizon is, basically, an artificial horizon on which the pitch and roll elements have been separated. The instrument may be a Mk 1 or Mk 2. The Mk 1 displays pitch changes of 40° above or below the horizon bar; the Mk 2 displays only 20° above and below.

11. The Mk 1 horizon has a compressed scale at the larger angles and only the first 10° either way are linear. The Mk 2 scale is linear throughout its range. Both instruments have complete freedom in roll, with a bank scale marked to 60° each side.

12. The components are: a. A horizon bar. b. A bank ringsight pointer and bank scale. c. A pitch ringsight pointer. d. A pitch scale. e. A glidepath pointer. f. An azimuth director pointer (ADP). g. A pitch scale setting knob. b. BEAM and glidepath (GP) flags and a pitch director indicator (P) flag. 1. An attitude warning flag.

13. Roll signals from the remote vertical gyro are fed to the horizon bar, which rotates in a conve.ntional sense to indicate bank. At the same time, the bank ringsight, operating at right angles to the horizon bar, moves over a scale to indicate the precise angle of bank; the scale is marked in 10° intervals up to 30° and then at 60°. The ADP, moving over the same scale, is normally controlled by a heading error signal, given by the difference between the actual aircraft heading and the setting of the heading index on the beam compass. When ILS is selected, the ADP is controlled by the resultant of heading error and beam displacement signals, or by heading error signals alone. In the PDI function, track error only is given. The relationship between heading error signal and bank displacement angle is such that a 10° error signal demands a zoo bank angle, except when ILS is selected; in this case, a 100 heading error demands a 10° bank angle. In the PDI function the relationship is variable but is normally set so that 10° of track error demands a 24° bank angle. The maximum bank angle demand is 30°. The demand is satisfied by applying bank in the indicated direction until the bank ringsight is over the ADP.

14. Pitch signals from the vertical gyros are fed to the pitch pointer which moves vertically over the pitch scale. Any one of four types of director horizon may be fitted, although the intention is to standardise by fitting the Mk 2 instrument eventually. The charac­teristics of the different instruments are as follows:

a. Director Horizon Mk 1 Type D. This instru­ment has a non-linear pitch scale which is not calibrated.

b. Director Horizon Mk 1 Type F. This instru­ment is electrically identical with the type D. How­ever, ad~brated markings are added on either side df :the pitch scale, ranging in 5° steps from zoo nose-up to 10° nose-d()Wn.

c. Director Horizon Mk 1 Type H. This instru­ment is identical with the type F in all respects except that a larger attitude failure flag is fitted.

d. Director Horizon Mk 2 Type B. This instru­ment !has a linear pitch scale movement. The

1-12 Page 2 RESTRICTED

ATTITUDE

FLAG

' SANK

AZ IMU TH DIRECTOR PO INTER

ATTITUDE FAILURE

FLAG

BEAM FLAG

Mk.l Type D.

AP IOIB-1902-15 Military Flight System, Autopilot, Flight Instruments

GLIDE PATH Fl.AG

PITCH Fl.AG

- HORIZON BAR

.-~~-----PITCH SCALE

SETTING KNOB

BANK RINGSIGHT POINTER

I""' '"" '"G

FLAG

HORIZON BAR

GLIDE PATH---------t--~--~------~-J[~~ PITCH POINTER

POl NTE R

BANK

AZIM UTH

POINTER

PITCH SCALE

P ITCH SCALE

SE T T ING KNOB

RINGSIGHT POINTER

Mk.2 Type B.

1-12 Fig 1 Director Horizons

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Page 3 (AL 10)

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athbrated markings range from 20° nose up to 10° nose down. The left hand markings are in 2 o steps and the right hand markings are in so steps.

15. The glidepath pointer, moving relative to the pitch scale, is controlled by ILS glidepath signals. Both the pointer and the pitch scale can be moved up and down together behind the pitch pointer, by various methods. With no glidcpath signals, the needle remains over the centre dot on the pitch scale. When the pitch scale is being servo-driven andjor a selection other than central is made on the pitch selector, the P flag shows. With the MFSJTFR switch on the co­pilot's instrument panel set to TFR, the glidepath pointer becomes the pitch director. The no-signal posi­tion is at the top of the pitch scale (fail-safe) and the pitch scale is servo-driven (continuous fast chase, P flag shows). TFR demands displayed by the pitch director give the pitch rate changes needed to maintain a selected height AGL (see Chapter 15).

16. The BEAM flag is permanently in view whilst adequate ILS localiser signals are received; the GP flag shows whilst adequate ILS glidepath signals are received. Inadequate radio signals are indicated by the appropriate flag pulsing.

Twin Azimuth Gyro Unit

17. The basis of each compass is an azimuth gyro below the 1st pilot's seat. Gyro limits are -+- 85° in pitch and roll. Erection is fully automatic within one minute of the transformers being turned ON. The gyros are powered by 115V AC from the port and starboard main transformers and provide azimuth sig­nals to the port and starboard compasses. When mag­netic monitoring is taking place, the arrows in the annunciator pulsate. If the DG facility is selected on a beam compass, monitoring is disconnected on that side, and the annunciator arrows centralise and remain locked. Monitoring is automatically disconnected during turns but the annunciator arrows continue to indicate monitoring current. Each amplifier unit contains a turn switch which, if the rate of tum exceeds 30° per minute, inhibits magnetic monitoring of its own com­pass; the two tum switches also operate in series to inhibit the compass comparator and initiate pitch/ bank erection.

18. Automatic synchronisation is at a rate of 3° per minute, but manual synchronisation at up to 720° per minute is achieved by pushing/turning a knob on the annunciator unit.

19. The amplifier units, beside the gyros, amplify and transmit alignment signals and magnetic data to all parts of the compass system.

Beam Compasses

20. Each beam compass is the remote indicator of its gyro-magnetic compass but also:

a. Acts as the heading monitor for the autopilot.

b. Acts as the heading selector for both director horizons and for the autopilot. c. Can show the displacement from a selected ILS localizer beam and the aircraft heading relative to that beam.

d. Can provide directional gyro information.

21. The components are: a. A rotatable compass scale. b. A heading pointer, with a miniature aircraft in the centre and a ringsight pointer at the tip. c. A heading index. d. Top and bottom datum marks. e. Radio-coupled range marks. f. A radio beam displacement bar and scale. g. A DG flag. h. A sense switch. i. A setting knob. J. A compass warning light.

22. Any movement of the aircraft in azimuth is shown by the heading pointer moving over the com­pass card. The compass card can be rotated by pushing in and turning the setting knob, to bring the desired heading against the heading index, which is normally set against a datum mark at the top or bottom of the dial. During setting the heading pointer moves with the card and continues to indicate aircraft heading.

23. The heading index can be moved round the com­pass card, by pulling out and turning the setting knob. An alteration greater than 180° should be done in two stages, especially when making procedure turns using the autopilot. A turn in the other direction is demanded if the tail of the heading index passes the heading pointer.

24. The radio beam displacement bar travels hori­zontally over a scale on the face of the dial to indicate deviation from the beam signals. The first dot on each side indicates -! 0 displacement. The beam bar cannot indicate more than 1° accurately. Provided that the localiser QDM is set at the top datum, the relative positions of the beam bar and the heading pointer indicate the relation of the localiser beam centre line to the aircraft. Radio signals are coupled with heading error signals and fed to the ADP of the director hori­zon and to the autopilot, only while the heading index

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RESTRICTED. AP 101B-1902-15 Military Flight System, Autopilot, Flight Instruments

QDM/Qt>R

SENSE KNOB ---++6.•

TOP RADIO COUPLED ---t--f---;~----?''---7""'""'r-----T­RANGE

-COMPASS SCALE ---+---.--e • -~:.. BOTTOM RADIO ---t----''-e----,:-::.~~-=-~=---=-~ ------:-'•=---+--+--- BEAM DISPLACEMENT

COUPLED RANGE SCALE

1-12 Fig 2 Beam Compass

is set to within 30° of either top or bottom datum marks (ie, within the radio-coupled range marks), and depending on the posicion of the sense switch. Nor­mally the radio and beading error 'iignals are in oppo­sition, so the aircraft is directed to tum until the signals are equal and opposite. This occurs when the aircraft is on the centre line or, if the aircraft was too far from the beam initially, when flying to intercept at 45<> to the beam (provided the heading index is set to top or bottom datum); as the aircraft enters the beam the pilot is directed to tum progressively on to the beam QDM or QDR as appropriate. If the aircraft settles down slightly off the centre line with a small heading error in the opposite sense and no demand on th~ ADP, the aircraft is drifting parallel to the centre line, which can be regained by placing the beading inde.x under the beading pointer and following subsequent azimuth demands. The aircraft should then fly along the centre line with the correct drift set.

25. The 3-position sense switch controls the coupling of heading error and localiser radio signals to the ADP of both director horizons and to the autopilot:

a. When pointing up, heading error and QDM localiser signals are fed to the ADP, if the heading index is between the top set of radio coupled range marks. QDR localiser signals (if selected by moving

the beading index between the bottom set of radio · coupled range marks), are suppressed from the ADP.

b. When pointing down, beading error and QDR localiser signals (if selected by the heading index) are fed to the ADP but QDM Jocaliser signals (if demanded) are suppressed.

c. When pointing horizontally, beading error and QDM or QDR localiser signals are fed to the ADP, depending on which type of signal is selected by the heading index.

26. A comparator switches on the compass warning lights when the beam compass readings differ by more than 5°. A thermal delay switch prevents operation of the warning lights until misalignment has been present for 20 to 45 seconds. If either compass is selected to DG, or the rate of tum exceeds 3Q<> jminute, the comparator system is inhibited and the lights do not operate. Certain power failures may also cause a light to come on.

27. The DG flag shows when the directional gyro function is selected on the annunciator unit.

28. To avoid undue running of the compasses on the ground, compass isolation switches at the nav /plotter or AEO positions are used when ground testing other equipment.

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Annunciator Units 29. An annunciator unit for eacM compass includes a COMP JDG switch, a synch~nising knob and a window showing the annunciating arrows. To syn­chronise, depress and turn the knob in the direction of the annunciating arrow until the arrow pulsates. The knob is also used to set a DG output.

MFS Selector 30. An MFS selector on the centre instrument panel carries a COMP switch for compass selection, a navi­gation selector and pitch selector.

31. The COMP switch selects the port or starboard compass to supply heading error signals to the ADP of both director horizons and to the autopilot. All heading selections must be made on the compass selected.

32. Navigation Selector. The navigation selector has five settings:

a. BOMB. Depending on the position of the nav/ radar VIS/BLIND switch, steering signals from the NBS or the bomb aimer's tum control are fed to the MFS. Normal heading error signals are suppressed. b. REMOTE. The ADP show only MFS heading error signals. The navigator can feed in variation and/or drift so that the beam compasses read mag­netic heading/track or true heading/track. c. With the switch central, all heading pointer indications are magnetic. MFS or HRS heading errors can be shown by the ADP depending on the MFSJHRS switch setting. d. LOC. Provided that the ILS is tuned in and operating, localiser beam signals are fed to the ADP if settings of the heading index and sense switch permit. BEAM flags on the director horizons show and the beam bars on the beam compasses indicate aircraft position relative to the beam (a true plan view of the approach is given on the beam com­passes provided ILS localiser QDM is set to top datum). Whether ILS is tuned in or not, selecting LOC gives 1 : 1 bank demand. e. GP. All LOC facilities are retained. The GP flags show on the director horizons Wld clle glideparh pointer moves relative to the centre dot to show angular displacement of the aircraft (represented by the centre dot) from the glidepath, whatever the position of t1he centre dot. Provided that DATUM and APPROACH have been selected on the pitch selector, the glidepath pointer also shows pitch demands to regain or maintain the glidepath.

33. Pitch Selector. Controls the servo-driven func­tions of the director horizon pitch scales. With any selection other than central the pitch (P) flag shows on the director horizons.

a. MACH. Inoperative.

b. HEIGHT. Pitch directions are given to main­tain the aircrnft at the altitude at which it was flying at the -time of selection (see para 36).

c. Centre (Normal). No pitch director signals are fed to d:te director horizons and the pitch scale can be adjusted to any required attitude datum, using the pitch scale setting knob:

(1) Emergency Setting (Mode 1). With the knob pulled out, the pitch scale can be manually adjusted up or down to indicate a required atti­tude. Only the scale of the instrument so operated moves, the other instrument being unaffected. If the knob is left pulled out, the scale cannot be servo-driven on that instrument.

(2) Trimming Setting (Mode 2). In its normal position, the knob has restricted movement against a spring in either direction but rotation causes the scale to move up or down slowly, and the P flag appears, until the knob is released. Operation of the knob in this fashion affects both instruments but this function is inoperative if any automatic pitch function is selected and if fast chase is selected on the other director horizon.

(3) Fast Chase (Mode 3). With the knob held in against spring tension, both scales move rapidly to align the centre dot with the pitch pointer. While the scale is moving, the P Bag appears. This facility is not accurate if the aircraft angle of climb or dive is greater than 10°. This method of operation is particularly useful when climbing away from an overshoot; it overrides any setting of the pitch selector, releasing it to its central position.

d. APPROACH and DATUM. The DATUM position is spring-loaded to the central position. The APPROACH position can only be retained if the navigation selector is set to LOC & GP.

Navigator's Controls

34. A track control unit (TCU), Type A or C, can supply variation andjor drift to the compass system automatically or manually. A compass selector on the TCU enables the navigator to select which MFS com­pass supplies information to the TCU and naviga­tional equipment (GPI, NBS); the selected compass must not be set to DG by the pilot unless the navi­gator is fully aware of the selection made. True track or heading is displayed on the navigator's compass repeater (NCR), or magnetic heading after depressing the TCU spring-loaded MAG HDG switch. With the Type A TCU, only port compass data is displayed on the NCR, although the starboard compass can still be selected as the heading source to the navigational equipment.

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PORT COMPASS

AMPLIFIER

r-----, I I I I I I I I I I

r----l I I I I I I I I I I I I

STBD I COMPASS!

AMPLIFIER I I I I

I STBD I

DETECTOR I

I I I L ___ _j

I MAG L _ TRACK_CONTROL _

UNIT

NAVIGATOR'S REPEATER

[

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~ 1 - 12 Fig 3 MFS System Diagram~

(Editorial Correction)

RESTRICTED

PORT 1---~-----l VERT I CAL

AUTO BOMB

GYRO

STBD VERTICAL

GYRO

AP 101B-1902·15 Military Flight Systt;m, Auto-pilot,Flight Instruments

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36. Pitch director functions are fed into the system when the pitch control on the MFS selector is moved from its central position. The functions are:

a. With HEIGHT selected, the pitch scale is servo­driven up or down, the position of the centre dot indicating corrections needed to maintain the pressure altitude as determined by a manometric unit. A pitch integrator modifies these corrections to account for varying speeds and AUW and to prevent over­controlling. As the variation of the pitch scale is relative to its position at the time of selection, allow the aircraft to settle to the required cruise conditions and operate fast chase before selecting HEIGHT. To avoid undue attitude changes the height facility disconnects automatically if deviation from the selected altitude exceeds 13 mb ( 400 feet at sea level). The selector then returns to the centre posi­tion. If HEIGHT is selected before the aircraft has settled to its cruising attitude, any signals followed in the interim result in a distorted flight path.

b. The DATUM position of the switch should be selected at glidepath interception, as it moves the pitch scale to a position such that the centre dot corresponds to the attitude for a 3 ° glidepath under average conditions, thus giving a reasonable demand from the glidepath pointer. When released, with LOC & GP selected, the switch returns to APPROACH. At APPROACH, the centre dot moves slowly in the direction of any pitch pointer movement, ie, it modifies the DATUM setting in keeping with the necessary corrections for existing windjspeedjweight conditions. This process is termed pitch integration. APPROACH also provides an auto-drift facility, which is designed to cope with small changes in drift down the glide slope. It achieves this by slowly wiping out steady heading errors, so that long term control of bank demand is due to localiser signals alone. APPROACH should not be selected unless the ADP are central and the heading error is less than 5o.

Power Supplies

37. The main system is operated by 115V AC at 400Hz. All system relays and control switches are operated by 28V DC.

38. Complete power failure to the system is indicated by:

a. The failure warning flags showing on both director horizons.

b. No compass annunciation and the system going dead.

c. No ILS BEAM or GP flag indications.

d. MFS pitch selector reverting to the central position.

39. Power failure to one side of the system is indi­cated by:

a. Failure warning flags on both director horizons.

b. No compass annunciation on dead side.

c. Starboard compass warning light if failure is in port system (no compass warning light if failure is in starboard system until there is so discrepancy between indicated headings).

d. No BEAM and GP flags.

e. Reversion of pitch selector to central position if failure is in the starboard system.

40. Partial power failure within the systems may also be indicated by:

a. The failure warning flags on both director hori­zons when a power failure leads to a discrepancy between the signals to the horizons. Certain power failures can lead to an immediate appearance of the failure warning flag on one or both director horizons, before the failure has led to a signal discrepancy.

b. Compass warning light indication on one or both beam compasses, either immediately or when com­pass indications differ by more than 5° for more than 20 to 4 5 seconds.

c. ILS BEAM or GP flag behaviour.

d. Faulty flight direction facilities indicated by failure to achieve the selected manoeuvre.

Pre-Flight Checks 41. The pre-flight checks are listed in the Flight Reference Cards. The full functional checks are at the end of this chapter.

42. If electrical ground equipment is in the vicinity of the aircraft, the beam compasses may synchronise on different headings, in which case the warning lights may come on.

43. The twin compass system must not be run unnecessarily while the aircraft is on the ground. Except in flight, or during functional ground tests, the two compass isolation switches at the nav /plotter's position must be at OFF.

Normal Flight Control 44. The beam compass should not be regarded as a steering compass but as an instrument for setting up courses to be steered by obeying the commands of the director horizon.

45. Select the required compass on the MFS selector. The heading index can be set to any position round the beam compass but, for nor­mal en-route operation, it is more convenient to

· set the index to the top datum and rotate the compass card so that the required course is in line with the index. For course changes, rotate

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RESTRIGrED AP lOlB-1902-15

the compass card to bring the new course against the heading index; the ADP on the director horizon indi­cates the necessary bank required to turn the aircraft on to the new heading.

46. En-Route Flying

a. To fly magnetic headings, set the MFS navi­gation selector to the centre position.

lb. To fly true trnck, magnetic track or true heading, set the navigation selector to REMOTE. The navi­gator can then feed in drift (from doppler) andjor magnetic variation to the beam compass heading

, pointers. Any changed selection by the navigator causes the heading pointer to move relative to the compass card and necessitates resetting the compass card if the previous course is to be maintained.

c. To fly HRS courses set the navigation selector central and move the MFS jHRS switch to HRS. The ADP now show demands to fly the HRS course set up by the navigator.

47. Let-Down and Approach. Detailed instructions for handling the MFS during let-down and approach are given in Part 3 Chapter 3. Note: After turns of more than 90° at low altitude, heading errors of up to 5{) and, exceptionally, up to 10° can occur. Elimination of these errors may take up to 8 minutes.

Failures 48. Attitude. If there is a power supply failure, or a difference of 10° of roll or 3to of pitch signals between the port and starboard systems, warning flags show on both director horizons. Check each director horizon against other flight instruments and disengage the autopilot if in use.

49. Heading. If the heading indications on the beam compasses differ by more than 5°, both amber warning lights come on after 20 to 45 seconds. Check both annunciator units and, if both are functioning correctly, ensure that the navigation selector is central to obtain magnetic heading and check each beam compass read­ing with the best magnetic heading available. Select DG on the faulty compass and notify the navigator that this has been done. Select the serviceable compass on the MFS selector.

50. Director Horizon Faults a. A director horizon, although supplied with cor­rect input signals, can develop a fault causing reduced sensitivity of the pitch, glide path andjor azimuth director pointers, causing:

(1) The pitch pointer to indicate a smaller pitch attitude than actually exists.

Military Flight System, Autopilot, Flight Instruments

(2) The glide path or azimuth pointer to demand a smaller correction than is actually required.

b. During flight, cross-refer to the two director horizons and the other flight instruments and, if a fault occurs> refer to the serviceable instrument, ie the director horizon presenting the greater demand or pitch indication.

c. If the fault occurs during an ILS approach, break off the approach and use a ground interpreted aid instead, referring to the serviceable director horizon.

ci As the demand signals to the autopilot are unaffected, an auto-ILS approach can be made suc­cessfully although the defective horizon presents false indications.

51. Pitch Pointer Fluctuation. The pitch pointer is liable to minor fluctuations. These may ·be accepted, provided that the vertical gyros are erect and that the variations do not exceed half the thickness of the pitch pointer bar.

52. Port Vertical Gyro Failure a. The possible indications of malfunction or failure of the port vertical gyro are :

(1) Appearance of the attitude flags. (2) Port director horizon rolling continuously due to gyro toppling. (3) Port director horizon display displaced due to a fixed fault. ( 4) Failure of the TFR.

Any of these faults may be caused by fuse failure, therefore the relevant fuses should always be checked. If a malfunction or failure cannot be corrected, the starboard MFS can still be used in all modes.

b. The port vertical gyro feeds pitch and roll signals to the port director horizon and pitch signals to the TFR.

53. Starboard Vertical Gyro Failure-a. The possible indications of malfunction or failure of the starboard vertical gyro are:

(1) Appearance of the attitude failure flags. (2) Starboard director horizon rolling continu­ously due to the gyro toppling. (3) Starboard director horizon display displaced due to a fixed fault. ( 4) Pitch scale oscillations following operation of the fast setting facility or the selection of HEIGHT or APPROACH.

Any of these faults may be caused by fuse failure, therefore the relevant fuses should always be checked.

b. The starboard vertical gyro feeds pitch and roll signals to the starboard director horizon, pitch signals to the pitch computer (used for director signals to

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both director horizons when fast chase, HEIGHT or APPROACH is selected) and roll signals to the autopilot (used as a roll datum for angles of bank of 5° to 45° except in the auto-ILS mode).

c. If a malfunction or failure cannot be corrected, the port MFS can still be used in normal flight and during instrument or auto-ILS approaches by use of the emergency (manual operation) function of the port director horizon pitch scale knob. A void low level flight except in cases of operational necessity. Do not use TFR (because incorrect pitch indica·tions will be given).

d. Do not use the autopilot in normatl flight; it may be used during an auto-ILS approach. If the autopilot is used with track selected when the MFS navigation selector is set at other than LOC & GP, a continuous roll may occur when a heading change requiring more than 5° of bank is selected; when LOC & G P is selected, the autopilot roll monitor function is inhibited.

e. To caay out an auto-ILS approach, position the aircraft manually on the approach about 5 miles before the start of the glide path. At the correct height and with the speed below 180 knots, set the QDM of the ILS on the port beam compass, engage the autopilot in the normal manner, select LOC & GP, pull TRACK and then:

(1) When the glide path pointer begins to move down the pitch scale, adjust the pitch scale so :that the centre dot is in line with the horizon bar. The pitch scale has no fixed relationship to attitude and must only be used in the emergency mode to help in judging relative movement between pitch pointer and horizon bar.

(2) When the glide path pointer is coincident with the centre dot, pull GLIDE, select airbrakes and power as normal and keep the autopilot trim indicator central.

Throughout, the flight instruments must be monitored carefully and speed, power and attitude must be checked regularly.

f. If an overshoot is necessary, make it manually, remembering that attitude must be interpreted from the horizon bar, pitch pointer, ASI and VSI.

AUTOPILOT General

54. The autopilot provides all the normal facilities. Three rate gyros, mounted at right angles to each other, detect disturbances to the flight path and pass signals, through amplifiers, to servo motors in the control runs.

55. Pitch and sideslip monitors in the autopilot plat­form and a heading monitor in the beam compasses

back up the rate gyros to detect slow rates of move­ment. The autopilot platform, with the rate gyros, is below the 1st pilot's seat. A crossfeed system enables yaw signals to be fed to both aileron and rudder circuits; corrections to heading are thus made with both these controls. An aileron/rudder crossfeed com­pensates for yaw due to aileron drag.

56. As safety measures, a spring strut in the elevator circuit and a roll error cut-out in the aileron circuit automatically disengage the whole autopilot if too great a demand is applied to the controls. A neutral detector ensures that, if a servo-motor is overloaded, power is cut off and cannot be re-applied until the fault is rectified.

57. The aileron artificial feel control stops are sig­nalled to the high speed position when the autopilot is engaged and rudder feel is increased by 45 %, to limit application. To allow sufficient control authority on the ILS approach, the elevator artificial feel is relieved by 20 knots and 1° more aileron movement is provided (but see para 64).

58. Autopilot Control Interlocks. The autopilot con­trols operate in the following order of precedence. Each control overrides all others lower in the order:

Azimuth Control Interlocks 1. TRACK 2. BOMB 3. Turn control 4. Heading monitor

Note 1: Track is de-selected whenever the .M.FS navigation selector is moved to or from the central position, or when MFS/HRS selections are made by the pilot or navigator. Note 2: All controls are inoperative if the aileron channel is disengaged.

Pitch Contt·ol Interlocks 1. GLIDE 2. Fast rate pitch control 3. ALT/IAS 4. Slow rnte pitch control

Note 1: GLIDE is only available if LOC & GP is selected on the MFS navigation selector.

Note 2: All controls are inoperative if the elevator channel is disengaged (but GLIDE remains sele~ed).

59. Power Supplies. The autopilot uses 115V AC from the port transformer, and 28V DC from the non-essential busbar.

Controls and Indicators 60. Autopilot Control Panel. The autopilot control panel at the rear of the retractable console carries the following items :

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RESTRICTED AP lOlB-1902-15

a. POWER Switch. When this switch is pulled out, 115V AC and 28V DC are fed to the autopilot. The gyros then accelerate to working speed, the pitch platform erects to horizontal and the roll frame erects to level with the aircraft wings.

b. READY Magnetic Indicator. Shows black/ white 60 + 15 seconds after pulling the POWER switch, indicating that the autopilot is ready for engagement. Shows black when the autopilot is engaged and black/white if, when the autopilot is engaged, any individual channel is disconnected.

c. Channel Switches. When the rudder (R), aileron (A) and elevator (E) channel switches are set to IN, the control surfaces engage with the auto­pilot, provided that the ENGAGE switch is pulled out. If any one switch is put to the out position, its channel disengages ; if all three switches are put to the out position, the whole autopilot disengages.

d. ENGAGE Switch. Pulled out to couple the autopilot to all three control surfaces (provided that the channel switches are IN).

e. IN Magnetic Indicator. Shows black when the autopilot is disengaged, white when it is engaged.

f. IASJALT Lock Switch. When selected and pulled out, maintains the aircraft at the existing air­speed or pressure height, whichever is selected. Overridden by the fast rate of the pitch control, by GLIDE or by setting the elevator channel to out. This switch overrides the slow rate of the pitch con­trol. Signals are supplied by the port pitot-static system.

g. BOMB Switch. Pulled out to feed signals from the NBS or the bomb aimer's turn control to the autopilot, to maintain the aircraft on the desired bombing track. Overrides the turn control. Aircraft bank is limited to 25°.

h. TRACK Switch. When TRACK is pulled out, the bank angle is limited to 25° maximum and the heading monitor is overridden. Heading errors con­trol the aircraft in azimuth. In addition, when LOC is selected (auto-ILS) and coupling takes place, the aircraft responds to the heading error flocaliser error resultant in the same way as the ADP of the director horizon. The turn control is overridden.

WARNING: When TRACK is pulled with the MFS selector set to LOC & GP, elevator feel is relieved by 20 knots, the aileron stops are reset to give 1 o

more travel, low speed gearing is introduced in the control channels and an increased value of yaw-to-roll crossfeed is introduced as the roll datum instead of the starboard MFS vertical gyro; for this reason, LOC & GP must not be selected with TRACK, even with the ILS off, unless the speed is below 180 knots, otherwise

Military Flight System, Autopilot, Flight Instruments

serious over~banking may result, accompanied by nosc­dow:n pitching.

i. GLIDE Path Switch. Pulled out to feed ILS glide path signals to the autopilot (MFS set up for ILS approach). Overrides ALT lock. When GLLDE is pulled, a programmed 3 o pitoh-down occurs ; :t!he

· aircraft is subsequendy controlled in pit.ch by gtlide path signals and lin azimuth by looaliser signals only, ie, heading error signals are ignored. Bank angle is limited to 10°.

'j. AJL PRIME Switch. Inoperative.

61. Turn and Pitch Controls. The turn and pitch control is at the forward end of the retractable console. A rotary switch turns over a scale marked from 0° to 4(}0 on either side of a neutral detent to provide lturn control; a DIVE/CLIMB switch provides pitch ' con­trol. These switches are used to alter the aircraft attitude.

a. Pitch Control. The switch is spring-loaded to the central (neutral) position and is moved forward to produce nose-down pitch change. Switch movement

· .is opposed by two spring rates, so that initial move­ment against a weak spring causes a slow rate change of attitude and further movement against a strong spring causes a fast rate of change. This control,

' and not the elevator trim, must be used to adjust pitch attitude.

b. Turn Control. To turn the aircraft, move the turn control to the desired bank angle. The control remains at the selected position and the aircraft maintains that angle of bank until the control is moved to a new position, when the aircraft follows the new selection. The control overrides any pre­selected heading on the beam compass, except when TRACK is pulled. Bank angles of up to 45° can be achieved with the turn control.

c. Bomb Aimer's Turn Control. To enable the bomb aimer to control the aircraft in azimuth during a visual bombing run, a further control is provided at the bomb aimer's station.

62. Instinctive Cut-out and Reset Switches. An instinctive cut~out switch is provided on each control column. Applying firm pressure to either switch cuts off the 28V supply to the autopilot and the ENGAGE and POWER switches release. The POWER switch cannot be reselected until the spring-loaded autopilot reset switch on the starboard console has been operated.

63. Trim Indicator. The trim indicator on ·t!he piJots' centre panel, is in ·the form of a side view of the aircraft, with a pointer to show any out-of-trim con­dition in pitch, movement of the pointer showing if the aircraft is trimmed nose or tail heavy. The indicator also carries flags to show when the autopilot is ready and engaged, the indications corresponding with those on the control panel.

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Page 11 CAL 10)

64. ArtificUzl Feel a. When the autopilot is engaged, the aileron feel actuator moves to the maximum speed condition and feel failure warnings in the aileron channel are normally inhibited (if the actuator overruns the maximum speed setting, a limit switch reactivates failure warnings); also the rudder feel actuator is moved to a higher speed setting. The artificial feel MI remains black and normal failure warnings are available in the rudder and elevator channels. Although both aileron and rudder channels should be IN at all times:

(1) If the rudder channel is switched out the feel MI changes to white.

(2) If the aileron channel is switched out the feel MI changes to white after about 10 seconds.

(3) If both rudder and aileron channels are switched out simultaneously, or the autopilot is disengaged, the feel MI shows ILS for about 10 seconds before turning to black.

Note: In some aircraft, the 2-position artificial feel MI shows white whenever the autopilot is engaged and reverts to black about 10 seconds after disengagement.

b. When TRACK is selected with the MFS navi­gation selector set to GP (see para 60h) full normal warnings are available in all feel channels and the feel MI shows ILS. If it remains black, the elevator LOC relay has failed to operate and, because the elevator feel has not partially relieved, the spring strut has less authority and is more likely to disengage.

c. If, on selection of TRACK with GP, the aileron feel fails to reset from cruise to approach setting, aileron feel warnings remain inhibited (unless the failure is due to power supply loss) and, after disengaging the autopilot, the feel MI remains at ILS. This should be regarded as an indication that the aileron feel has failed to the maximum speed condition and the appropriate action must be taken.

65. Autopilot Disengagement. The autopilot will be completely disengaged as a result of:

a. Power off or non-essential loads shed. b. Operation of either instinctive cut-out. c. ENGAGE switch selected down. d. All three channel switches out. e. Spring strut collapse. f. Roll error cut-out operation. g. Neutral detector operation.

Autopilot Manoeuvre Envelope

66. The maximum angles of bank which can be applied when using the autopilot are:

Using tum control Using BOMB or TRACK With GUDE selected

The maximum rate of roll is 4·8° fsec

67. The maximum pitch angles are: ClUnb 25° Descent 400 The maximum rates of change of pitch are: Slow rate . . . 0·5<> fsec Fast rate 1·4° /sec

68. The elevator spring strut is designed to cut out the autopilot when a force of about 26 lb is applied to the control column.

Pre-Flight Checks 69. The pre-flight checks are given in the Flight Reference Cards. The full fun<:tional checks are at the end of this chapter.

Use in Flight 70. In addition to the limitations quoted in Part 2, the following restrictions are imposed on the use of the autopilot:

a. The autopilot may be engaged in climbing, descending or level flight, provided that the aircraft is in trim. Engaging the autopilot in a tum is not recommended.

b. In conditions of severe turbulence, neither the AL T nor the lAS lock may be oogaged.

c. The autopilot may normally only be used when all three channels are engaged and serviceable. However, in cases of operational necessity the air­craft may be flown with the elevator channel dis­engaged. In .these circumstances, the aircraft is not to be flown in .the ILS mode.

d. For rake-off and whenever the autopilot is not in use, the instinctive cut-out must be operated.

e. If malfunction of the pitot-static system is suspected, me AL T fiAS locks must not be used

f. If the MFS starboard vertical gyro has failed the autopilot may only be used if TRACK is selected with LOC & GP set on the MFS navigation selector and the aircraft established on the ILS beam.

71. Set up the aircraft as for pre-flight checks but with the control forces trimmed out for the existing flight oonditions. Oheck the trim lindicator frequently when the autopilot is engaged and adjust the trim as necessary to keep the indication central.

72. To turn the aircraft: a. Select the required angle of bank on the turn control (up to 45°); as the heading is approached, adjust bank accordingly.

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b. With TRACK pulled and the MFS naviga­tion selector central or at REMOTE, the aircraft can be turned by se~ting the required course with the heading index of the beam compass. With LOC and TRACK selected, the heading index within the radio coupling markers and the beam compass sense switch in the appropriate position, ILS localiser signals combined with heading error signals control the aircraft directionally. With the radio signals suppressed, or with the heading index outside the coupling range, the aircraft follows the heading index.

73. To climb or descend, move the pitch control backwards or forwards until the desired pitch attitude is reached. This control must always be used to

change the pitch attitude ; the elevator trimmer is then used to keep the trim indicator central.

74. If the ALT lock is to be used, engage it when the aircraft is in an approximately level attitude. If the auto­pilot has only been switched on after reaching cruising altitude, do not engage the AL T lock for at least 5 minutes, to allow the equipment to stabilise.

75. At maximum altitudes, if it is desired to make a cruise climb, trim the aircraft carefully and then engage the autopilot. The aircraft may take several minutes to settle down to a constant mach number. Alterna­tively, a stepped cruise climb can be made, using the speed lock and readjusting the datum speed as the mach number increases.

76. The autopilot must not be used above 0·87M if the mach trimmer is inoperative, or an unstable pitch oscillation may develop. With the mach trimmer working, the autopilot can be used up to 0·90M (see para 85).

77. If the bomb doors are left open for longer than the period normally used in auto-operation, or if re­peated bomb door operations are made, rudder displacement with resulting yaw may occur, because of temperature effects on the control runs passing through the bomb bay.

Auto-ILS Approach 78. To make an auto-ILS approach, set up the MFS for an ,LLS approach; follow the instructions in Part 3, suppressing the radio coupling as required with the sense switch, and pull TRACK. When the aircraft is in a suitable position, set the localiser QDM to the top datum, couple onto the looaliser by rotating the sense switch to the up position and setting the heading index to the localiser QDM. If the aircraft is less than about 15 miles from the airfield, provided ·that the speed is below 180 knots, select LOC & GP on the MFS, to bring in the low speed autopilot control; the feel indicator shows ILS (or white in the case of the 2-position Ml).

Military Flight System, Autopilot, Flight Instruments

79. As the glide path needle descends to the centre dot of the pitch scale, select GLIDE on the autopilot, set the airbrakes and throttles as required. The aircraft first changes pitch attitude by 3° nose-down, after which it is controlled in pitch by glide path signals. Control in azimuth is now solely from localiser signals, the heading errors being ignored. The bank angle is limited to 10°.

80. If a race track pattern is to be followed for subsequent auto-approaches, leave the sense switch pointing upwards but if the beam QDR is to be flown, followed by a procedure tum, set the sense switch horizontally. The heading index should now be used to manoeuvre the aircraft and the beam QDM left at top datum.

81. If the autopilot is inadvertently disengaged after GLIDE has been sele<:ted, re-engagement of the auto­pilot and reselection of GLIDE causes a further 3° nose-down pitch. It is recommended that the auto­pilot is not re-engaged.

Use of the Autopilot after Engine Failure 82. Following engine failure, carry out the following drill:

a. Physically pre-load the rudder pedals and switch off channel R. Retrim the rudder as required to keep the aircraft straight and then re-engage channel R. b. This drill must be repeated after any further change of asymmetric power.

83. Alternatively, disengage the whole autopilot, retrim the aircraft and re-engage the autopilot.

Malfunctions 84. Unselected Engagement. If the autopilot engages without being selected, or if it fails to dis­engage, press the instinctive cut-out. The autopilot must not be used again.

85. Disconnects and Runaways a. If operation of the spring strut disengages the autopilot, there is nose-up or nose-down change of trim which can be easily overcome. b. If the autopilot cut-out occurred below 0·87M with the automach trimmer servos retracted and it is thought to be because the aircraft is out of trim, the autopilot may be re-engaged after retrimming the aircraft and resetting !the automach trimmer by pressing the COMPARATOR RESET button. If the aircraft was not out of trim, the autopilot must have malfunctioned and must be switched off for the rest of the flight. c. If the autopilot cut-out occurred above 0·87M wid1 automach trimmer servo(s) extended, the pilot must decide whether the automach trimmer or the autopilot is at fault. Reduce speed to 0·85M, switch off the automach trimmer and retract the servos;

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if the extensions are found to be significandy different, a runaway servo should be suspected (see Chapter 7 para 63). Switch on the automach trimmer, press the COMPARATOR RESET button and, if the servos remain retracted, accelerate to 0·9M, checking that the blue extension lights come on approximately together. If the automach trimmer appears to be unserviceable continue as in Chapter 7 para 63. Autopilar flight may be continued with the automach trimmer S1.Vitched off, provided that the speed is kept below 0·87M. If the automach trimmer is serviceable, continue as in sub-para b.

d. If the aircraft rolls unexpectedly and the auto­pilot disengages, the roll error cut-out must have operated; switch off the autopilot.

86. Yaw Damper Failure. If the aircraft yaws unexpectedly, the fault could be caused by yaw damper or autopilot malfunction. If the autopilot continues to work normally it is likely that the yaw damper is at fault. With the aircraft in straight and level flight, and with the feet on the rudder pedals prepared for a trim change, switch out the autopilot; if there is a sudden change of trim, it is likely that the yaw damper has failed. Select the other yaw damper, retrim the rudder and re-engage the autopilar.

87. Indications. If the autopilot fails or disengages, the IN indicator shows black and the READY indicators show black/white. If a power failure occurs, both flags on the trim indicator show black.

OTHER FLIGHT INSTRUMENTS

Pitot-Static System

88. Pressure Heads. Two pressure heads, below and on either side of the nose, are heated using 115V AC from one phase of the 200V supply (No 2 busbar port, No 4 busbar starboard) via two ganged switches on the inboard edge of the starboard console which also con­trol the heating supply for the OAT probe and the supply to the time delay for the hydraulic power pack. When TFR is installed, they also control ADD probe heating. A magnetic indicator at the top of the centre panel shows white when the port heater is switched off, or if the fuse has failed. The port and starboard systems are independent.

89. Static Vents. On either side of the nose is a static vent plate with three inlets; only two inlets on each side are in use. From each vent, a line crosses the nose

to connect with the vent on the other side. The two pipes supply the port and starboard pitot-static systems.

90. Covers. Covers are provided for the pressure heads and plugs for the static vents.

91. System Supplies. The services supplied by the two systems are :

Port System

1st pilot's panel

Autopilot

Monitor panel,

autostabilisers

Mach trim panel

Artificial feel failure warning

Starboard System

Co-pilot's panel

MFS manometric unit

NBS calculator and computer

Monitor panel, autostabilisers

Navigator's panel

Mach trim panel

True airspeed unit

Fatigue meter airspeed switch

Artificial feel units

TFR transducer Airspeed warning transducer

92. High Airspeed Audio Warning System. An air­speed transducer, operating from the pitot-static system, produces through an audio generator a warning note on ·the intercom, at an airspeed of 375 to 380 knots. A test button is outboard of the port console. The system uses 28V DC. An isolation switch is on the 1st pilot's console.

Pressure-Operated Instruments

93. Each pilot has an ASi at the top left side of his flight instrument panel.

a. The 1st pilot's instrument is a Mk 15A.

b. The co-pilot's instrument is a type KAB 0506K and incorporates a warning flag which pulsates if the speed is below 160 knots and the undercarriage is not locked down.

c. There is a further ASI (type KAB 0505K) at the nav /plotter's station.

d Correction cards for each instrument are dis­played in the cockpit.

94. There is a Mk 3A machmeter at the top of each pilot's instrument panel.

95. Each pilot has a VSI on the right-hand side of his instrument panel. These may be either Mk 3P or Mk 3Q instruments.

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1st PILOTS PANEL

FATIGUE METER

AP 101B-1902-15 Military Flight System, Auto- pilot, Flight Instruments

...._-1-~-- BOMB FUZE AIRSPEED SWITCH

GROUND TEST

PJTOT PRESSURE STATIC PRESSURE

AIRSPEED SWITCH---~--+-"

!T--1H -------,..ARTIFICIAL FEEL UNITS

1-12 Fig. 4 Pitot-static System

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96. a. Each pilot has an altimeter on the left of his instrument panel; a further altimeter is at the nav J plotter's position.

b. The 1st pilot's Mk 29B altimeter is a servo­operated instrument with automatic reversion to cap­sule operation if electrical supplies fail or if selected by a SjR knob on the right-hand bottom corner of the instrument (S refers to the standby or capsule­operated mode, and R for reset or servo mode). In the standby mode, an aperture above the veeder counter shows STBY, and a vibrator operates on 28V DC ESSENTIAL. The range of the instrument is from minus 1000 feet to 60,500 feet. Indications are by a pointer and drum showing hundreds of feet, and two counters showing thousands and tens of thousands of feet; the left-hand counter being marked with diagonal black and white stripes below 10,000 feet, and diagonal red and black stripes below zero feet.

c. The co-pilot's Mk 30A altimeter, operating on 115V single phase AC from the No 4 load busbar, has a height encoding facility used with channel C of the IFF jSSR and a pressure error correction unit; an aperture above the veeder counter shows PE if this unit fails. The range of this instrument is from minus 900 feet to 60,500 feet. The indications are similar to the Mk 29B except that the left-hand counter shows red and white stripes below zero feet.

d. The navigator's Mk 19F altimeter bas a range of minus 1000 feet to 60,000 feet. The display is of the 2-needle type showing hundreds and thousands of feet, with a white disc moving around in the centre of the instrument to show tens of thousands of feet.

e. It is essential that frequent comparisons are made between the readings of the 1st pilot's, co­pilot's and navigator's ASI and altimeters and between the pilot's macbmeters, especially at high indicated mach numbers, high or low indicated air­speeds and low altitudes.

97. The true airspeed unit in the starboard wing provides T AS to the NBC and drives an AMI at the navigator /plotter's panel.

98. Fatigue Meters a. Two fatigue meters are fitted in tbe bomb bay. One h switched on automatically by an airspeed pressure switch operating at speeds above the range 135 to 170 knots; the other, a low level fatigue meter, is controlled by a switch on the navigator's panel. The meters use 28V DC non-essential supply.

b. A fatigue consumption indicator on the panel between the navigator I plotter and AEO is used with a transmitting fatigue meter Mk 18A in the bomb bay.

Military Flight System, Autopilot, Flight Instruments

Standby Instruments

99. The 1st pilot has a standby horizon gyro unit Mk SA on the right of his flight inruument panel. With power applied the gyro erects within 90 seconds, unless the gyro is more than 10° from the erect datum, when the fast erection button must be used. The fast erection button need not be held in for use but should be checked out and free to turn, prior to flight. A striped flag shows when no power is available or when the instrument is malfunctioning. 115V 3-phase power is used, from a separate 40 VA transformer connected to No 4 busbar.

100. Each pilot has an E2B standby compass, on either side of and above the central windscreen.

Miscellaneous Instruments

101. HRS Mk2 a. A beading reference system (HRS) consists of a master reference gyro (MRG), a heading repeater unit (HRU), a control unit, a navigator's compass repeater type B and a navigator's beading unit (NHU). Heading error signals can be fed to the MFS via an HRSjMFS selector on the starboard side of the fuel contents panel, provided that the MFS navigation selector is central; the system can also be used with NBS. The system uses 115V AC from the NBS transformer and 28V DC from the non-essential busbar.

b. The NHU combines aircraft heading with drift to give a track output for comparison with the desired track set on the unit dial. The NHU output is fed to the autopilot and pilots' ADP in the form of fly left/right demands.

102. Accelerometer. A Mk 2A accelerometer is on the left of the pilots' centre instrument panel.

103. Slip Indicators. Each pilot has a bubble-type slip indicator at the bottom of his instrument panel.

104. Air Temperature Gauge. A ratiometer-type air temperature gauge is at the nav /plotter's position. The probe is on the starboard side of the fuselage, just aft of the radome. A 28V DC de-icing heater is controlled by the port pressure-head beater switch.

105. Clock. There is provision for an eight day clock at the bottom of .the 1st pilot's instrument panel.

106. TailjGround Warning. A hinged arm is fitted underneath the aircraft tail cone and two red lights are installed on the coaming in front of the 1st pilot. If the arm touches the ground, during a landing run for example, the lights come on, warning the pilot that the tail is too close to the ground.

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FULL FUNCTIONAL CHECKS

Military Flight System

107. After 100 hours flying, or before flight ·testing, make the checks in the following paragraphs.

108. With power supplies available but not on, check on the MFS that the attitude warning Bag is in view on each director horizon ; switch on powu and check that the pitch pointer and horizon bar on each director horizon oscillate violently, settling down as the gyro­scopes erect.

109. After about 30 seconds, check that the pitch and roll attitudes indicated on the two horizons are the same and that both attitude warning flags have cleared. The actual pitch and roll attitudes indicated depend on the aircraft ground attitude.

110. Set the COMP knob fully anti-clockwise (arrow pointing to port) and the navigation and pitch selectors central (vertical). Check on both pilots' director hori­zons that the setting knobs are not in the fully-out position.

111. Switch ON the 1st pilot's and co-pilot's compass isolation switches and set both annunciator unit switches to COMP after one minute. Synchronise both compass systems to centralise the indications of both port and starboard annunciator arrows.

112. Push and tum the setting knob of each beam compass to set each compass card with the aircraft heading at the top compass datum. Check that both heading pointers indicate approximately the same air­craft heading as the standby magnetic compass. Some disparity ·in the readings can be expected if the aircraft is in a hangar or if magnetic material is in the area of the wingtips.

113. Set the port annunciator switch to DG and check that the DG flag appears on the port beam compass. Repeat on the starboard system.

114. Turn the synchronising knob on the port annunciator unit to set the port heading pointer at least zoo clockwise from the starboard heading pointer.

115. Set both annunciator switches to COMP. Check that both DG flags have cleared, and that both compass warning lights come on within 45 seconds.

116. Slowly rotate the port synchronising knob to bring the port heading pointer gradually towards the same indication as the starboard pointer. Check that the warning lights remain on while the readings differ by at least 7° but go out •before the difference is

reduced to 3° . Repeat the check after using the port synchronising knob to set the port heading pointer at least zoo anti-clockwise from the starboard pointer.

117. Turn the port synchronising knob to set the port heading pointer to a reading sufficiently different from the starboard pointer to bring the warning lights on. Set the port annunciator switch to DG and check that the warning lights go out. Restore the switch to COMP and check that the lights come on again. Repeat on the starboard system.

118. Set the beading pointers on both beam com­passes simultaneously in clockwise rotation by turning both annunciator unit synchronising knobs clockwise. Check that the warning lights go out while both pointers are rotating and come on again when the pointers stop. Repeat this check in the reverse direc­tion by simultaneously turning both synchronising knobs anti-dockwise.

119. Resynchronisc both compasses to centralise the indications of both annunciator arrows.

120. Pull out and tum the port beam compass setting knob to set the heading index to the top compass datum. Push in and tum the setting knob to set the heading pointer coincident with the heading index. Check that the ADP of both director horizons are at zero.

121. Pull out and tum the port beam compass setting knob to rotate the heading index up to 20° either side of the top compass datum and check that both azimuth director pointers move freely over their full range of movement. Check that the AD P move to the right when the heading index is set to the right and to the left when set to the left.

122. Set the COMP knob (MFS selector) fully clock­wise (arrow pointing to starboard) and repeat the checks in para 120 and 121 on the starboard system.

123. Check that no Bags are visible on either director horizon and check that the radio displacement bars of both beam compasses are central on their scales.

124. Check the operation of the pitch scale on the port director horizon by means of the setting knob:

a. Pull out the knob fully and check that, by rotat­ing it, the pitch scale can be manually adjusted both up and down ; leave approximately central.

b. Push the knob and release, allowing it to return to its centre position. Rotate the knob clockwise against spring pressure and note that the pitch scales on both director horizons move slowly to the top of their travel and that the pitch flags appear. Release the knob.

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· c·. Push the knob fully home and note that the pitch scale centre dots on both director horizons move rapidly to a position directly under the pitch pointers and that the pitch flags appear. Release the knob.

d. Rotate the knob anti-clockwise against spring pressure and note that the pitch scales on both director horizons move slowly to the bottom of their travel and that the pitch flags ~ppear. Release the knob.

e. Push in · the knob fully and note that the pitch scale centre dots on both director horizons move rapidly to a position directly under the pitch pointers and that the pitch flags appear. Release the knob.

f. Repeat on the starboard director horizon. Leave the pitch scale centre dots off-set from the pitch pointer.

125. Select the MFS navigation selector to LOC and check that the BEAM flags on both director horizons pulse fully in and out of view. Select the LOC & GP position and check that both BEAM and GP flags pulse.

126. Select HEIGHT on the MFS pitch selector and note that the pitch flags appear on both horizons; the pitch scale should move slowly towards the pitch pointer. Pull out the setting knobs on each horizon in turn and note that the flags on the respective instru­ments disappear. Return the pitch selector to the centre position and return both horizon setting knobs to the centre position.

127. Rotate the pitch scale setting knob clockwise on the port director horizon until both pitch scales are at the top of their travel ; release the knob.

128. Select and hold on the DATUM position on the pitch selector and check that both pitch scales move rapidly to the centre of the instrument and that the pitch flags appear.

129. Release the pitch selector and note that it locks in the APPROACH position. The pitch scale should move slowly towards the pitch pointer. If the heading pointer is offset from the heading index, the ADP should move towards zero bank (drift unit function).

130. Select the navigation selector to its centre posi­tion and check that the pitch selector releases to its centre position and that the pitch flags disappear.

131. If the aircraft is not to be flown, switch off b<)th compass switches and the power supply.

Autopilot

132. After 100 hours flying, or after autopilot mal­function, make the checks in the following paragraphs.

Military Flight System, Autopilot, Flight Instruments

133. Engagement. With power supplies on, pull out the POWER switch and check that the READY indicators show black/white after 60 ± 15 seconds. Set the R, A and E switches IN. Check ·that both beam compasses are synchronised and that the flying controls are centralised, ·then pull out the ENGAGE switch. Check that the READY indicator shows black, that the IN indicator turns white, that the IN flag shows on the remote trim indicator and that the artificial feel magnetic indicator remains black.

134. Individual Channels a. Check the operation of each channel switch in tum. With any channel switched off, the READY indicators show black/white and the appropriate control is free. When an individual switch is returned to IN, check that the control re-engages. Switch out all three channels and check autopilot and control disengagement. Reselect all three channels IN and check that the autopilot does not re-engage until the ENGAGE switch is pulled out.

b. When the aileron channel is selected out, the feel magnetic indicator turns white after 10 seconds and turns black when IN is selected. When the rudder is selected out, the indicator turns white immediately. If the aileron and rudder channels are .bo'th disengaged, the feel indicator shows ILS for 10 seconds and then black.

135. Disengagement. Operate the instii.nctive cut-out on !the p<)rt contJrol ooluinn checking ilhat t!h.e contrOls become free, 'lihe POWER butJtOn drops in, the auto­pilot indications revett to power off and the artificial feel indicator shows ILS for 10 seconds and then shows black. Operate the autopilot reset switch on the star­board console, pull up the POWER button and check that operation of the starboard instinctive cut-out causes the POWER button to drop in.

136. Trim Indicator and Spring Strut. Before start­ing this check, select feel relief and cancel the main warning ; check that the feel indicator is white. Engage the autopilot and push the elevator control forward, checking that the trim indicator shows nose-heavy. Increase the push-force to approximately 25 lb, when the spring strut should operate and disengage the auto­pilot fully. On disengagement, the mach trim amber light and the main warning lights come on and the autostabiliser magnetic indicator goes white. Press the comparator reset button and check that mach trim indications return to normal. Repeat the checks in the opposite sense.

137. Roll Error Cut-out and Aileron Control Stops. Re-engage the autopilot and, after 10 seconds, move the turn control quickly to full travel. The ailerons should move to about t travel and, after a slight delay, the autopilot should disengage fully. Repeat the check in the opposite direction.

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138. Compass Monitoring. Engage the aileron channel only. If the aileron control is moving slowly in either direction, rotate the heading pointer with the syn­chronising knob, in the correct sense to stop the move­ment. Disengage the autopilot, resynchronise the compass and re-engage the autopilot. The aileron control should now be stationary ; rotating the heading pointer slowly in either direction (using the synchronis­ing knob) starts the aileron control moving in the appropriate sense. Resynchronise the compass and repeat the check with the other pilot's compass selected.

139. Tum Control. Move the tum control slowly from the detent, making sure that the aileron control moves in the correct sense. Set the heading index and heading pointer to the top datum and pull TRACK. Check that the turn control is inoperative and leave it fully deflected. Push in TRACK and check that the turn control is inoperative until it has been returned to its detent.

140. Pitch Control. Engage the autopilot and check that movement of the pitch control moves the elevator control in the correct sense. Select and pull AL T and check that the pitch control is inoperative on slow rate but that on selection of fast rate the AL T switch drops in and the slow rate is once more operative. Repeat the check with lAS selected.

141. Heading Error Steering. With the MFS navi­gation selector central and the heading pointer over the heading index: at the top datum, pull TRACK. Check that rotating the heading pointer either way moves the control in the correct sense. Check that

GLIDE cannot be selected. Select LOC on the MFS navigation selector and check that the TRACK switch drops in. Reselect TRACK and recheck the heading index operation and that GLIDE cannot be selected. Select LOC & GP on the MFS; the TRACK switch should stay up. Recheck the heading index operation. Select GLIDE and check that the ailerons no longer respond to movements of the heading index. Select the navigation selector to central and the TRACK and GLIDE switches should drop in.

142. Elevator Partial Feel Relief. With the MFS heading pointer and index at the top datum, switch off the aileron and rudder channels and reset their artificial feel systems ; check that the feel indicator shows black. Select LOC & GP and pull TRACK ; the feel indicator should show ILS.

143. GLIDE Interlocks. Select ALT and then GLIDE. The· AL T switch should drop in, the eleva­tor control should move forward and the spring strut may operate, disengaging the autopilot. Repeat the check with lAS selected. With GLIDE selected, check that the pitch control is inoperative.

144. REMOTE Interlock. With the MFS naviga­tion selector central and TRACK pulled, select RE­MOTE on the MFS selector; the TRACK switch should drop in. Reselect TRACK and, by moving the heading pointer with the track control unit, check that the ailerons move in the correct sense.

145. Shut-down. Disengage the autopilot, switch out the channel switches and push in the POWER switch.

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• RESTRicrED AP lOIB-1902-lS

PART I

CHAPTER 13-PNEUMATIC SYSTEMS

Contents

General

General

ENTRANCE DOOR SYSTEM Entrance Door System Supplies Emergency Door Opening Canopy Jettison ... Door Closing Door and Canopy Seals De-Icing Tank Pressurisation

OTHER PNEUMA TIC SYSTEMS Scanner Pressurisation and NBS Engine Air Pressure Nitrogen Pressurisation ...

IDustrations Door, Canopy Seals NBS and H2S Supplies ... Engine Air Supplies Nitrogen Pressurisation

1. There are five, separate, pneumatic storage systems in the aircraft as follows :

a. Emergency door opening and canopy jettison.

b. Entrance door closing, door and canopy seal inflation and bomb aimer's window de-icing.

c. H2S scanner and NBS.

d. Undercarriage emergency lowering (Chapter 10).

e. Engine starting (Chapter 5).

2. The engines supply compressed air for the follow­ing systems:

a. Bomb bay seal inflation.

b. Hydraulic reservoir and power pack pressurisa­tion (pre-mod 2321).

c. Fuel tank and recuperators pressurisation.

d. Equipment in rear fuselage.

e. Anti-icing.

ENTRANCE DOOR SYSTEM

Entrance Door System Supplies 3. Three storage cylinders, charged to 2000 PSI from an external supply, are on the port side of the crew compartment. Their charging points and pressure gauges are on the front bulkhead in the nosewheel bay.

Para I

3 5 9

10 12 13

14 17 21

Fig 1 2 3 4

The forward cylinder supplies pressure for emergency ~door opening and canopy jettison; the remaining two

bottles supply door closing, door and canopy seal in-~ flation and pressurisation of the bomb aimer's window de-icing tank.

4. A ground servicing cock, on the underside of the crew's floor, is normally locked in the open position by a red cover. When turned off, it isolates the services supplied by the rear cylinder.

Emergency Door Opening 5. Pressure at 1200 PSI for operating the door jacks is controlled by the EMERGENCY position of the door opening lever, on the forward end of the door frame on the port side. With the undercarriage down, the nitrogen passes through a restrictor, to control the rate of movement.

6. The rear crew switches (see Chapter 9 para 2) allow nitrogen to pass to a jack. The jack rotates a Jayshaft which withdraws the bolts and moves the lower portion of the door handle to the emergency position. This operates the emergency door opening valve, allowing nitrogen to pass to the jacks. The cabin should be depressurised first. Loads on the door bolts are such that, using the switches alone, the door cannot open until the differential pressure has dropped to 1 · 5 PSI (30 seconds at 43,000 feet, 9 · 5 seconds at 27,000 feet); if the manual control is used at the same time, the

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door opens at 2 · 65 PSI (20 seconds at 43,000 feet, 5 · 5 seconds at 27,000 feet). However, unless escape in the minimum time is essential, the manual control should not be used simultaneously with either of the switches.

7. When the door has been opened by either switch, the door opening lever is carried to the gate while mechanism below the lever moves further to operate the emergency door-opening valve. A 28 volt fault could result in the door closing again under slip­stream pressure. T o protect against this, the first man at the door must ensure that the lever is gated in the emergency position. The emergency door opening switch is on the vital busbar.

Note: A worn gate may allow the lever to pass to the emergency position.

8. If the door locking pins become scored, the door may fail to open when the navigator's switch is used. In this case, use the manual control ; closing the cabin air switches may assist by reducing the time required to depressurise the cabin.

Canopy Jettison

9. The canopy jettison pneumatic valve is operated by pulling any one of the pilots' ejection seat firing handles. When the valve is operated, nitrogen at 1200 PSI passes to a ram which opens the jaws of the canopy attachment and operates the jettison gun. For mechanical jettison, see Chapter 9.

Door Closing

10. The door closing valve is operated either by a toggle handle or, externally, by a pushbutton near the door handle. Either selection must be held until the door is locked closed, otherwise nitrogen pressure is lost from the jacks. The toggle handle is stowed under a hinged panel on the starboard side of and below the centre crew seat.

11. When the valve is operated, nitrogen at 400 PSI is fed to the up side of the door jacks.

Door and Canopy Seals

12. The door and canopy seals are supplied with nitrogen from the rear cylinder at 25 PSI. The door seal is inflated automatically when the entrance door is closed. The canopy seal is permanently inflated to 25 PSI by a mushroom valve which is held open when the canopy is locked. The seal deflates when the canopy starts to move during jettison.

De-Icing Tank Pressurisation

13. The de-icing tank is pressurised with nitrogen at 14 to 15 PSI. A pressure-maintaining valve in the fine cuts off the supply when the pressure in the main system falls to 150 PSI. The supply is also fed to the

probe de-icing tank but this system is inoperative. The ground servicing cock must be closed before filling the tank and the residual pressure of 14 to 15 PSI must be relieved by pressing the pressure release valve on top of the tank.

OTHER PNEUMATIC SYSTEMS

Scanner Pressurisation and NBS

14. The storage cylinders, charged to 1800 PSI, are on the starboard side of the nose section. The common charging point and individual pressure gauges are behind an access panel on the port side of the nose.

15. H2S System. Nitrogen at 30 PSI is passed to a regulator, the supply being controlled by an ON/OFF switch at the navigator's station. Also at this station is an absolute pressure gauge, showing the pressure delivered from the regulator to the scanner.

16. NBS System. Nitrogen at 30 PSI is delivered to the NBS differential pressure regulator, the supply being controlled by an ON/OFF switch at the navi­gator's station. From the regulator, it passes to a manifold, to supply the various components of the NBS. A pressure gauge at the navigator's station shows the pressure at the manifold.

Engine Air Pressure

17. The air supplied by the engine compressors can be divided into two distinct systems, ie engine and airframe supplies.

a. Engine Supplies. Air, tapped from various points on the engine, is used to supply the following engine systems :

System Source of air

(1) CSDU oil tank pres-surisation Intermediate casing

(2) Bulkhead seals

(3) Bearing pressurisation

( 4) Turbine cooling rear face

(5) Other turbine sur-

Intermediate casing

Intermediate casing

3rd stage HP com­pressor

faces 3rd stage HP com-

(6) Induced cooling

(7) Engine anti-icing

pressor

HP compressor casing

HP compressor casing

b. Airframe Air Supplies. All the airframe ser­vices use air tapped from the HP compressor delivery casing. Most have switched controls but some are

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•

CA~ ~ SCHRADER

VALVE SEAL VALVE

SEAT Cfi ARING --..... HANDLE :

CANOPY : CONT~OL VALVE 1 -an- I

CANOPY 1

JETTISON 1

OPERATING RAM :

SEAT~ I FIRING _ _ .J

HANDLE

CAIJOPY SEAL

- CYLINDER P!tESSURE 2DDOPSI

- ~~S~~"JE\rY~ \'1!J"M"D DOOR CLOSING 400 PSI

DOOR AND CANOPY SEALS 25 PSI

DE-IClH G TANk PRESSURISATION 14- 15 PSI

CHARGING POINT

PRESSURE RELIEF VALVE 2000

PSI

GAUGE IN NOSE

GAUGE IN NOSE

RESTRICTED

CAM-OPERATED DOOR SEAL

400/25 PSI

gJ

AP 101 B-1902-15 Pneumatic Systems

DEFLATION VALVE PRESSURE ~ MAINTAINING VALVE

TO &OMS AIMERS WINDOW 150 PSI DE-ICING TAN~~ ~

CA&LE C! 400/14 ·15 CONTROL DOOR CLOSING PSI

PUSH VALVE BUTTON!}-

NITROGEN BOTTLE

Jl GROUND J SERVICINGW

COCK DOOR SEAL I

DOOR OPENING VALVE

-E3I I DOOR

OPERATING JACK

DOOR OPERATING

JACK

-E3'I I I

RESTRICTOR

-an--OPERATING ON DOOR

LAY SHAFT

CONNECTED TO __ R~S!_Rf.£T_2R_ -+ NOSEWHEEL UNIT

BY· PASS SHOWN WITH NOSEWHEEL DOWN

1-13 Fig 1 Door, Canopy Seals ~ T4 bombsight supply deleted •

TEST POINT

~ PRESSURE

FORWARD REDUCING NITROGEN VALVE BOTTLE

EJ NON-RETURN VALVE

(]) PRESSURE GAUGE

:::::fl TEST POINT

)CX:::X:D:::::::X:~ Dl FFERENTIAL R PRESSURE REGULATOR

PRESSURE REDUCING

VALVE 1800/30 PSI

ELECTRICALLY OPERATED

COCK

CALCULATOR POWER UNIT

PRESSURE REDUCING

VALVE 1600/30 PSI

ELECTRICALLY OPERATED

COCK

TEST POINT

VENT TO CABIN

PRESSURE

REGULATOR

1-13 Fig 2 NBS and H2S Supplies

R ESTRICTED

AIR FROM NOSEWHEEL BAY

AUTO POWER UNIT CALCULATOR

- CHARGING PRESSURE 1800 PSI

m:: N 8 S SYSTEM PRESSURE 30 PSI

zzzzzz. H 2 S SYSTEM PRESSURE 30 PS I

EB NON-RETURN VALVE

GAUGE IN CABIN

SCANNER CONNECTION

1-13 Page J. (AL6)

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FUE; TANKS PRESSIJRISATION

1

~ ............................ ~P--- 10Ms&AYFUEL ~~~~~ ~':!KS PRESSURISATION t t

t

N"l ENGINE

N"2 ENGINE

/ /--- -PRESSURE REDOCING} • // VAlVE 200!'15 PSI 1

RECUPERATORS// ~ / /

/

/ /

/

(RESTRICTOR

I I I I I I

: RESERVOIR POST-MOD 2321 I

L --- ---- - -- _s~:_ ~~~ ~ -- j

I I

I I

2

CONTROL VAlVE

PRESSURE REDUCING VALVE 200(10 PSI

1- 13 Fig 3 Engine Air Supplies Mod 2321 embodied

CHARGING VALVE

200/20PSI

t

• + RECUPERATORS

- PlAIN rii!SSUJ! . ZOO PSI

..:::. Ws1~v~~:~~rmAULIC ZZZZZ IOH& DOOft SEAl ri\ESSllltE .

10 PSI ~ RAriO EXHAUST VAlVE

1:::£1 MOM·RETUftM VALVE

GAUGE WAVEGUIDES

NITROGEN BOTTLE

CHARGING VALVE

NITROGEN BOTTLES

1- 13 Page 4

PRESSURE REDUCING VALVE

NON RETURN VALVE

GAUGE

TEST POINT

AR I 18146

SHUT-OFF VALVE

PRESSURE RELIEF VALVE 40 PSI A

1- 13 Fig 4 Nitrogen Pressurisation Systems Mod 2193 and 2321

RESTRICTED

AERIALS COUNTERPOISE

PLATE

WATER TRAP

PRESSURE RELIEF VALVE 22 PSI

AERIAL TUNING UNIT

• RESTRICTED AP lOlB-1902-15 Pneumatic Systems

automatic on engine start. The airframe services are:

Service Engine air

(1) Ground starting other engines

(2) Air ventilated suits .. .

(3) Cabin conditioning .. .

( 4) Airframe anti-icing .. .

(5) Bomb bay heating

~~

(6) Main fuel tank pres­surisation

(7) Fuel recuperators

(8) AAPP fuel tank

(9) Bomb door seals

(10) ECM pressurisa;ion

(11) Hydraulic tank pres­surisation (main and EHPP, pre-mod 2321)

All or as selected

All

All (or selected pairs)

Engines in associ-ated wing

Inoperative ~

Each group from associated engine

Associated engine

No4 Nos 1, 2 and 3

Nos 1, 2 and 3

Nos I, 2 and 3

~ 18. a. Before air is supplied to the airframe services (1) to (4), engine air switches and the individual service control must be ON. Although fuel tank~ pressurisation does not depend on the engine air switches being open, the tank pressurisation switches must be on.

b. Pre-mod 2321, services (7) rto (11) are auto­matica'lly supplied with air on engine start. Post-mod 2321, services (7) to (10) are automatically supplied with air on engine start.

19. The CSDU oil tank, bulkhead seals, all bearing seals and the turbine cooling system are supplied with air automatically on engine starr. Induced cooling is accomp'lished by air from the HP compressor casing,

passing through an electrically-operated induced cooling valve. The valve is controlled by a microswitch on the main undercarriage leg but will not operate if the undercarriage is lowered by the emergency system.

20. Engine anti-icing is controlled by one of two switches on the co-pilot's anti-icing panel. The single­pole switches, labelled PORT and STARBOARD, are

~used for engine anti-icing on the ground. Each switch controls its associated pair of engines. The other method~ of switching on anti-icing for the engines is to operate the airframe anti-icing switches to AUTO or MANUAL. Since these switches also operate airframe anti-icing, they are not to be used on the ground,

~ except when required for the take-off and landing run. ~

Nitrogen Pressurisation

21. a. Post-mod 2193, two separate nitrogen systems are provided in the rear fuselage to pressurise radio equipment and, post-mod 2321, the hydraulic system.

b. Two storage cylinders, in the roof structure aft of the ECM equipment, are charged with nitrogen to 3000 PSI, through an external charging point. From the starboard panel, pressure is reduced to 30 PSI for ARI 18076 and ARI 5874; from the port panel, pressure is reduced to 23 PSI for ARI 5952. A light on the AEO's panel comes on when the ARI 5952 pressure is above 18 PSI. A shut-off valve is provided on the port side of the tail cone.

c. Post-mod 2321, this system . is also used to pressurise the main hydraulic reservoir and the EHPP at 15 PSI. See Fig 4. A shut-off valve on the port side of the tail cone enables hydraulic pressuri­sation to be iso1ated.

d. A single storage cylinder, in the rudder PFC compartment is charged with nitrogen to 1800 PSI, through an external charging point, and supplies a pressure of 20 PSI to ARI 18146.

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PageS (AL 9)

• RESTRICTED AP lOlB-1902-1 5

PART 1

CHAPTER 14-RADIO AND RADAR

Contents COMMUNICATIONS RADIO

UHF/VHF PTR175 (RT1) UHF I ARC52 (RT2) Tone Release HF Intercom Station Boxes

NAVIGATIONAL RADIO AND RADAR Rllldio Compass (AD F) ILS Tacan Doppler and GPI Radio Altimeter Mk 7B R<adar Altimeter Mk 6A

OPERATIONAL RADIO AND RADAR NBC and H2S ECM IFF

COMMUNICATIONS RADIO

UHF /VHF PTR175 (RTl)

1. The PTR 175 installation provides 3500 channels between 225 and 399 ·95 MHz (UHF) and 369 channels between 117 · 5 and 135 · 95 MHz (VHF) or 19 preset channels, one of which is tuned to 243 MHz. In addition, a separate receiver allows a preset guard frequency between 238 and 248 MHz to be super­imposed on any selected channel. MCW transmission is available.

WARNING: UHF transmissions must not be made in the ILS glide path frequency band of 328 · 6 to 335 · 4 MHz, as the ILS display of the parent or any other aircraft within 5 miles may be affected during an ILS approach.

2. The control unit at the AEO position carries the following controls :

a. A 20--position rotary switch giving 'selection of 18 preset channels, the guard frequency (channel G) and MANUAL.

b. Three manual control switches. The first selects the first two figures (hundreds and tens) of the desired frequency in tens, the second selects the third figure (units) and the third selects the decimals. These switches are only operative when MANUAL is selected on the 20-position switch.

c. A volume control.

Para 1 6

11 12 14 16

20 21 22 23 25 30

31 36 38

d. A 7-position function switch, giving selection of:

OFF

T I R, normal transmission and reception

T /R + G, with guard frequency superimposed on reception

grT } inoperative

T/R ON-DjL OFF

3. On the port console, beside the radio altimeter limit switch, are the aerial selector and the tone switches. The aerial selector switch is labelled RT1 UPPER/RT2 UPPER. When set to RT1 UPPER, RT'1 is connected to the upper aerial and RT2 to the lower. When RT2 UPPER is selected, the connections are reversed.

4. It is possible to reset the preset channels in the air if necessary. To do so, the cover plate on the control unit is removed by undoing the milled screws. The channel selector is then turned until the number to be selected is indicated on the resetting panel (this will not correspond to the selector switch indication). The resetting pins are then moved to the required positions and the cover plate closed.

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Page 1 (AL 8)

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5. Power Supplies. RT1 (PTR 175) uses 28 volt DC from the essential bus bar.

UHF ARC/52 (RT2) 6. The UHF transmitter/receiver is in the nosewheel bay and the controls are on the 1st pilot's port console. It is possible to select 1750 channels at 0 ·1 MHz intervals or 19 preset channels, one of which is tuned to 243 MHz. In addition, a separate receiver allows a guard frequency of 243 MHz to be superimposed on any selected channel. MCW transmission is available if required.

7. The remote control unit on the port console, operated by the 1st pilot, carries the following controls:

a. A 20-position rotary switch, giving selection of 18 preset channels, the guard frequency (channel G) and MANUAL.

b. Four manual control switches. The first selects either 200 or 300 MHz, the second selects 0 to 90 MHz in tens, the third selects 0 to 9 MHz and the fourth 0 to 0 · 9 MHz. These switches are only operative when MANUAL is selected on the 20-position switch.

c. A volume control.

d. A 4-position function switch, giving selection of:

OFF T /R, normal transmission and reception T JR +G, with the guard frequency superimposed on reception ADF, inoperative,

8. The preset channels may be reset in the air in nhe same way as those of rhe RTl.

9. Two ONJOFF tone switches are on the port con­sole, RT1 for the PTR 175 and RT2 for the ARC/52.

10. Power Supplies. RT2 (ARCJSZ) uses 200 volts AC and 28 volts DC.

Tone Release 11. Tone release facilities are available for simulated bombing practice and may be obtained from either VHF or UHF. The controls are on the bombing con­trol panel and consist of a 3-position RTIJOFF JRT2 switch, start switch and a light. When the start switch is pressed, the service selected radiates a continuous 1 kHz note until rhe bomb is supposedly released.

HF 12. HF fSSB. The remote control unit is at the AEO position and carries a mode switch, digital fre­quency selection and RF gain control. SSB, HFRT and CW facilities are available. Also at the AEO posi­tion is the aerial remote control unit.

13. Power Supplies. HF JSSB uses 115 volt, 400 Hz, single phase AC and 28 volt from the essential bus bar.

lntercom 14. The intercom system operates through an A1961 amplifier, controlled by a NORMAL/EMERGENCY j OFF switch at the AEO position. In addition, a separate Al961 amplifier is provided for conference intercom and is controlled by an ON/OFF switch at the AEO position. If the normal intercom system fails, setting the control switch to EMERGENCY provides intercom facilities through the amplification stage of the RTl.

15. External intercom points are provided in clle nosewheel bay and in the tail section, the system being controlled by an ON/ISOLATE switch at the AEO position.

Station Boxes 16. All crew members have identical type 7681 station boxes, the controls on which allow any crew member to:

a. Select any one of five SPEAK/LISTEN services.

b. Mix incoming radio and intercom signals without interfering with selections made by other crew members.

c. Call all other crew members on intercom irrespective of the services they may have selected.

17. The station boxes carry the following controls:

~ a. Seven Listen Only Controls. These provide listening services on :

HF Conference If C Airborne Warning Normal I/C RTl RT2 ADF /TacanjiLS

b. A Rotary SPEAK/LISTEN Switch at the Bottom of the Box. This switch allows selection of the following five services:

Conference IJC HF RT2 RTl Normal I/C

c. A Spring-loaded CALL Switch on the Right of the SPEAK/LISTEN Switch. This switch is used, in conjunction with the intercom position of the SPEAK/LISTEN switch, to call all crew members, regardless of their selections.

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•

RESTRICTED AP lOlB-1902-15 Radio and Radar

d. A NORMAL/OFF /DIRECT Switch, to the Left of the SPEAK/LISTEN Switch. On NOR­MAL, incoming signals are fed through an amplifier, powered by a fused 28V supply. The fuse and a spare are on the front of the box. If fuse or amplifier failure occurs, selection of DIRECT by­passes the amplifier and switches it off. Only the selected SPEAK/LISTEN facility is available (at reduced volume) and the LISTEN ONLY controls are inoperative. The incoming call facility is lost, ie, the crew member affected can use his call switch to call other crew members but cannot be called by them.

18. Both pilots have press-to-transmit buttons on their control columns. The AEO has a press-to-ttansmit button and a morse key.

19. Control of Services. The control of the various services fed to the station boxes is:

a. Pilot (1) RT2 (ARC/52) selection. (2) ILSjThcan/ ADF audio selection. (3) RT1/RT2 aerial changeover switch. b. ABO (1) Airborne warning. (2) HF. (3) ILS ON/OFF switch. ( 4) ILS channel selection. (5) 1/C NORMAL/EMERGENCY JOFF switches. (6) Conference I/C ON/OFF switch. (7) ADF power switch. (8) Extema:l intercom. (9) RTl (PTR175). c. Navjplo.tter

ADF

NAVIGATIONAL RADIO AND RADAR

Radio Compass (ADF)

20. The radio compass and manua'l loop controllers are in '!lhe roof above the nav /plotter's position, while rhe selection of ILS, Tacan or A:DF audio signals is controlled by a switch on the 1st pilot's port console. A repeater indicator is on the co-pilot's instrument panel. An ON/OFF power switch on the ABO's panel oontrols ·the 28V DC and 11 SV 400 Hz single-phase supplies.

ILS

21. The ILS master ON/OFF and channel selector are 'at the ABO's position. The AUDIO ILS /T ACAN / ADF selector is on the 1st pilot's port console, beside t!he radio 'a'ltimeter limit switch. A marker light is on each pilot's instrument panel. There is no separate ILS indicator, the signals being fed into the MFS When the ILS is selected on and the MFS navigation selector is set to LOC & GP. The localiser signals are

then shown on the beam compasses and the glide path signals on the direcror horizons (see Ohapter 12). ILS signals are fed to the autopilot when the TRACK and GLIDE switches are pulled out. The system uses 28V DC (non-essential).

Tacan 22. The Tacan oonttoller and indicator are at the nav /plotter's station, with a repeater indicator on the pilots' centre panel. Audio signals may be obtained via <the ILS fT ACAN / ADF switch on the 1st pi1ot's console. The system uses llSV AC, from the star-

~ board main transformer, and 28V DC. Mod 2465 introduces a new Tacan aerial. ~

Doppler and GPI 23. Decca Doppler Type 72 provides groundspeed and drift which is oompounded with heading infor­mation in a groundspeed resolver to produce aircraft velocities in cartesian co-ordinates. The installation comprises an indicator control unit at the nav /plotter's station, and a computer and transmitter j receiver antennae system in the bay immediately behind the port main undercarriage. The system uses 115V single­phase, 400 Hz AC from No 4 busbar, and 28V DC.

24. GPI 6 or 6A at the nav jplotlter's position uses aircraft velocity in NJS and E/W components from the groundspeed resolver to drive a ground position Iatitude and longitude read-out. Velocity sources available to the GPI are obtained from the doppler, the HRS and the Navigation and Bombing Computation System. The GPI operates normally below latitude 84°; otherwise the grid navigation facility is used. The system uses llSV, 3-phase, 400Hz AC from the NBS transformer and 28V DC. GPI 6A provides the addi­tional facility of longitudinal oounter reversal.

RadW Altimeter Mk 7B 25. A radio altimeter Mk 7B measures t~he height of the aircraft AGL/ ASL up to a maximum of 5000 feet in two ranges, 0 to 500 feet, and 0 to 5000 feet. Instru­ment aocuracy on the 0 to 5000 foot range is ± 30 feet

~or -+- 3%, whichever is the greater. Accuracy on the 0 to 500 feet range is ± 3 feet or ± 3% whichever is the greater. ~

26. The master indicator is on the 1st pilot's flight instrument panel, above the VSI. A duP'licare indicator is at the navigator plotter's position. A limit indicator (consisting of 'three coloured lights) is on the inboard side of the co-pilot's panel. The transmitter/ receiver and the transmittingjre~iving aerials are in the lower starboard wing, out'board of the nosewheel bay. Power requirements are 115V AC from the port main trans­former, and 28V DC.

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Page 3 (AL 10)

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27. The controller on the l&t Pilot's port console carries:

a. An ON/OFF switch. b. A RANGE selector. c. A limit lights selector. d. A spring-loaded TEST swiroh.

28. The indicator has a red failure light at the bottom left-hand comer of the casing.

29. The altimeter is ready for use approximately one minute after being switched on. When the TEST switch is operated, t!he needle moves to 65 + 10 feet.

Radar Altimeter Mk 6A

30. The radar altimeter Mk 6A indicator 1 control unit on the nav /radar's upper panel is used as a reference for NBS height finding and for low-level terrain clearance. The system uses 28V DC, from the non-essential busbar, and 115V, single phase, 1600 Hz AC from No 2 frequency changer. The supply switch is on panel 12P, together with the oscillator tuning switch. The accuracy of this equipment below 10,000 feet is + 100 feet. It is inaccurate below 500 feet AGL/ ASL.

OPERATIONAL RADIO AND RADAR

NBC andH2S

31. The navigational bombing system (NBS Mk IA) is primarily a blind bornbing system but also incor­porates facilities for rapid fixing and homing. The system consists basically of the H2S Mk 9B and NBC Mk 2A. The controls are at the nav fradar's position.

32. The system uses 200V AC from No 4 (or No 1) busbar, llSV 400 Hz AC from No 1 transformer, llSV 1600Hz from No 2 (or No 1) frequency changer, 112V DC and 28V DC. The T AS unit uses llSV 400 Hz from the port maio transformer. The whole system is de-energised when automatic 'load shedding occurs.

33. The H2S components are: a. Indicator 301C. It carries the display in the CRT and associated controls. b. Wave Form Generator 68C. Carries the circuitry for producing the correct timebase waveforms. c. Analyser Unit. Establishes North-stabilised PPI rotating t'imebase. Provides plan range information for calculator SA and bearing angle values for production of a bearing marker on t and k million ·scales and a heading marker on t and 1 million scales on PPI display. d. Scanner Assembly. Mounted on a gyro-stabi­lised platform which. also carries the modulator transmitter f receiver and wave guides.

e. Calculator Type 5A. An electro-mechanical device which solves the plan-rangejslant-rangef height triangle, and produces rhe t and i million scale range markers. f. Control Unit 626. A joystick control used to apply movement to <t>he display. g. Control Unit 595. Carries scanner and trans­mitter controls.

~ h. Control Unit 6204. The swept gain unit has adjustable preset'S for optimum gain at the higher gain levels. SEM 046 (Mod 2500) in the Vulcan B Mk 2 (MRR) introduces an IF BIAS voltmeter so that the presets, in conjunction with a calibration card for the particular receiver, can be set simply ·to a predetermined optimum gain for prevailing conditions. ~

34: The NBC Mk 2A components are: a. Calculators Types 1 and 2. Synchronous in­formation is fed into these units to combine with air speed and wind information for DR navigation and steering and range computing. b. Calculator Type 3. Carries the height and air­speed servo mechanisms and the ballistic computers. c. Navigation Panel. The display end of the system, giving up-to-date information of T AS, track, groundspeed, latitude, longitude, range along track, time to go and wind velocity. It also provides a facility for monitoring computer voltages. d. Control Unit 585. The maio control unit of the NBS, containing a function switch with associated relays, shift work potentiometers with associated mechanical systems and offset potentiometers. It also carries height setting controls.

35. Signals from the NBS or the bomb aimer's tum control can be fed to the MFS and the autopilot (see Chapter 12).

ECM 36. The ECM inStallation is controlled by the AEO. Audio signals can be superimposed on the intercom by selection of the appropriate control on the station boxes. Alarm signals are fed to all crew station boxes when the AIRBORNE WARNING volume control is selected. The MONITOR/ ALARM switches are inoperative.

37. The power supplies required for the instaUation are 200V, 3-phase 400 Hz AC from all four busbars and 28V DC. If any two alternators fail, part of the ECM loads are shed automatically and the remainder when the RAT is pulled. Full ECM cannot be regained unless at least three alternators are running normally ; some of these loads are available if the NORMAL/ RESET switch is put to RESET.

IFF 38. IFF fSSR. IFF /SSR 1520 controls at the AEO position include :

1-14 Page 4 RESTRICTED

RESTRICTED AP 101B-1902-15 Radio and Radar

a. An OFF fSBY f LOW fNORMfEMGY switch, with the following functions:

OFF: No power supplied.

SBY: Equipment is warmed up and receiver switched on. Ready to transmit after 50 seconds.

LOW: Operating on reduced sensitivity.

NORM: Full .transmission and reception.

EMGY: Transponder accepts Modes 1, 2, 3A or B interrogations, whether selected or not and transmits emergency replies.

b. Four toggle switches, MODE 1, 2, C, D and a rotary switch 3AfOFF fB, for mode selection. Modes 1, 2 and 3 are military modes and the others are civil, modes 3 and A being identical. Mode C is provided to pass coded height info11mation from the co-pilot's Mk 30A altimeter. Mode D has not yet been allocated a purpose and is reserved for future expansion.

c. A CIVIL/ MIL switch, which is only operative with EMGY selected and controls the replies on Modes 3A and B. d An IP switch (identification of position), spring-

~ loaded to off (centre). Post-SEM 054, an additional IFF IP IDENT switch is fitted. ~

e. Two code selectors, for Modes 1 and 3. f. A TEST pushbutton embodying a green light. When pressed, a self-test system is initiated and, if the operation is satisfactory, the green light comes on.

g. The system uses one phase (115V) from No 4 busbar 200V AC and 28V DC.

39. A red IFF FAIL light is on the panel in front of the AEO. The light comes on:

a. When the set is switched off (steady). b. If the equipment is at SBY and the test button is pressed (flashing) or if the transponder is being interrogated (flashing). c. If any faults occur (steady). d. If the TEST button is pressed when LOW is selected (steady).

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Page 5 (AL 10)

RESTRICTED AP lOlB-1902-15

PART I

CHAPTER 15-TERRAIN FOLLOWING RADAR

Contents

General

General ...

DESCRIPTION TRFU Range Loop Height Loop Programme (Attitude Hold) Pitch Rate Control Climb High Function Acceleration Limiters Safety Devices ... ADD Electrical Supplies

OPERATION In-flight ... Malfunctions

Illustration TFR System Diagram

1. The TFR system is designed to enable the pilot to follow approximately the contours of the terrain at a height selected in flight. The components of the system are:

a. A radar pod in the nose (TFRU), containing a transmitter/receiver, an adjustable antenna, a pitch rate gyro and computing circuits. The pod is cooled

~ by air from the H2S radome. ~

b. Airstream direction detectors (ADD). Two slotted probes, protruding horizontally from each side of the nose, measure the angle of attack.

c. An airspeed transducer, supplied from the air­craft starboard pitot-static system.

d. A control unit, on the 1st pilot's console.

e. Indicator and warning lights.

Para 1

3 4 5 7 8 9

10 11 12 14

15 20

Fig 1

2. The system is used in conjunction with the follow­ing items of aircraft equipment:

a. MFS port vertical gyro

b. MFS pitch computer Chapter 12 c. Glide path pointer of the director horizon

d. Radio altimeter M.k 7B (Chapter 14).

DESCRIPTION TFRU

3. The radar pod measures the slant range forward of the aircraft against a preset datum. The antenna angle is adjusted by the height selector and the angle of attack measured by the ADD. Dive demands are initiated when range exceeds the datum and climb demands when the range is less than the datum. Com­puters in the unit analyse and compare the various signals and pass them to the pilots' instruments as dive or climb signals.

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Page 1 (AL 10)

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Range Loop

4. The measured slant range is fed to a range com­puter which feeds a pitch rate demand signal, pro­

~ portional to the range error, to the majority circuit para 8). ~

Height Loop

5. Over flat terrain or calm water, radar returns may be insufficient to activate the range loop. To prevent this apparent loss of range initiating a dive command, height information from the Rad Alt 7B is modified by aircraft attitude and antenna angle signals, to pro­vide a pitch rate command to the majority circuit.

6. When intending to use the TFR below 500 feet the Rad Alt 7B should not be selected to the 500 feet scale until below 500 feet to avoid spurious climb demands. When the aircraft reaches the set TFR height, Rad Alt 7B can be selected to the 500 feet range to provide improved accuracy when the height loop is in operation.

Programme (Attitude Hold)

7. As the aircraft climbs towards the summit of a hill, the radar beam clears the crest. To prevent a dive signal being given before the aircraft is over the summit, a delay computer demands a steady attitude computed from the speed, pitch angle, angle of attack and last measured range. After this delay, other sub­systems take command and, in practice, the most posi­tive pitch rate command is a minus 0 · 5g (applied) dive rate limit known as the pushover. A programme is not initiated if the signals increase progressively through the maximum range as opposed to disappear­ing instantly. This prevents inadvertent programme being caused by aircraft pitch-up.

Pitch Rate Control

8. Signals from the range loop, the height loop, the programme, the dive rate limiter and the fail-safe system are fed into a majority circuit computer, which selects the highest climb demand of these signals. The majority circuit output goes to a comparator as a pitch rate demand for comparison with the output from the pitch rate gyro; any difference is the final pitch rate

~ command fed to the MFS director horizons via the MFS pitch computer when the MFS/TFR switch is selected to TFR. Demands are shown on the pitch

director (the GP pointer during non-TFR operations), and the pitch scale is in continuous fast-chase mode. ~

Climb High Function

9. If for any reason the aircraft pitch angle is greater than 4° nose-up, the antenna is off-set down­wards, thus tending to shorten the range returns and causing the aircraft to fly higher, the effect increasing in proportion to the gradient. This prevents ballooning on the far side of a summit and gives a greater safety margin over steeply rising ground.

Acceleration Limiters

10. To limit acceleration to between +O · 75g and minus O· 5g (applied), airspeed signals are fed into two acceleration limiters. Signals from the limiters limit the pitch rate demands from the height loop to +0·75g (applied) and from the dive rate limiter to minus 0 · 5g (applied).

Safety Devices

11. Fourteen internal monitors continuously check range loop, antenna servo loop, pitch and pitch rate circuitry, as well as the monitoring system itself. If any monitor detects a failure the green light goes out, the red FAIL light comes on and a pitch up of 1 · 89°1 second is demanded. Those circuits which cannot be monitored are duplicated, both operating simul­taneously and either one capable of operating the system. Failure of the radio altimeter activates the fail-safe circuit.

ADD

12. The ADD provide angle of attack information by rotating two slotted probes to maintain equal pressure either side of internally mounted vanes. This rotation is measured by a potentiometer and passed to the TFRU, where antenna angle is correspondingly adjusted. To compensate for airflow distortion the ADD outputs are paralleled and a mean signal taken. The ADD probes are electrically heated via the pressure head heater switches.

13. A discrepancy between the two ADD brings on the ADD MONITOR amber light at the AEO's position and initiates TFR failure indications. I\,~ I';M ""()"'~ fault has cleared the warnings may be cancelledi_~the ~---.--J AEO's by pressing the ADD monitor light and the pilots' by selecting RESET. A TEST /NORMAL

1-15 Page 2 RESTRICTED

e

.... I .... V\

~ :;d trl

.... Cll - . ~ ~ :;d 1!:1 ......

!:11:1 Q {I)

gs '< "' ... (l> a t::' ;· ~ a

"0,....

tfJI (),._.

WVl

B

'~,

I I ;--a I < >

-

111T::~.s·111 I

TFR GO LIGHT

I ---· I I I ;

e ·

RANGE LOOP

HEIGHT LOOP

PROGRAMME (ATTITUDE HOLD)

FAILSAFE

PITCH RATE CONTROL

!MAJORITY H COMPARATOR I CIRCUIT

II POWER Ill PITCH SUPPL I ES ~~~~

TFR WARNING LIGHT

REDUNDANT CIRCUIT FAIL LIGHT

e

;;l ~ er cil?d :::: ...... oo ~ ...... -· t:d ::l •

0<1 .....

~:g o:>N P.• ., ...... .... Vl

RESTRICTED AP 101B-1902-15 Terrain Following Radar

Table 1 Controls and Indicators

Control f Indicator Location

SUPPLY, ON/OFF switch Port console . ..

Radar control STAND BY I Port console ... ON switch

Magnetic indicator Port console ...

Height selector Port console ...

FAIL-SAFE, TEST /NOR- Port console ... MAL/RESET switch

TFR VIDEO or GO lights One each pilot's panel (green), press-to-test

TFR FAIL lights (red), One each pilot's panel press-to-test

TFR WARNING lights One each pilot's panel (amber), press-to-test

MFS /TFR mode selector switch

Redundant circuit fail light, press-to-test

Co-pilot's panel

On rear crew panel, between navigator and AEO

Function

Energises relays which connect 200V AC and 28V DC to the TFRU. The radar is then ready to be switched on after approx one minute delay

At STANDBY, radar is on but not transmitting. At ON, radar transmits subject to the delay above

Shows OFF when supply switch is OFF, STBY when radar is in the standby mode and black when control switch is ON and the radar is transmitting

Enables required terrain following height to be selected in 100 foot increments

At TEST, a simulated fault is fed to fail-safe circuits; TFR FAIL lights come on and max climb demand is shown on pilots' indicators. RESET position, used to cancel test or transient warnings, does not cancel warnings if a genuine fault exists. Spring-loaded from

~ RESET to NORMAL. An additional TFR RESET button to the right of the throttle quadrant gives the co-pilot an independent facility to reset the system after failure indication. ~

Come on in the air when the radar is on and monitored circuits are serviceable

Come on if a failure is detected in any of the monitored circuits, provided that the control switch is ON. Should

· be accompanied by a maximum climb demand

Flash when a nose-up demand of 1° /sec has been ignored for three seconds (greater demand reduces period). Goes out when demand is less than 1° /sec. Also shines steadily when TFR first switched on

When set to TFR, demand signals are fed to the glide path pointers on the director horizons and the pitch scale is in the fast chase mode.

WARNING : If the switch is set to TFR when the radar is off or the pod is not fitted, the pitch scale remains

~ in fast chase and the GP pointer parks at the top of the scale. Must be at MFS for normal aircraft operation. ~

Comes .on if there is a discrepancy of 1· 2° /sec or more for three seconds between the duplicated circuits of the TFRU

ADD fail light, press-to­reset

Above redundant cuit fail light

cir- Comes on if there is discrepancy between ADD outputs

ADD warning TEST /NOR- ADD trim panel MAL switch

~ (Both lights have an adjustable iris post-SEM 058) ~

For ground test purposes only. Must be set to NORMAL before flight

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Page 5 (AL 10)

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switch on the ADD trim panel below the pilots' floor must be at NORMAL 'before flight. It should be noted that any yaw or sideslip may give an ADD warning and that the ADD warning is operative whether TFR is in the standby condition or on.

Electrical Supplies

14. The TFRU uses 200V, 3-phase, 400Hz AC from No 4 :busbar via panel 75P. The system also uses 28V DC from panel 4P. The ADD probes are heated by 28V DC, via the pressure.Jhead heater switches. The warning lights use 28V DC from the TPRU while their press-to-test facility is supplied from the main aircraft system. The radio altimeter uses USV AC from the port main transformer. If No 1 or No 4 a:lternator fails or is switched off, the TFR may fail but can be reset immediately.

OPERATION

In-Flight

15. Carry out the cheeks in t!he Flight Reference Oards.

16. If the system is switclled on above 20,000 feet pressure altitude, the TFR remains in t!he standby condition. The fail-safe system is energised by loss of in-track signals from the radio altimeter. The system must be selected to STANDBY at least 10 minutes before use.

17. With the system on and serviceable and the air­craft height within limits for the radio altimeter, demand signals are presented on t!he director horizons. By following 'these demands, the pilot can maintain the aircraft at .the correct height above the rerrain, with the foMOWiing exceptions :

a. Slim towers and/or transmission lines may not be detected.

b. Let-downs over smooth water may result in an undershoot of up to 50 feet below the set height. \

~A

c. Drift angles of more than f.0 may result in the radar not detecting obstacles in the flight path.

d. Terrain following performance deteriorates with bank angles in excess of 20° over mild terrain and in excess of 10° over rough terrain.

e. Clouds with heavy moisture content, or heavy precipitation, may appear as range returns and result in climb commands.

f. 'l'he system design logic can be defeated and reduced hill clearances obtained, by certain com­binations of hills at spacings of less than one NM. The extent of such reductions varies depending upon aircraft speed, height, setting, hill height, and till separation but between the nominal velocities 250 and 350 knors, peak clearances are within 50% and 200% of set height. In all but the worst cases, clearances are within 80% and 140% of the set height.

g. Rapid accelerations when height loop signals are being followed (para 5) may give dive demands resulting in a height loss of up to 200 feet.

18. During pushover, do not alter heading.

19. Pilots should resist any tendency to 'interpret' TFR demands or apply any prediction to them. The TFR already produces the required pitch rate demands, therefore such techniques are unnecessary and can lead to dangerously low terrain clearance.

Malfunctions

20. If any monitored part of the system fai:ls, the fail­safe circuit comes into operation, the TFR FAIL light comes on and a max climb should be demanded. If, at any time, the TFR FAIL light comes on, a climb must be initiated even if a maximum climb has not been demanded. If the failure is transient, the system can be reset by putting the TEST /NORMAL/RESET switch to RESET momentarily. If the failure occurs during a turn, the FAIL light comes on but the climb demand may be masked by the pitch rate of the turn. If failure occurs in the TFRU, the radio altimeter con­tinues to function.

21. If one of t!he duplicated circuits fails, the system continues to function on a single circuit but the redundant circuit fail light comes on. This condition cannot be reset manually, but if the circuits resume compatibility the light will go out.

22. If both ADD fail me scanner is driven to its operational limit and the limit switches initiate full failure warnings.

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RESTRICTED AP 101B-1902-15 Terrain following Radar

35. If the 28 volt supply to the TFRU fails, there wHI be neither green nor red indicator lights. This can be checked by operating the press-to-test facility for the lights, as the DC supply for this is from a separate source; if the lights come on, the TFRU DC supply is at fault.

36. If the port MFS gyro fails, a large pitch input may be given to TFRU. If flying over smooth water (ie inadequate radar returns), TFR commands result

in incorrect height separation. A TFR failure signal may not be given but the MFS attitude failure flag should show.

37. Whenever MFS attitude failure warning is given, the TFR should be switched off.

38. If the TFR warning light comes on, 1evel the wings, apply 80% power and initiate a climb imme­diately, regardless of TFR demands.

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e

PART 2

LIMITATIONS

List of Chapters

Title

AIRFRAME LIMITATIONS

ENGINE LIMITATIONS

MISCELLANEOUS LIMITATIONS

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AP lOlB-1902-15

Chap.

I

2

3

• RESTRICT ED AP lOlB-1902-15

PART 2

CHAPTER I-AIRFRAME LIMITATIONS

Contents

General Speed and Mach Number Limitations Crosswind Limitations G Limitations Weight Limitations CG Limitations Aircraft Approach Limitations Autopilot Limitations Air-to-Air Refuelling Limitations ... Airframe Anti-icing Limitations Bomb Bay ...

Para

1 4 7 8 9

12 13 14 16 19 20

The limitations given in this Part are taken from the Release to Service ~ Document to AL 69 standard. The Release to Service Document must be~

consulted to ascertain the latest release standard.

General 1. The Vulcan B Mk 2 is designed for manoeuvres appropriate to the role of a medium bomber, in world­wide conditions. Aerobatics, stalling and spinning are prohibited. Speed must not be reduced below that for the onset of pre-stall buffet and in any case not below the threshold speed for the weight less 5 knots. In manoeuvres at altitude, acceleration should not be increased beyond that for the onset of buffet where this occurs before maximum g is attained.

2. Height Limitations. There is no height restriction on the aircraft because of airframe limitations. How­ever, the maximum operating altitude is limited by the oxygen equipment, as follows :

Pres- Max sure Anti-c Cabin

Regulator 7erkin Suit Alt (ft) Remarks

Mk 17F No No 50,000 -Mk 21A Yes Yes 56,000 Oxygen contents more orB than i of total

Mk 21A Yes No 52,000 -orB

Mk 21A No No 45,000 Above this altitude, orB oxygen pressure is

Mk 2 or above 30 mm Hg 2A (max l u n gs can

stand in comfort)

Mk 3 or No No 50,000 -3A

Mk 2 or Yes Yes 56,000 -2A

Note: In the worst case, eg following loss of the canopy, or if the entrance door opens, aerodynamic suction can cause the cabin altitude to exceed the aircraft altitude by up to 5000 feet.

3. Arresting Gear. The aircraft has unrestricted clearance to trample the following runway arresting gears, provided that the configuration has not been altered by modifications later than mod 2240:

a. SPRAG. b. CHAG. c. RHAG Mk 1. d. PUAG Mk 21. e. Bliss BAK 9, BAK 12, 500S.

Speed and Mach Number Limitations (see also Release to Service)

4. a. With all PFC working and all autostabilisers operative: Maximum speed above 15,000 feet-330 knots or 0 · 93M (0 · 92 with Mk 301 engines), which- AJ,43 ever is less. (Elevator forces are not to be trimmed :r-' out above 0 · QOM 'I ""~ s.~ ~ 15ca::>P-r · 1iJu.

.b-.-r~ ... ~ .:.Ffl-T€-~~5 pgm~ ~ ~ b. With one or more PFC inoperative-0·90M.

c. With one or both servos of the mach trimmer inoperative-0 · 90M unless specifically authorised. If specifically authorised, 0 · 93M (0 · 92M Mk 301 engines).

d. With one pitch damper inoperative-0·93M (0 · 92, Mk 301 engines).

e. With two or more pitch dampers inoperative--0·90M.

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f. With feel relieved or failed-250 knots or 0 · 90M, whichever is less. E:l'fitreme care is necessary to avoid overstressing the aircraft.

5. Maximum Speeds for Operation of the Services. The speed for operating a service also applies to flight with the service extended.

a. Airbrakes

b. Bomb doors ...

c. Undercarriage

d. RAT

e. Tail parachute streaming

No restriction.

Up to the normal limiting speed of the aircraft.

270 knots (0 · 90M above 40,000 feet).

330 knots or 0·93M (0·92, Mk 301 engines).

145 knots (max). Any parachute streamed above 13S knots is to be examined before re-use.

f. The tail parachute must be jettisoned at speeds between 50 and 60 knots. In an emergency the parachute may be retained until the aircraft has stopped.

6. For low level limitations, see Release to Service.

Crosswind Limitations

7. Maximum crosswind component for take-off, landing or streaming brake parachute: 20 knots.

G Limitations

8. The following accelerometer Teadings are not to be exceeded:

Max Indicated g with AUW Aileron Angle

(lb) IMN Negligible Large

Angle Angle

Up to Up to 0·89 2·0 1·8 160,000 0·89 to 0·93 1·8 Prohibited

160,000 to Up to 0·89 1·8 1·5 190,000 0·89 to 0·93 1·5 Prohibited

Above Up to 0·93 1·5 Gende man-190,000 oeuvres only

Note 1: Full aileron may be applied up to the indicated Mach numbers quoted, but aileron is not to be applied rapidly.

Note 2: Manoeuvres involving simultaneous application of large aileron angles and normal acceleration are not to be executed at indicated Mach numbers greater than 0·89.

Note 3: Manoeuvres under zero or negative g conditions are prohibited.

Weight Limitations 9. Maximum for <take-off and emer-

gency landing Normal landing ...

204,000 lb 140,000 lb

10. If, in emergency, the aircraft is landed at 195,000 lb or more, the rate of descent at touch-down must be kept to a minimum and the angle of bank on the approach must not exceed lS0 •

11. Simulated asymmetric flying is not permitted at weights above 195,000 lb.

CG Limitations

12. The CG limitations (undercarriage down) are: a. At weights up to 142 to 156·9 inches

195,000 lb aft of datum b. At weights above

195,000 lb 148 to 151·3 inches

aft of datum

Aircraft Approach Limitations 13. The aircraft approach limitations are:

T1-ue Height (Above Touch­

down) a. Precision radar 250 feet

ILS, Auto or Manual* (in-line localiser) 250 feet

ii,LS, Auto or Manual* (off-set localiser) 270 feet

*It is advisable that all ILS approaches should be Tadar 01onitored.

b. Visual Committal Heights (VCH) True

Height

One engine inoperative 150 feet

Two engines inoperative

c. Engine Out Allowance (BOA) One engine out

200 feet

Two engines out (up to 18S,OOO lb) Two engines out (above 185,000 lb)

0 feet SO feet

100 feet

Autopilot Limitations 14. The autopilot limitations are:

a. Speeds Maximum airspeed Maximum mach number

Mach trimmer operative Mach trimmer inoperative

350 knots

0·90M 0·87M

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I RESTRICTED AP 101B-1902-15

Airframe Limitations

Maximum airspeed with TRACK and LOC & GP selected (ie feel partially relieved) 180 knots

b. Minimum altitude (except dur-ing ILS approach) 1500 feet AGL

c. The artificial feel must be functioning correctly.

15. With the autopilot engaged, the following con­ditions must be observed:

a. Longitudinal trim is to be maintained so that the autopi1ot trim indicator is within the safe range.

b. One pilot is to be strapped into his seat at all times.

c. The autopilot may be used with the elevator channel disengaged (if operaoionally essential) ; in the ILS mode the elevator channel must be engaged. Neither the aileron, nor the rudder channel may be disengaged separately.

Air-to-Air Refuelling Limitations

Note: Although the aircraft is cleared for air-to-air refuelling to the following standard, there is no opera­tional requirement at present and therefore it is not to be practised.

16. The aircraft is cleared for air-to-air refuelling by day and by night, using Victor tankers and by day only using Boeing KC135 tankers, with Mk 8 equip­ment, subject to the following conditions:

a. Speed. The speed of the tanker at and during contact should be:

(1) KC135 ...

(2) Victor maximum speed

Minimum speed

Maximum altitude

260 to 275 knots (at heights up to 30,000 feet)

300 knots up to 31,000 feet

0 · 80M between 31,000 feet and 35,000 feet

220 knots

35,000 feet

b. Airbrakes. MEDIUM drag airbrakes are recommended for contact.

c. Fuel Trim. The CG control switches may be used to maintain the fuel CG. Contact must be broken if either needle of the CG indicator gQes into the red sector. The A and E tank gauges are monitored during refuelling to ensure that a spurious EMF has not opened the refuelling valves.

d. Night Contacts (Victor). Before flights involv­ing night contacts, the probe lighting must be cor­rectly focussed on the forward third of the probe.

17. Only one transfer pump in the KC135 is to be used.

18. All the necessary flight refuelling system modi­fications must be embodied in the aircraft, including the Mk 8 probe and probe lighting, and the modi­fication which ensures the ground refuelling master switch is off.

Airframe Anti-Icing Limitations 19. Subject to the embodiment of all necessary modi fications, the wing anti-icing system may be used at all levels but the fin system must remain inoperative. For the physical limitations of the system and for operating instructions, see Part 1, Chapter 11 and Part 3,

~ Chapter 5. ~

Bomb Bay

20. Bomb Bay Tanks. The bomb bay tanks may be used in the following configurations:

a. Two cylindrical tanks

b. One cylindrical tank in forward position

c. One A tank in forward position and one cylin­drical tank in rear position

d. One A tank in the forward position of the two forward locations

e. One A tank forward and one E tank aft.

21. Bomb Bay Panniers. 750 lb and 4000 lb panniers may be carried with bomb bay fuel tanks to the follow­ing configurations:

Forward Centre Aft

Cylindrical 750 lb Pannier Cylindrical Tank Tank

A Tank 750 lb Pannier E Tank

4000 lb Pannier

4000 lb Pannier

Cylindrical Tank

A Tank

750 lb Pannier

4000 lb Pannier

4000 lb Pannier

22. Opening of Bomb Doors in Flight. The bomb doors may be opened in flight:

a. When the tanks have been emptied of all usable fuel.

b. During an operational emergency with fuel in the tanks.

c. Provided no panniers are fitted in the bomb bay.

d. Provided that time with the bomb doors open is kept to a minimum.

23. The weapons simulator must not be used on the ground if there are any signs of fuel leaks and may not be switched on in the air until the tanks are empty of all usable fuel.

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PART 2

CHAPTER 2-ENGINE LIMITATIONS

Contents

Fuel and Oi.l Specifications

RPM and JPT Limitations, Olympus Engines

Oil Pressure Limitations, Olympus Engines

AAPP JPT Limitations

AAPP Oil Limitations

Engine Anti-icing

lllustrations

Resonance Frequency Band

Fuel and Oil Specifications

1. The permitted fuel and oil specifications are given at Preliminaries, pages 9 and 10.

RPM and JPT Limitations, Olympus Engines

2. Mk 200/Mk 301 Series Engines. RPM and JPT for Mk 301 engines are shown in brackets, where different.

Engine Speed Max

Condition TimeLimil %RPM 1PT oc

Maximum for take-off and operational necessity 10 mins 101 (100) 670 (625)

Maximum continuous Unlimited 97·5 610 (570)

Ground idling minimum 29•5 (24•5) 610 (570) (alternator on Unlimited (minimum load) ISA, SL)

Overs peed 20 sees 104 (107) -During start - - 700

3. Relighting the Mk 200 series or ~ 301 engines should not normally be attempted above 35,000 feet.

4. To avoid resonant frequencies which could affect ~engine fatigue life, the RPM band 95% + lt% is to

be avoided up to FL300 (Fig 1). Furthermore, on the Mk 301 engines the RPM band 78% to 85% is also to be avoided below 5000 feet. The engine handling

Para

1

2

5

7

9

10

Fig

1

procedures recommended in Part 3 of this Manual meet the restrictions placed on 200 series engines. For Mk 301 engines, whenever 80% RPM (below 5000 feet) is recommended, 78% RPM should be used if the OAT is 15° or below and 85% RPM if the OAT is above l5°C. ~

Oil System Limitations, Olympus Engines

5. Mk 200 Series Engines

a. Norma:! lilt 90% RPM and above

b. Minimum at 90% RPM and above:

Sea level to 20,000 feet Above 20,000 feet

c. Maximum consumption rate

d. Minimum oil tempera,ture for starting

6. Mk 301 Engines

a. Normal at 90% RPM and

55 to 60 PSI

50 PSI 45 PS'I

1! pt/hr

above 55 to 65 PSI

b. Minimum at 90% RPM and above

Sea level to 20,000 feet Above 20,000 feet ...

c. Maximum consumption rate

d. Minimum oil temperature for starting

50 PSI 45 PSI

It pt/hr

minus 26°C

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Page 1 (AL 10)

RESTRicrED

AAPP JPT Limitations

7. The following maximum JPT limitations apply to fiight and ground running under any load:

a. For max of S minutes

b. For max of 1 hour

c. Continuous

d. During AAPP start High transient for 10 seconds max

8. When ground running, the AAPP must be stopped every 10 hours for an oil check.

AAPP Oil Limitations 9. The minimum oil pressure is 4 PSI. The pressure varies rapidly with temperature and should not be less than 12 PSI after a cold start.

Engine Anti-Icing 10. If icing conditions are met, the anti-icing must be used and the instructions in Part 1, Chapter 11 and Part 3, Chapter S are to be followed.

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RESTRICTED AP 1018-1902-15

PART 2

CHAPTER 3-MISCELLANEOUS LIMITATIONS

Contents

Navigational, Operational and Radio Limitations

Electrical System Limitations

Para 1

11

Navigational, Operational and Radio Limitations 1. NBS Mk 1A. The NBS Mk 1A is cleared for use as a navigational and bombing system.

2. Radio Compass. The radio compass is cleared for unrestricted use.

3. HRS. The HRS is cleared for use, subject to the following conditions :

a. During start-up, with fast erection, HRS is to be switched on at least 2i- minutes before taxying.

b. If airborne starting, the MRG is not to be switched on until initial take-off acceleration is over. The aircraft is to be kept on a steady heading for at least 2~- minutes while the MRG runs up.

c. Unless it fails, the MRG is not to be switched off until the aircraft is stationary.

4. Decca Doppler. Decca doppler is cleared for ~ unrestricted use. With the undercarriage down it will

operate only in the memory mode. It is compatible with the autopilot when operating normally in the track steering mode. However, prior to doppler unlock spurious drift signals may cause the aircraft to take up a false heading. ~

5. ECM. The ECM installation is cleared for use up to 55,000 feet in temperate climates.

6. MFS. The MFS is cleared for Service use.

7. IFF. IFF JSSR is installed and is cleared for unrestricted use.

8. Tacan. Tacan is cleared for use. ..~

9. Radio Altimeter. The following limitations apply to the radio altimeter :

a. Mk 7B High range Low range

b. Mk 6A

100 to 5000 feet 0 to 500 feet. Unrestricted use above 5000 feet

10. Radio Communications a. The PTR 175 is cleared for unrestricted use. b. The ARC 52 is cleared for unrestricted use. c. The HF /SSB is cleared for tmrestricted use.

Electrical System Limitations 11. The maximum continuous load per alternator is 32 kW, subject to a maximum continuous CSDU oil temperature of 120°C. This oil temperature must not be exceeded and, if necessary, height or ~oading must be reduced to keep the oil temperature within limits.

12. AAPP a. Operating altitude... Ground level to 30,000

feet, undercarriage up

Ground level to 5000 feet, undercarriage down

b. There is no guarantee of a successful start above 30,000 feet. The alternator must never be put on load above 30,000 feet.

c. Maximum Loads Altitude

feet 10,000 to 30,000,

undercarriage up

0 to 10,000, undercarriage up

0 to 5000, undercarriage down

Ground running

Load Time kW mins

17 30

32 30

23 4 32 (up to 45°C

OAT)

d. Air Bleed. Maximum electrical load for AAPP with airbleed is 10 kW.

13. RAT Operating altitude ...

Max load ...

Speed range

Time limit ...

20,000 to 60,000 feet

17 kW

Maximum speed 0 · 93M Minimum speed 0·85M/

250 knots, whichever is greater

10 minutes on load above 30,000 feet

10 minutes off load above 50,000 feet

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• RESTRICTED

PART 3

HANDLING

List of Chapters

Title

STARTING, TAXYING AND TAKE-OFF

HANDLING IN FLIGHT

CIRCUIT AND LANDING PROCEDURES ...

ASY~ETRIC FLYING

OPERATING IN ICING CONDITIONS

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AP IOlB-1902-15

Chap

1

2

3

4

5

(AL 8)

RESTRICTED AP 101B-1902-1S

PART 3

CHAPTER I-STARTING, TAXYING AND TAKE-OFF

Contents

General

Ground Starting the AAPP

Starting a Main Engine Using an External Air Supply

Para

1

3

8

Starting the Remaining Engines Individually Using Air Crossbleed

Quick Starting of Engines

13

16 18

21

24

29 30

31 37

40

43

General

Rapid Starting of Engines

Starting Failures-Normal Start

Starting Failures-Rapid Start

Dry Motoring Cycle

After Starting all Engines (all Conditions)

Taxying

Take-off

After Take-Off

Aborted Take-Off Procedure

Climbing

1. Throughout, it must be remembered that the limitations at Part 2 must be observed, and that the relevant checks in the Flight Reference Cards must be made at the appropriate times.

2. The air supply required for starting main engines may be obtained from an external source (Palouste) or by crossbleed from a running engine. Alternatively, engines may be started individually or simultaneously using air from the rapid start compressed air installa­tion. As any one, or a combination of more than one of these sources of starting air may be used to start the main engines, the starting order may be varied as required. However, the airflow patterns into the com­bined main engine air intakes make it advisable to use an outboard rather than an inboard engine to supply air for starting the remaining engines.

Ground Starting the AAPP 3. The electrical supply required for ground starting the AAPP may be obtained from an external 200 volt supply providing 28 volt through an aircraft TRU, or from an external 28 volt supply. H no external supplies are available, the internal aircraft battery may be used to provide a 24 volt supply.

4. The AAPP ground starting checks are given in the Flight Reference Cards.

... ~ 45 ~

5. Failure to Start. If the AAPP fails to start satis­factorily, or if any malfunction occurs during starting, set the MASTER SWITCH to OFF.

6. AAPP Dry Motoring Cycle. For an AAPP dry motoring cycle, the HP cock override and ignition isolation switches must be held down prior to and dur­ing the whole of the starting drill ; the switches must not be released until the master switch is off.

7. After Starting. When running satisfactorily, the AAPP may be used to provide electrical power together with air supplies for the A VS. If it is used to provide any air supplies, the electrical loading must not exceed 10 kW. There is no time limit on ground running the AAPP except that it must be stopped after 10 hours for the oil contents to be checked.

Starting a Main Engine Using an External Air Supply

8. When using an external air supply, the engines may be started in any order. If it is intended to start the remaining engines using air supplied by the run­ning engine, this must be No 1 or No 4 engine. After completing the relevant internal checks in the Flight Reference Cards, confirm with the crew chief that the external air supply is connected and that it is clear to start engines.

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9. The checks before starting are given in the Flight Reference Cards.

10. Press clle starter button and check that the indicator light in the button comes on, showing that the air control valve has opened. Wait for 10 seconds, checking that the oil pressure and RPM are rising, then move me HP cock lever towards the idling gate untH a fuel flow of 8 to 10 lb/min is achieved. When the JPT rises pause for one second and then move the HP cock lever slowly to the idling gate. During starting, the JPT normally rises to 300° C to 400°C and then falls to approx 250°C as the engine accelerates. If the rate of rise of the JPT indicates rthat 600°C may be exceeded, meter the fuel supply by moving the throttle lever slightly behind the idling gate until the JPT decreases and then slowly advancing it to the normal idling posi­tion. If the JPT continues to rise rapidly above 600°C and it appears likely that 700°C will be exceeded, close the HP cock and isolate the starter motor by switching OFF the engine master switch. After a normal start, the RPM may stabilise below the correct ground idling speed; in order to achieve the correct idling speed, advance the throttle slighcly and then return it to the idling gate. The starting cycle is terminated by the overspeed switch. The light in the starter button goes out when the air control valve has closed.

11. The checks during starting are given in the Flight Reference Cards.

12. Starting the Remaining Engines a. The three remaining engines may be started individually and in the same manner using air sup­plied by the Palouste.

b. The Paloustc may be removed and the three remaining engines may be started individually using air crossbleed from the running engine, provided that its speed is set to 70% RPM.

Starting the Remaining Engines Individually Using Air Crossbleed

13. The checks before starting are given in the Flight Reference Cards.

14. Set the RPM of the running engine to 70% and check that its engine air switch is open. Open the engine air switch of the engine to be started, and then start it in the manner described in para 10. The AEO should check the voltage and frequency of each alternator and switch on as required, noting that when three alternators are on line the pilots' red AL T FAIL light is steady and when the fourth alternator is switched on the pilots' AL T FAIL light goes out.

15. If, for any reason (eg sandy airfield, confined space), it is desirable to restrict RPM, when two engines are running their speed may be set to 60% RPM to start the remaining engines as an alternative to one engine at 70% RPM.

Quick Starting of Engines 16. After starting No 1 or No 4 engine, the three remaining engines may be started simultaneously using air from the running engine, provided that its speed is set to 93 % RPM. Carry out the checks in the Flight Reference Cards.

17. Start the three remaining engines simultaneously, using the technique described in para 10. As the fuel flow can only be monitored for one of the engines t!he HP cock levers for the other two must be moved forward carefully using the lever of the monitored engine as a guide. Each engine should start in the normal manner. Because of the number of indicators to be watched, .it is essential that both pilot:s monitor the .fire warning lights and indications of oil pressure, RPM and JPT. If any engine malfunctions, close its HP cock independently of the other engines but leave the engine master switch ON until the starting cycles of the other engines are complete. When all engines are running satisfactorily, the AEO switches on alternators in the normal manner.

Rapid Starting of Engines 18. General

a. 200 volt AC and 28 volt DC eleetrical supplies are required while starting the engines and are normally obtained from an external source. Elec­trical supplies may be obtained from the AAPP when starting engines individually but this source must not be used when starting all engines simultane­ously. b. The rapid-start compressed air installation, when fully charged, should provide sufficient air for three individual engine starts on each side. If the engines are started simultaneously, sufficient air should remain in the system for one individual start per side (minimum pressure 1100 PSI). Not more than two combuster starts on any one engine are to be attempted in any one 30-minute period. c. Owing to the risk of high JPT's and the difficulty of monitoring and adjusting four engines at the same time, simultaneous rapid starting must not be attempted at temperatures below minus 15°C. In such temperatures, No 1 or No 4 engines should be started by the combuster starter and the remaining engines started by using air crossbleed as described in paras 14 or 17.

19. Simultaneous Rapid Starting a. In order to gain full benefit from the rapid start installation, a complete combat readiness check should be carried out before engine starting. On completion of the combat readiness check, leave all systems selected as required for take-off. Carry out the checks in the Flight Reference Cards.

Note: The gyro hold-off button relay is only energised when RAPID is selected on the air selector switch and the master switch is OFF.

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b. Before starting the engines, at least one booster pump per group must be on. To start the engines, move all ,throttles to the 50% RPM position, se'lect the master switch ON and press the master RAPID START button. Combuster ignition is indicated by each individual starter button light coming on. Engine light-up is indicated by rising jet pipe temperature after approximately five seconds. As each engine accelerates, its starter disengages (indicated by the starter button lights going out) on overspeed-trip or on timer control after 12 seconds. During acceleration, check the indications of oil pressure, RPM, JPT and fire warning. The JPT should rise to between 400°C and 550°C in approximately 18 seconds and then begin to fall. When the JPT on any engine has stopped rising, wait a further two seconds and then close all throttles to the idling posj.rion; if ·the JPT on any engine reaches 600°C, all throtrles must be closed imme­diately to the idling position. In either event il:he throttles, having been closed, can be reset immedi­ately to give the required RPM.

c. During the start, the alternators come on line as engine RPM increase and, in addition, the flying control motors start automatically, the rudder first, followed by the ailerons and elevators together. All flying control motors should be started within 13 to 15 seconds of pressing the engine start button. The MFS gyros also start automatically as soon as the master start switch is selected ON.

20. Individual Rapid Starting

a. The rapid start installation permits individual starting of engines, provided that the air tempera­ture is above minus 26°C, without the automatic starting of PFC, etc. One or more engines may be started as required by using the individual start buttons. The engine starting indications and pro­cedures are detailed in para 19b. It must be remembered that, if the remaining engines are to be started using air crossbleed, an outboard engine should be started first.

b. The checks for a single rapid start are given in the Flight Reference Cards.

c. The remaining engines may be started in a similar manner. Alternatively, they may be started by using air crossbleed from the running engine. Select NORMAL on the air selector switch and start as described in paras 14 or 17.

Starting Failures-Normal Start 21. Starter Button Light Fails to Come On. If, during a normal start, the starter button light does not come on, cancel the start by switching the engine master switch OFF. Change the bulb in. the starter button, recheck all selections and try again.

22. Engine Fails to Light Up. If an engine fails to light up normally, close the HP cock and cancel the starting cycle by switching the master switch to OFF.

Note : If a normal start has failed, and it is necessary to use the combuster starter, it is essential to check that the master switch is OFF before selecting RAPID.

23. Starter Bution Light Remains On. If the starter button light remains on, it indicates that the air control valve has not closed. The following actions are to be taken ~ .. and completed by 22% RPM:

a. Select the master switch OFF, air selector switch NORMAL (light goes out).

b. Reselect the master switch ON (do not press starter button). If the light remains out the valve has closed ; the system is defective but safe and should be rectified after flight. If the light comes on again, the valve has not closed and the system is not safe ; shut down the engine.

Starting Failures-Rapid Start

24. Starter Combuster Fails to Ignite

Indication-No light in individual start button and bottled air supply cut off after 2 seconds.

Action- Investigate. Press the individua1l start button again, or crossfeed from another engine. Operationally, press the appro­priate individual start button imme­diately. If further failure occurs, crossfeed air from another engine.

25. Bulb Failure

Indication-No light in individual start button after 2 seconds but rapid start cycles normally.

Action- Do not interfere with the start. Check that the bottled air supply is cut off on overspeed or after 12 seconds. Record fault after landing.

26. Engine Fails to Ignite

Indication-No rise in JPT after approximately 5 seconds.

Action- Close the HP cock and put the master switch OFF. Investigate. Carry out two dry motoring cycles and attempt a further rapid start or crossfeed air from another engine. OperationaUy, leave the throttle set to the 50% RPM position and as soon as the light goes out press the start button again. If this attempt fails, crossfeed air from another engine.

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27. Engine Fails to Accelerate to Idling after Starter Disengages

Indication-No further increase in RPM or oil pressure.

Action- Close the HP cock. Wait until the engine bas stopped rotating. Reset the controls and try another rapid start. Operationally, close the HP cock and wait 7 seconds. Reset the throttle to the 50% RPM position and press the individual start button.

28. Excessive 1PT Indication-JPT reaches 700°C or the rate of tem­

perature rise indicates that this limit is likely to be exceeded.

Action- Close the HP cock immediately and monitor the fire warning 1ight. Investi­gate (ground crew). Operationally, if there is no fire or obvious damage indicated and it is essential to restart, wait at least 15 seconds for the engine to run down, reset the controls and press the individual start button.

Dry Motoring Cycle 29. a. If an engine fails to light up, after the HP

cock has been opened, a certain amount of fuel will have accumulated in the engine. Clear this by carrying out two dry motoring cycles before making any further attempt to start, otherwise high tempera­tures, or fire, with the risk of damage to the engine will result. The procedure for a dry motoring cycle is similar to that for a NORMAL start except that the ignition switch is OFF and the HP cock remains CLOSED.

b. Proceed as follows: 1. Air selector switch 2. Ignition switch 3. Throttle/HP cock 4. Engine master switch 5. Palouste

If crossfeeding : Engine RPM Appropriate engine air

switches 6. Clearance to motor 7. Individual start button

8. Master start switch

NORMAL OFF SHUT ON Ready

70%

OPEN

Pressed, 'felease when indicator light comes on OFF after 30 seconds and before 60 seconds from the time the start button was pressed. Check light out.

If crossfeeding : Throttle back -to idling.

Note 1 : Allow the motored engine to stop turning before starting the second motoring cycle or before a further attempt to start.

Note 2 : During the motoring cycle the oil pressure and RPM should register (approx 3% RPM).

After Starting All Engines (All Conditions) 30. Carry out the After-Starting checks in the Flight Reference Cards.

Taxying 31. Ensure that the parking brake is fully off before taxying.

32. As visibility from the cockpit is restricted, it is advisable to inspect the area before entering the air­craft, especially if it is intended to taxy in confined spaces. Take particular note of objects likely to be blown by the jet efflux.

33. Before taxying, the scanner must be st,abilised or secured and the relevant checks in the Flight Reference Oards car.cied out.

34. The thrust required to overcome the inertia of the aircraft and tyre set varies with the AUW and the surface but large amounts of thrust are rarely needed. Once the aircraft is in motion, sufficient thrust for normal taxying is obtained with all engines idling. At light weights, it is difficult to keep the speed down with all engines running. It is recommended, therefore, that on the completion of a sortie, the outboard engines are shut down to reduce the brake load.

35. As soon as convenient after moving, ·the brakes and nosewheel steering should be checked. To operate the nosewheel steering, press the selector button at the base of the control handle and offset the rudder by the required amount. Very little is achieved by using asymmetric thrust in turns. Differential braking however, can be used to assist in tight turns but care should be taken not to turn too tightly by this method, otherwise the steering and centring jacks may be dam­aged. It is possible to complete a 180° tum fairly comfortably on a 50 yard runway.

36. a. In emergency, it is possible to taxy using only differential braking to tum ; the pilot should always be ready to steer by this method should the nosewheel steering fail.

b. Do not operate the bomb doors while taxying, as the nosewheel steering becomes ineffective until their operation has ceased (approximately six seconds).

c. The brakes are very effective but it is possible to use them unevenly and thus to overheat one side by inadvertent differential braking, when taxying or during the landing run.

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Take-off 37. Complete the checks before take-off before enter­ing the runway. Align the aircraft with the runway and, with the brakes applied, open the throttles to 80% RPM, making sure that the brakes hold. Check for significant discrepancies between individual engine indications. When ·the engines are stabilised, switch on airframe anti-icing if required (30 seconds maximum before take-off). Ensure that the parking brake is off, release the brakes and then open up the tt:hrottles to full thrust. If the brakes are released suddenly, the nose tends to rise but it is unlikely that the nosewheel will leave the runway. If, in emergency, the JPT limiter is set to OFF, the RPM must be restricted during take-off (according to the table below) in order to avoid exceeding the JPT limits.

Olympus Olympus 301 engines 200 Series (but see Limitations)

Ambient %RPM Ambient %RPM Temp •c Limit Temp •c Limit

- 10 99 -Sand below 102

- 5 98·5 0 101· 5 0 98 5 101 5 97·5 10 101

10 97·5 15 100·5 15 97 20 100 20 96·5 25 99 ·5 25 96·5 30 99·5 30 96 35 99 35 96 40 98·5 40 95·5 45 98 45 95·5

38. There is no tendency to swing and any small deviations in the early stages can be corrected by nose­wheel steering. The rudder starts to become effective above 60 to 80 knots. Care must be taken when using either nosewheel steering or rudder to prevent over-controlling in the early part of the take-off run. Accelerallion is good, even at high weights, and is very rapid if full power is used at lighter weights (below 160,000 lb AUW). Above 100 knots, nosewheel steer­ing is almost ineffective unless weight is maintained on the nosewbee~ therefore the stick should be held well forward of the central position if nosewheel steering is used continuously throughout the take-off run (see Part 3 Ohap 4 for handling following engine failure on take-off).

39. At the rotation speed (see .table), ease the contra! column back so that the aircraft becomes airborne. (As weight is reduced below 170,000 lb, correspondingly

Jess backward movement of the control column is re­quired). Apply the brakes for 4 seconds and select undercarriage up; allow the aircraft to accelerate to ;the initial climb speed as the undercarriage is retracting and continue to accelerate to climbing speed.

Rotation and Initial Climb Speeds

Rotation speed Initial climb AUW (lb) (knots) speed (knots)

150,000 and below 135 148 160,000 139 148 165,000 141 149 170,000 143 151 180,000 148 156 190,000 153 160 195,000 155 163 200,000 157 165 204,000 162 170

After Take-off

40. Keep slip and skid to a mmunum while the undercarriage is travelling, in order to reduce stresses on the undercarriage door brackets. The under­carriage retracts in 9 to 10 seconds and no difficulty is experienced in achieving a clean aircraft before the undercarriage limiting speed of 270 knots is reached. Whenever possible, the undercarriage should be com­pletely retracted before exceeding 200 knots. There is no appreciable trim change during take-off or under­carriage retraction but, as speed increases, a steadily increasing push-force on the control column is neces­sary, because of the rapid increase in speed. This QUsh-force can be trimmed out easily in increments as the aircraft is accelerated to its climbing speed. Make a visual inspection of the undercarriage after UP selection, using the periscope.

41. At a safe height, throttle the engines to 93% and select CRUISE on the TAKE-OFF /CRUISE selector. Carry out the after take-off checks as soon as possible after take-off. Engine RPM creep in the climb, and 93% must be maintained by use of the throttles up to FL 300. Above FL 300 set and maintain 95% until top of climb is reached. 95% is the maximum per­mitted RPM for day-to-day operation in order to prolong engine life. Under operational conditions, or when specifically authorised, open the throttles fully and climb at maximum continuous power.

42. During take-off, yaw damper malfunction does not noticeably affect handling characteristics but, as speed increases to about 170 knots, the effects become apparent. Select the other channel immediately or, if both are defective, switch OFF.

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Aborted Take-Off Procedure

43. In all instances where the take-off run has to be aborted the following actions are to be taken:

a. Warn crew aborting. b. Close <the throttles. c. Select airbrakes to HIGH DRAG. d. Stream the tail brake parachute if speed is between 75 knots and 145 knots. e. Apply maximum continuous braking at or below normal maximum braking speed. f. Garry out engine firejfuilure drills as appropriate. g. Inform ATC "Aborting". h. Clear rUJilway if feasible. 1. Carry out brake fire drills before further taxying.

44. After an aborted take-off using NMBS it may be possible to clear the runway. In which case the aircraft should be raxied slowly with minimum application of toe brakes; the parking brake should not be used to hold the aircraft stationary during checks, or for park­ing.

WARNING: If maximum continuous braking is used above NMBS, the aircraft should be stopped and

evacuated as soon as practicable. Because of the high risk of tyres bursting the aircraft should not be ap­proached for at least 2 hours unless it is necessary to extinguish any fire.

Climbing 45. The recommended climb speed is 250 knots to 20,000 feet and then 300 knots up to a height where this speed coincides with 0 · 86M.

46. If the JPT limiters have been overridden because of unserviceability, the JPT must be watched very care­fully if the limits are not to be exceeded during the climb.

47. Between 10,000 feet and 15,000 feet, the pressu­risation failure warning hom may blow because the aircraft rate of climb is greater than the rate of increase of cabin pressure but, by 15,000 feet, the selected pressure should have been achieved. If cabin pressure surging occurs, check that the duct relief flap is in the

~ closed position; if surging persists, close one of the in-use engine air switches and leave closed until the top~ of climb.

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PART 3

CHAPTER 2-HANDLING IN FLIGHT

Contents

Para

Engine Handling 1 Engine RPM 2

JPT 5

Engine Idling Speeds 8

Combustion Chamber Failure ... 9

Gauge Failures 14

P1 P3 Switch Failure 16

General Flying ... 17 Trimming 22 Airbrake Characteristics 24

Engine Handling I. General. Engine life is affected by the frequency and extent of temperature changes caused by increasing ~md decreasing thrust ; therefore, to prolong engine life and maintain performance, throttle movement should be smooth and changes in thrust as few as possible. Slam decelerations to IDLING may be made

~at all altitudes. With 301 engines, care should be taken to ensure that the inboard throttles are not inadver­tently closed beyond the detent position when above 15,000 feet. Slam accelerations should be avoided._ but, in an emergency, can be made from any throttle setting at all altitudes. It is recommended that throttles should not be opened from IDLING to OPEN in less than 3 seconds.

Engine RPM 2. Thrust should not be reduced from TAKE-OFF by means of the TAKE-OFF /CRUISE selector switch alone. Before moving the switch to CRUISE, reduce the engine speed to 93% RPM. Then, after setting the switch <to CRUISE, maintain 93% up to FL300 then set 95% and maintain until top of climb is reached. If maximum cruise thrust is required, open the throttles fully.

3. Similarly, when increasing thrust from the CRUISE setting to take-off, close the throttles until the RPM begin to fall before selecting the CRUISE/ TAKE-OFF selector switch to TAKE-OFF. The throttles can then be opened fully (but see Engine Limitations, Part 2, Chapter 2).

4. The selector switch must be set to T AK.E-OFF (200 series only) before entering the circuit, to ensure that full power is available for overshoot.

Para

Bomb Doors 25

High Speed Flight 27

Pitch and Yaw Dampers Inoperative ... 29

Approach to the Stall ... 31

Stalling in Turns 34

Flight in Turbulent Air 35

Flight with the Entrance Door Open ... 36

Fuel Handling 38

Descending 39

JPT 5. The JPT •limiters must always be switched ON, unless they are proved to be defective. Each is capable of keeping -the JPT within ± 5°C of the selected limita­tion, when controlling, but regular checks should be made for excessive temperatures, especially when taking off in high ambient temperatures, climbing and changing power at high altitudes. The JPT limiters do not prevent excessive turbine temperature during rapidly changing engine conditions with CRUISE selected but, at T AK.E-OFF, their sup­pressed datum prevents overswing of turbine tempera­ture during acceleration at all ambient conditions.

6. Under varying ambient conditions and with the JPT limiters working properly, each engine does not necessarily indicate the same RPM and JPT as the others. An engine may reach its governed RPM before reaching the JPT limit or vice versa. However, to be within the limits in steady conditions with all engines at full throttle, no engine should be more than 30°C hotter than the mean of the others, or more than 2% RPM slower than the mean of the others.

7. Symptoms of Malfunction (at Governed Limits) a. If rthe JPT or RPM of any engine fall outside the limits of 30°C above or 2% RPM below the others or, alternatively, if the RPM and/or JPT are fluctuating continuously, either the engine is mal­functioning or its JPT limiter is unserviceable, or more air is being bled from this engine than the others.

b. The following is the recommended procedure to discover whether the engine is at fault:

(1) Check AC supplies to limiters (two engines affected).

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(2) Check ENGINE AIR switches all ON. (3) Check CABIN AIR switches both on (two engines affected). ( 4) Check wing and engine anti-icing switches as required (two engines affected).

c. If the fault is not cleared, select 93% RPM by manual use of the thrott!les.

d. Switch off the limiters; if after 2 minutes the engine is still outside the JPT limits, it must be assumed that the engine is unserviceable and it is advisable to shut it down, as it may have suffered damage to its compressor, combustion system or turbine. If, however, the engine RPM and JPT come within the limits with .the JPT limiters switched off, it is probable that the limiter is defective.

e. If, with the JPT limiters switched off, small fluctuations of RPM and JPT persist (within the limits stated in para 7(a)), accompanied by similar fluctuations in fuel flow to the affected engine, the presence of air in the fuel chassis is the probable cause. The engine may be kept nmning and the defect should be reported after landing. Air in the fuel chassis may also cause fluctuations at cruise power settings.

f. Having discovered the extent of the malfunction, the original engine settings may be restored and it is preferable, whenever possible, to have the limiters switched ON.

Engine Idling Speeds 8. Idling speed varies with altitude and forward speed. The characteristics of the 200 series engines may be different on the incorporation 'Of Mod 1785. They may result in undershoots of idling RPM after a de­celeration, or a downward idling reset following a period of idling conditions, both of the order of 5% LP RPM. Engine response from these conditions will be normal in that movement of the throttle will give inunediate engine acceleration, but the time to maximum RPM could be extended by as much as 1 second. The idling speeds under varying conditions are listed below:

% RPM

Condition 200 Series 301 Engines Engines

Static sea level idling 32 no load 27 oo load (ISA) 29 · 5 full load

on altema-tors 24·5

Approach idling 37 to 41 31 to 35

Idling at 50,000 feet 76 to 7S 7S to 81 (O·S6M)

Windmilling at 0 ·SSM 15 to 20 16 to 19

Combustion Chamber Failure

9. Any of the following symptoms could indicate combustion chamber failure:

a. A reduction or restriction of approximately 5% RPM at idling or at maximum power.

b. A change in JPT of approximately 30°C (usually an increase, but a decrease could occur).

c. Sluggish acceleration.

10. To detect a faHed combustion chamber, set the engines accurately to the same RPM and check the fuel flow rate. If the variation is 5 lb/min or more, record:

a. The RPM.

b. The four engine fuel flow rates.

c. The four engine JPT.

11. Do not check at RPM of more than 90% or when the throttles are fully closed. It is particularly important that RPM are settled and synchronised.

12. Record the flow rate variations by zeroing the lowest fuel flow rate and expressing the other three readings as 'plus' increments.

13. A variation of 8 lb/min or more on any engine can indicate the possibility of combustion chamber failure. Proceed as follows:

a. Check for possible gauge sticking by moving the throttle of the suspect engine and noting a sensible associated change in fuel flow rate.

b. If the flow rate system is not faulty repeat the recording required art para 10.

c. If the variation is still 8 lb/min or more, alter the flight conditions slightly to give a challge of engine RPM of at least 3%. Again repeat the recordings, being particularly careful that RPM are settled and synchronised.

d. If the variation is still 8 lb/min or more, a failed combustion chamber should be suspected and the engine should be flamed out.

Gauge Failures

14. If the RPM gauge of an engine fails, the engine must be shut down as it may have suffered damage to its auxiliary gear-box.

15. If either the oil pressure or JPT gauge of an engine fails, the engine should be shut down as it is no longer possible to monitor it fully. The engine may only be kept running for reasons of range or opera­tional necessity.

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PI P3 Switch Failure

16. If the Pl P3 switch fails (200 series engines only) the maximum RPM attainable above 15,000 to 20,000 feet will be approximately 80%. Carry out the following drill :

a. Throttle back another engine to the same RPM and compare .the fuel flows and JPT of the two engines. If the fuel flows and JPT are within 5 lb/ min and 30°C respectively, the engine may be kept running and full RPM may be attainable at a lower level.

b. If the engine is outside the above limits it must be assumed that the engine is unserviceable and it is advisable to shut it down.

General Flying

17. Flight with powered flying controls failed and flight without artificial feel is described in Part 1, Chapter 7.

18. The elevator stick force varies as the square of the speed but it is possible to overstress the aircraft during manoeuvres in the pitching plane at high air­speeds. At high mach numbers, the elevator stick

~forces become heavier. Longitudinal stability is notice­ably reduced with the CG at the aft limit. For low altitude training the CG should be kept forward of the mid-position to reduce the possibility of overstressing the aircraft in pitch. ~

19. The rudder forces are very heavy except in the low speed range when the use of rudder is necessary for correctly balanced turns.

20. The ailerons are very light and effective and some experience may be needed before a tendency to over­control at higher speeds can be avoided. T he stick force is constant over the speed range but the maxi­mum deflection is limited progressively from 150 knots onwards.

21. a . Coarse use of aileron produces large amounts of adverse yaw and, with the undercarriage down, the resulting sideslip can cause the load limits of the undercarriage doors to be exceeded.

27.

b. Adverse yaw is most marked at 200 knots and heavy, co-ordinated rudder application is required to counteract it. When changing the direction of roll, large angles of sideslip are produced and co­ordination is particularly difficult. During man­oeuvres, the slip indicator tends to under-read due to the heavy damping of the ball.

Trimming 22. Care is needed to trim the aircraft accurately. The best results are achieved if small increments of trim are applied and time is given for them to take effect before any further adjustments are made.

23. Change of Trim

Raising and lowering UC Negligible

Airbrakes (medium drag) Slight nose-down

Airbrakes (high drag) Further nose-down

Airbrakes (in) Nose-up

Bomb doors Negligible

Airbrake Characteristics

24. At high airspeeds at lower altitudes, the airbrakes are very effective and cause only mild buffet in · the HIGH D RAG position, with a marked nose-down change of trim. At higher altitudes, dose to the limiting mach number, the HIGH DRAG position produces marked buffet, accompanied by severe airframe vibra­tion, as well as the nose-down change of trim. If the limiting mach number is exceeded, tllis vibration makes cockpit instruments unreadable. At low airspeeds, ,the airbrakes are much less effective but do assist during the approach and landing. The airbrakes take approxi­mately 5 seconds to move from IN to MEDIUM DRAG and a further 2 seconds from MEDIUM DRAG to HIGH DRAG.

Bomb Doors

25. When bomb doors are opened at high airspeeds and mach numbers, moderate buffet occurs. Buffet is correspondingly less at lower airspeeds and mach numbers. With the bomb doors open, a slight increase in engine RPM is necessary to maintain a given air­speed in level flight.

26. If the aircraft is flown at alritude for more than 30 seconds with the bomb doors open, temperature effect on the control runs may result in gradual but noticeable changes in trim, particularly in the aileron sense.

High Speed Flight

WARNING : When flying at high indicated airspeeds or mach numbers, frequent comparisons must be made between the readings of the 1st pilot's, co-pilot's and navigator's ASI and between the pilots' machmeters.

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28. At High Level/High Mach Numbers a. Mach Trimmer Irn>perative. Unless specifically authorised, the speed should be restricted to a maximum of 0·9M if the mach trimmer is inopera­tive. The aircraft readily accelerates to maximum mach number in level flight, except at high AUW when a slight descent is necessary. Very little change of trim occurs from 0·85M to 0·88M but, beyond this, a nose-down trim change develops which requires a substantial pull-force at 0·93M. At higher mach numbers, the nose-down trim change increases rapidly and there may be very little control left for recovery above 0·95M. If mach number increases further, the nose drops even with the stick fully back. If control is lost, recovery is effected by closing the throttles (leaving the airbrakes IN) and holding the stick fully back until, with decreasing height and attendant drop in mach number, control is regained. The trimmer can be used, if necessary, to relieve the pull-force but do not relieve the elevator feel ; over­trimming must be avoided, otherwise there is a strong tightening of the pull-out as mach number is reduced. If the elevator feel does become inoperative, extreme care must be taken during the recovery phase to prevent serious overstressing of the airframe.

b. Mach Trimmer Operative. No difference in handling is noted from 0 · 85M until the mach trimmer comes into operation at about 0 · 87M. Beyond this point, a steadily increasing nose-up change of trim develops (due to the progressive up elevator applied by .the mach trimmer), which should not be trimmed out above 0·9M. The maximum permitted mach number is 0·93M (0·92M for 301 engines), at which speed a substantial push-force may be required, dependant on CG position. The mach trimmer still operates above 0 · 93M (0 · 92) until full extension of the servos occurs at about 0·95M at which speed the elevators have reached the almost fully up position. The aircraft now behaves, and recovery from above the limiting mach number is effected, in exactly the same manner as described in :the •last sub-paragraph. However, over­controlling is more likely to occur. This is due to the extension and retraction of the mach trimmer with changing speed on recovery.

c. Engine Handling. Disturbances in air intake flow during recovery from high speed runs may lead to engine surge or flame-out particularly with 301 engines. If surge occurs, partially close the HP cock; if surge symptoms persist, close the HP cock fully and take normal cold relight action. If flame-out occurs, attempt a hot relight ; if unsuccessful, close the HP cock and take cold relight action.

Pitch and Yaw Dampers Inoperative

29. Pitch Damper. The foss of one channel of the pitch damper is almost unnoticeable and no restriction

is imposed but speed must be restricted to 0 · 90M if more than one pitch damper channel becomes inoperative.

30. Yaw Dampers. The loss of the yaw damper is most noticeable when accurate course keeping is neces­sary or during an instrument approach. No limita­tion is imposed if the yaw dampers become inoperative.

Approach to the Stall

31. Stalling is not permitted. Speed must not be reduced below the pre-stall buffet and in any case not below the threshold speed for the weight less 5 knots.

32. a. As speed is reduced, with airbrakes in and depending on weight, pre-stall buffet may be experi­enced at the corresponding threshold speed.

b. At high angles of attack the rudder is masked by the mainplane. This results in a reduced rudder response and larger than normal rudder movement is required to maintain directional control.

c. As speed is further reduced, a rate of sink develops which increases rapidly and may exceed 4000 fit:jmin.

d. At 115 knots all controls are light and effective but rudder response is greatly reduced. In the land­ing configuration with approach power, any buffet is masked by the airbrakes.

33. At speeds below 115 knots at aft CG and with the stick fully back, directional instability may occur, causing considerable yawing and rolling. If this is allowed to develop, it is impossible to control until the stick is pushed forward. During recovery, the direction of yaw and roll reverse. Excessive height is lost before recovery is complete.

Stalling in Turns

34. It is not possible to stall this aircraft in turns unless either the g limitations are exceeded or airspeed is extremely low. However, at high altitudes the loading in the turn should not be increased after the onset of the initial buffet.

Flight in Turbulent Air

35. Flight in turbulent air should be avoided but, if this is not possible, the speed should be maintained between 180 and 300 knots, preferably at 220 knots.

Flight with the Entrance Door Open

36. The entrance door may safely be opened in flight at speeds up to 220 knots and higher in emergency.

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RESTRICI'ED AP IOIB-1902-15 Handling in Flight

Above 220 knots there is a danger of the forward bulkhead collapsing. Below 220 knots, the handling qualities are not materially affected. There is a slight nose-up change of trim.

37. The noise level is very high at all speeds and intercommunication is difficult ; therefore it is advis­able that the captain's orders and intentions are made clear before the door is opened.

Fuel Handling 38. Part 1, Chapter 8 ~ ~ lists the occasions when fuel checks are recommended. If flying for maximum range, reduce trim drag by maintaining the CG index between minus 2 · 0 and plus 2 · 0.

Descending 39. Cruise Descent. Descent using the airspeeds quoted in the Operating Data, with throttles fully closed, is theoretically the most economical method of losing height but the improvement in range is marginal and is easily lost through inaccuracies of navigation. The rate of descent, particularly at high altitudes, is low and the time to descend from high altitude is over an hour. With the throttle detents fitted (301 engines), the inboard engines flight idle at 5% RPM higher than the outboard engines at 50,000 feet. The detent can be manually overridden in emergency, eg rapid descent or relighting.

40. Normal Descent. Close the throttles, select MEDIUM DRAG airbrake and maintain a speed of 250 knots. With 301 engines and detents operative use HIGH DRAG airbrake down to 15,000 feet. These settings achieve a comfortable rate of descent at the correct speed for the standard let-down.

41. Rapid Descent. Before descending, warn the crew and have all loose articles stowed.

a. Using Airbrakes Only. Close the throttles, extend HIGH DRAG airbrake and dive the aircraft at 0 · 90M/300 knots (0 · 88M/300 knots if the auto­mach trim is unserviceable). Buffet is heavy.

b. Emet·gency Descent. In emergency, select HIGH DRAG airbrake and, provided that the speed is less tl1an 0 · 90M and 270 knots, ~ower the undercarriage and descend at 0 · 88M/260 knots to 40,000 feet. In this configuration, the time to descend from 56,000 feet to 40,000 feet is approximately 1 t minutes. Flight with the undercarriage down at speeds above 200 knots may cause some damage to the undercarriage door assemblies. It is therefore recommended that the undercarriage is not lowered during practice descents.

42. For icing in the descent see Part 3, Chapter 5, paragraph 9.

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PART 3

CHAPTER 3 - CIRCUIT AND LANDING PROCEDURES

Contents

Joining the Circuit Visual Approach Manual Let-Down and ILS Approach Low Level Manual ILS Procedure Holding Pattern QDRJQDM Procedure PARJGCA Approach Autopilot Let Down and ILS Approach Low Level Auto-ILS Procedure ... Landing Braking Emergencies During Landing

Para 1 2

8

Landing With CG Outside Recommended Range Crosswind Landing

15 1'6 17 18 19 20 21 26 33 35 38 39 40 41

Landing Without Using A.i.J:lbra:kes Overshooting Roller Landings

Joining the Circuit 1. Before joining the circuit, carry out the airfield

~ recovery checks. Pre-Landing checks should be~ completed on the downwind leg. The aircraft handles comfortably at threshold speed plus 30 knots (see Table) on the downwind leg. Allow 100 feet pressure error correction to all circuit heights when using the Mk 19F, Mk 29B and Mk 30A altimeters.

Visual Approach 2. While visual circuits are possible in conditions of poor visibility, the restricted view from the cockpit, particularly during the line-up phase, does not lend itself to tills procedure ; whenever possible, make an instrument approach in these circumstances.

3. Select the required compass on the MFS selector and set the runway QDM at the top datum on the selected compass. All changes of heading can be made by setting the index to the required course, leaving the setting knob out between selections.

4. MEDIUM DRAG airbrake is normally selected when leaving the downwind position and HIGH DRAG before crossing the runway threshdld.

5. a. Make the final turn at pattern speed (threshold +30 knots), adjusting the speed to approach plus 10 knots by the mid-point of the tum; aim to acrueve approach speed when lined up with the runway. Cross the threshold with power on at the correct speed for the AUW.

b. From considerations of directional control, the minimum recommended approach speed is 135 knots. However, when landing at weights of 110,000 lb and below, the threshald speed may be reduced to 120 knots to avoid excessive float. c. At AUW above 150,000 Ib, the approach speed is 15 knots rugher than the threshold speed.

Circuit Speeds

Pattern Approach Threshold Speed Speed Speed

AUW (lb) (knots) (knots) (knots)

120,000 and below 155 135 125 130,000 160 140 130 140,000 165 145 135 150,000 169 149 139 160,000 173 158 143 170,000 177 162 147 180,000 181 166 151 190,000 185 170 155 200,000 189 174 159 210,000 193 178 163

6. The moderate nose-up attitude in the latter stages of the approach does not affect forward visibility, which is good. Control response is good and speed main­tenance on the approach is not difficult. However, the rate of descent must always be controlled by use of the throttles. The use of elevators alone to reduce the rate of descent can induce a rugh rate of sink, wruch can only be checked by a large increase in thrust.

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7. a. At high AUW, directional contral is poor in the approach configuration, unless sideslip is kept to a minimum by careful co-ordination of rudder and aileron. Limit angles of bank to a maximum of 15° during the final approach.

b. If, as a result of poor control co-ordination, large side-slip angles are induced, discrepancies can occur between the 1st pilot's and the co-pilot's ASI.

Manual Let-Down and ILS Approach Note: The following MFS selections are recom­mended until experience is gained: ; mereafter either or both compasses may be set for me final approach at ooy oonvenient stage of t!he pl:'()<CCf({ure at the pilo~· discretion.

8. On approaching the descent point make the following selections:

a. Set the navigation selector central.

b. Adjust the compass card to give the required magnenic heading (index at itop datum).

c. Select HEIGHT (if desired).

d. Check both radio 'Sense switches are pointing upwards.

c. Compass selector to non-flying pilot.

9. At clle descent point:

a. The non-flying pilot selects the dosoent heading to top datum.

b. The flying pilot follows up on his rompass and oommences descent (Wrbrakes <to MEDIUM DRAG, throttles closed and speed to 250 knots IAS).

c. Flst datum is relected when .the dosoent attitude is steady (cancels HEIGHT).

10. During the descent the flying .pilot sets the localiser QDM to top datum, pulls out the compass setting knob and thereafter sets dle index to ·the required aircraft beading.

11. At the inbound tum, or when cleared Ito join the localiser as appropriate, the non-flying pilot:

a. Rotates me compass cmd (in two stages if necessary to bring localiser QDM to top datum) and rthen pulls out his rompass setting knob.

b. Sets the navigation selector to LOC.

c. Offsets the headmg index t'O allow for drift (if necassary) once the aircraft is steady inbound.

12. At check height:

a. Seloot HEIGIIT (if desired).

b. Select LOC & GP.

c. Oheck that drift is correctly allowed for.

d. Seleot airbrakes IN and set power to give pattem speed.

13. When approaching the gl!idepat'h. select the air­brakes to MEDIUM DRAG and adjust power and speed as required. At glidepatlh intercept relect DATUM: check that the pitch selector reverts to APPROACH when DATUM is released. (If the heading error is more than 5°, or if the ADP are showing a tum demand, the pitch selector must be moved back to central.)

14. On overshool'ing the flying pilot sets the engines to 80% or 93%, as appropriate, and selects the required attitude. The non-flying pilot:

a. Selects airbrakes IN when ordered. b. Selects the navigation selector central. c. Selects a osen:sible !heading with the lindex (usually aligns the index with clle pointer). d. Seleot!S fast datum when :requested.

Low Level ManualiLS Procedure 15. It is recommended that, at low level, a race-traCk pa11tem is flown. At me beginning of the downwind leg carry out the pte-'landing checks and adjust power ro maintain pattern speed. Make the foNowing MFS selections :

a. Set the localiser QDM to top datum on both beam compasses.

b. <lleck both radio sense switches are po.int'ing upwards.

c. Set the heading .indices ro the downwind heading (including an allowance for drift).

d. Select LOC & GP.

e. Select HEIGHT (if desired).

f. When at a suitable diS11ance downwind initiate a tum towards clle beam by rotating me !heading ·index to localiser QDM plus drift allowance (in ·two stages if necessary). g. Wihen -steady on me beam adjust the drift allowance if necessary and reduce speed to approach speed.

h. Continue as in para 13.

Holding Pattem 16. Use normal MFS selection procedures, as in para 8, until the !holding pattern inbound heading is at top datum. Then pull out 'the compass setting knob and rotate me index to demand aM subsequent Jtums until olea:red ro leave the pattern.

QDRJQDM Procedure 17. a. On rlle approach ro intercept the beam:

(1) Set !the Iocaliser QDM to top datum. (2) Set ·the heading index to the required heading. (3) Select LOC. ( 4) Point the radio sense swirob upwards. (5) ·Seleot HEIGIIT (if desired).

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b. When crossing th~ localiser tberun (as shown by beam bar movement), tum the heading index to 15° past QDR. As the beam is re-entered, set the index to QDR and point the radio sense switches across.

c. Select LOC & GP.

d. If a steady offset is established, adjust the head­ing index for drift.

e. At the designated position outbound, rotate the heading index to the required heiading <to initiate the procedure tum {if !the index is still within me radio coupled range marks, point the radio sense switches upwards). On leaving the beam, commence timing (45 seconds plus wind allowance).

f. After the appropriate time, initiate a tum towards the beam by rotating the heading index to localiser QDM plus drift allowance (in two stages).

g. When steady on the beam, adjust the drift wUowance if necessary and reduce speed to approoch <Speed.

h. Continue as in para 13.

PAR/GCA Approach 18. Initially rotate the rompass ca:rd to ·the heading given by the GCA controller with the heading index vertical. When the aircraft is on the centre line the QDM of the runway is set at top datum and the compass setting knob is pulled out. New courses are then set with the heading index (see also Note before para 8).

Autopilot Let-Down and ILS Approach

Note: Trimming during autopilot approaches requires close attention.

19. In addition to the MFS selections listed in para 8 to 14, make the following autopilot selections:

a. To reach the descent point, engage the auto­pilot and select TRACK and ALT. When the airfield recovery checks have been completed, reselect TRACK.

b. Initiate the descent by using fast and slow rates of dive as required ':in ·conjunction with otosing the tllb.rott'les and seleCting the airbrakes to MEDIUM DRAG. When the speed is at 250 knots select IAS.

c. Re-select TRACK afit:er LOC is selected on the MFS.

d. Level out using ~ast and slow rates of dimb as required. Select AL T when level at rhe required height. Select HEIGHT on the MFS if desired. Select airbrakes IN and set power to give the pattern speed. Check that drift is correctly aHowed for on the beam compass. Ensure that the speed is below 180 knots and select LOC & GP on the MFS.

e. At glidepath intercept, continue as in para 13 but select GLIDE on the autopilot before selecting DATUM on the MFS.

f. At decision height the flying pilot uses his instinctive cut-out; the non-flying pi1ot checks the indications of disengagement.

Low Level Auto ILS Procedure

Note: Trimming during autopilot approaches requires close attention.

20. In addition to the MFS selections listed in para 15 and 17, make the following autopilot selections:

a. Race T-rack Pattern (1) When established on the downwind leg, engage the autopilot and select TRACK (provided the speed is 'below 180 knots) and ALT.

(2) At glidepath intercept, continue as in para 13 but select GLIDE on the autopilot before selecting DATUM on the MFS.

(3) Att decision height the flying pilot uses his instinctive cut-{)ut ; the non-flying pilot checks the indications of disengagement.

b. QDRJQDM Procedure (1) To intercept the beam, engage the autopilot and ·sdec:t TRACK and ALT.

(2) When established outbound, select LOC & GP; provided the speed is below 180 knots, reseleot TRACK.

(3) At glidepath intercept, continue as in para 13 but select GLIDE on the autopilot before selecting DATUM on the MFS.

( 4) At decision height the flying pilot uses his instinctive cut-out; tthe non-flying pilot checks the indications of disengagement.

Landing

21. The maximum AUW for landing is given in Part 2. If it is necessary to land at an AUW greater than 140,000 lb, a runway of 9000 feet or more should be used. The tail brake parachute (TBC) may be streamed at 135 knots (145 knots maximum) and should be jettisoned between 50 and 60 knots. In emergency, it may be retained until the aircraft has stopped and then removed by ground servicing personnel. However, if it is essential that the TBC be removed before the aircraft has been finally shut down the inboard engines may be opened up to approximately 40% and jettison selected, or the TBC jettisoned while the aircraft is taxying (if appropriate). Jettisoning the TBC with the aircraft stationary and shut down is permissible when absolutely necessary but some minor damage to •the rear fuselage may result.

22. a. Fly the circuit and approach at the speeds recommended for the weight. A safe margin for contro1 of the aircraft is allowed with up tto 30° of bank angle at pattern speeds and 20° bank angle at approach speeds. (15° at approach speed a~ve

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195,000 lb). During the later stages of the approach, (approx 200 feet), but not before decision height on an instrument approach, HIGH DRAG airbrake may be selected and speed reduced so as to cross the threshold with power on at the recommended speed. Maintain the correct approach speed by careful use of the throttles. At 195,000 lb and above, the rate of descent at touch-down must be kept to a minimum.

b. When the mean surface wind strength is 15 knots or above, .ir:relipective of direction, the threshold speed should be increased by oneJthlrd of the mean wind strength. If the threshold speed then exceeds the approach speed, the latter should be increased to equal the former.

23. Normal Landing. Aerodynamic braking may be used at all weights. After touch-down, when both bogies are firmly on the ground, raise the nose pro­gressively as speed is reduced, until the control column is fully back. Because of the small ground clearance at the wing tips and the high angles of incidence asso­ciated with aerodynamic braking, any mishandling in the lateral sense may result in damage to the wing tips. Bank angles in excess of 3!0 are significant in this respect. Aerodynamic braking must not be continued below 85 knots if the headwind is greater than 25 knots, since there is a possibility of the tail being scraped. Two red lights, on the roaming in front of the 1st pilot, come on when the tail of the aircraft is roo close to the runway. With headwind oomponents of less t!han 25 knots, aerodynamic braking may be oontinued down to 80 knots. Lower the nosewheel onto the runway and apply the brakes as reqUJired.

24. Overweight Landings. At weight!S below 140,000 lb, norma'l approach and landing techniques should be employed. At weights above 140,000 lb the speed at touch-down may be higher t!han the normal maximum braking speed (NMBS) and the maximum permissible speed for streaming the TBC. In such cases hold the nose as high as possible after touch-down and use aerodynamic braking until the speed falls to 145 knots. At 145 knots stream the TBC and when it deploys, lower the nose. Start braking when the nosewheel is on the runway and the speed is at NMBS. Provided that the AEO has been warned, the outboard engines may be shut down. Under favourable conditions of runway length, headwind or AUW, the pilot may, at his discretion, delay streaming the TBC until 135 knots. (Any parachute streamed above 135 knots may be scrapped.)

25. Short Landi1zg. Cross the runway threshold at the lowest safe height and at the calculated threshold speed. Provided that the speed is below 145 knots, stream the TBC as soon as the main wheels are on the runway. Use aerodynamic braking until the TBC has deve­loped and the speed .is 5 knots above NMBS, then

lower the nose. When the nosewheel is on the runway and the speed is at NMBS apply maximum continuous braking.

Braking

26. Both front and rear wheels of the mainwheel units must be firmly on the ground before the wheel­brakes are applied, as the maxaret units do not operate until the wheels are rotating. As a safeguard against locking tlhe wheels during a bounce after Iandtng, the ma.x:aret units ,remain inoperative for a few seconds. Apply brake pressure smoothly and progressively when the speed is below the braking speed for the weight (see Operating Data). On dry surfaces, the. maxaret units normally prevent the wheels from lockmg when excessive brake pressure is applied but, unless the shortest possible landing run is required, more geode use of the brakes is recommended.

27. On wet surfaces, braking action may be severely reduced according to the degree of wetness of the surface. Use light intermitJtent brake application initially. As tSpCed is reduced, brake application may be progressively increased and held continuously. If s?p or skid is suspected, release the brakes momentarily and then re-apply them gradually.

28. A drastic reduction in braking action must be expected on flooded or icy surfaces; whenever possible, avoid mese oonditions. If a landing has to be made on a

~flooded or icy runway, the TBC is to be used, cross­wind conditions permitting. If the use of the brake parachute is not possible, aim for a firm touch-down using aerodynamic braking down to at least 90 knots. Braking action must not be taken until the nosewheel is firmly on the ground; care should be exercised during braking to avoid wheel locking. ~

29. To prevent possible damage, it is necessary to limit the speed at which continuous braking is applied. This speed is referred to as the Normal Maxunum Braking Speed (NMBS) (see ODM, Section 10). It is defined as l'he maximum airspeed from which maxi­mum continuous braking can be applied and the air­craft brought to rest without damage to the brakes. However, heavy braking will markedly reduce brake effectiveness as the heat absorption limit of the brakes is approached; even moderate -braking at light weight and slow speed can have the same effect if prolonged, eg lengthy taxying using brakes against power. Follow­~ any such cases of excessive braking, sufficient cooling ·time should be allowed to resture the brakes to full capacity.

WARNING: If maximum continuous braking is used from NMBS or speeds approaching NMBS, to avoid subsequent damage to the brakes the aircraft should not normally be brought completely to rest, The air-

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Circuit and Landing Procedures

~ craf·t should be taxied slowly to dispersal with the mini­mum use of brakes and parked without applying the parking brake. If it is necessary to ·stop before reaching t!he dispersal, the minimum application of roe brakes should be used; the parking hra'ke should not be used.~

30. The emergency maximum braking speed (EMBS), defined in the Take-off Section of the Operating Data Manual, does not apply to landing. Tests have shown that the use of EMBS does not give a corresponding reduction in the landing run.

31. If the brake parachute fails to deploy, the appropriate NMBS must be used.

32. Brake Cooling Times (Training Case Below 140,000 lb AUW). If an intermediate landing is made

~during a sortie, the following conditions are to be ~ observed:

a. On the intermediate landing, brakes must not be applied above 80 !-'!lots.

b. Great care must be taken during taxying to minimise any further heating of the brakes.

c. A minimum of 15 minutes is to elapse between the end of the landing roll and the start of the take­off run.

d. On the subsequent take-off, the undercarriage must not 'be retracted for a minimum period of 5 minutes after unstick.

e. If the subsequent take-off is aborted, the TBC must be streamed but the brakes should not be applied until the speed has fallen to the NMBS without parachute figure. If practicable brake appli­cation should be delayed until the speed is below 80 knots.

Emergencies During Landing 33. Landing With Hydraulic Failure If, foHowing a hydraulic failure, nhe braking capacity available is reduced nhe following technique should be used:

a. Reduce AUW to near the minimum permitted.

b. Select a runway that is into Wtind and is at least 9000 feet long, preferably with a smooth over-run area. c. Make a normal approadb. aiming to cross the nmway threshold at the correct speed for the AUW.

d. If the pressure in the brake accumulators falls to 3000 PSI delay using the EHPP until just before touchdown in order to ·keep fluid loss to a minimum.

e. After touchdown stream the TBC once the speed has fallen to 145 knots.

f. With some brake pressure remaining, lower the nose and commence braking taking care Ito avoid maxare~g.

g. During t:he landing run attempt to recharge me brake accumulators if the pressure falls to 3000 PSI.

h. Wilth no brake pressure available considemtion can the given t:o maintaining the aerodynam!i.c braking attitude after streaming the TBC. However, rudder effectiveness in this attitude is markedly reduced and, with any crosswind component, better initial direc­tional oontrol will be available by lowering the nose immediately.

i. Provided the AEO has been warned the outboard engines may be closed down after the parachute has deployed.

j. Stop on the runway and apply the parking brake. If it is necessary to taxy off the runway for opera­tional reasons, taxy slowly so that only one application of the brakes is needed to stop the aircraft.

34. Failure of Parachute to Stream. If the pilot decides to roll, following failure of the parachute to stream, the switch must be put to JETTISON imme­diately and the ABO must observe through the peri­scope, or A TC confirm, that the parachute has jettisoned, before power is applied.

~Landing With CG Outside Recommended Range 35. The recommended CG index range for landing is + 2 to -2 landing gear down. At low fuel s·tates it may not be possible to mainmin the CG within these limits. At any fuel state, if the non-essential loads are not reset after load shedding, the CG will move rapidly aft if the No 1 and 2 tank boostrer pumps are feeding. An aft movement of about 5 units of index can be ex­perienced during a normal low level instrument circuit. The tail proximity warning lights are inoperative with non-essential loads shed.

36. Landing With Aft CG. Landing with the CG at rhe aft limit presents no undue difficulty. On the ap­proach to land, sufficient elevator control is 'available and speed stability is not markedly affected. When the CG index is aft of +4, it is recommended that the threshold speed is increased by 10 knots and that roe aircrtaft is flown gently onto the runway rather than flared. Aerodynamic braking should not be attempted but the TBC may be used as required.

37. Landing With Forward CG. When the CG index is forward of minus 4, it is recommended !!hat the approach to land is slightly shallower than normal, and that the threshold speed is increased by 10 knots to improve elevon effectiveness for the flare. After touch­down it may not be possible to maintain aerodynamic braking attitude down to 80 knots. ~

Crosswind Landing 38. A crosswind landing, using the crab technique, presents no special difficulty in crosswind components up to the limitation of 20 knots. When yawing the

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aircraft into line winh the runway prior to touchdown, there is a marked tendency for the into-wind wing to rise ; this tendency may be countered by prompt application of aileron. Aerodynamic braking may be used within the limiu imposed by the tendency of the into-wind wing to rise. Lower the nose to allow braking below NMBS. The braking parachute may be used after the nosewheel has been lowered ; any swing which develops due to the use of the braking parachute is best controlled by using nosewheel steering and brakes. If any difficulty is experienced in maintaining oontrol, jettison the parachute. Avoid landing on very wet or icy runways in strong crosswinds ; if the cross-­wind component in these conditions exceeds 7 knots, it is inadvisable to stream the parachute.

Landing Without Using Airbrakes

39. When landing without airbrakes, use the normal procedure but a longer approach is advisable. To avoid high sink rates developing if the engines are throttled back to the slow response range, any neces­sary increase in power must be anticipated.

Overshooting

40. Overshooting from any height presents no diffi­culties. Engine response is rapid on the approach but is comparatively slow (approximately 5 seconds to maximum RPM in normal temperatures, up to 13 seconds in tropical temperatures) when going round

again from the runway. Open the throttles as neces­sary and climb away. At a safe height, if leaving the circuit, complete the overshoot checks. Under normal conditions, an overshoot from ground level followed by a visual circuit and landing requires 1200 to 1500 lb of fuel, and an overshoot from ground level fol­lowed by a low level instrument approach and landing requi,res 2000 to 2500 lb of fuel. At low AUW, the aircraft accelerates rapidly if full power is applied on overshoot. To avoid an extremely steep climb-away and to prolong engine life, it is recommended that power is restricted to 80% RPM.

Roller Landings

41. When making a roller landing, hold the nose­wheel close to the runway. Retract the airbrakes and open the throttles smoothly to a minimum of 80% RPM, being prepared for some difference in response from each engine. Avoid any tendency to overcontrol on the rudder. During acceleration, avoid a high nose-up attitude and any tendency to take off below the rotation speed (135 knots up to 150,000 lb). See also Engine Limitations.

42. When making a roller landing after an asymmetric approach, lower the nosewheel onto the runway. Before the throttles are opened for take-off they must all be in the idling position; it is essential that RPM on all engines are equal. ~eli::J g ~-kr ScJ.se...s. -& 43. i;h ~ wnov,t,d~ ~ ~ o:.~~ G..MY ~~~ /)A-~ a:J-~ 2..

•

v~ ~ c~I,,P~~ ~.

b~ ~J ~.<> I.e~ T~"'id #~~Oo ~.,d ~ J -. ( /rNA 4.

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PART 3

CHAPTER 4 - ASYMMETRIC FLYING

Contents

Engine Failure During Take-Off Shutting Down or Failure of an Engine in Flight Relighting in Flight

Para 1 3 6

General Asymmetric Landing and Overshoot

Engine Failure During Take-Off 1. Single Engine Failure. Single engine failure on take--off presents no handling problems. However, in cases of mechanical failure it may result in double failure and should be treated as such.

2. Double Engine Failure a. If the failure occurs before the decision speed is reached, abandon the take-off, using the appro­priate FRC drill.

b. If failure occurs above the decision speed and the take-off is continued, maintain directional control, using rudder or a combination of rudder and nosewheel steering until rotation speed is reached. Rotate the aircraft, aiming to reach the initial climb speed at 50 feet. Apply rudder to keep the aircraft straight and, on becoming airborne, apply up to 5° of bank towards the live engines, accepting a slip indication of half a ball's width towards the live engines ; the use of coarse aileron must be avoided, as the adverse yaw induced affects directional control. At rotation, be prepared for the aircraft to move sideways towards the dead engines.

Note: Minimum control speed on the ground cv~!CG) is 135 knots (145), using rudder alone for directional control, but control can be maintained from 115 knots (125) using nosewheel steering. (Speeds quoted are

~for 200 series engines at take-off power, or 301 engines at cruise power, speeds in brackets are for~ 301 engines at rake-off power.)

c. For optimum climb gradient after take-off, raise the l·anding gear and climlb straight ahead at •the initial climb speed, using 2° of bank towards the live engines. Any reduction of speed below initial climb speed results in a marked penalty in climb performance. At a safe height, accelerate to the pattern speed and carry out the appropriate FRC drill. An improved rate of climb is achieved with a further increase of speed. It is recommended that the aircraft is climbed straight ahead to 1000 feet to provide a satisfactory height for rear crew escape should the need arise. Ensure that the pattern speed is achieved before turning on to the downwind leg, restricting bank to not more than 25<0 at AUW above 150,000 lb and 15° at AUW above 195,000 lb.

12 15

Shutting Down or Failure of an Engine in Flight 3. If an engine fails or is being shutdown in flight, close the throttle to the HP cock SHUT position, close the appropriate ENGINE AIR switch, adjust the relevant booster pumps and crossfeed cocks as necessary and inform the AEO. T,he LP cock should be closed if the situation demands it.

4. If an engine fails in flight when it is unlikely to have suffered, and is not showing symptoms of, mechanical damage, and is windmilling at a steady and reasonable speed, an attempt may be made to relight it. During relighting, keep a careful watch on the engine indicators and, at any sign of malfunc­tion or difficulty, shut the HP cock immediately; do not relight.

5. a. The buried wing root installation of the Olympus engines makes the adjacent engine of a pair liable to mechanica+l damage should one engine suffer structural failure. To minimise damage to the airframe, containment shields are fitted outboard of the outer engines to protect the wing, inboard of the inner engines to protect the bomb bay. They are not fitted between each engine and captains must always be aware of the possibility of damage occur­ring to the adjacent engine whenever fire or failure indications for one engine are received.

b. The following points should be considered when­ever engine fire or failures occur in flight.

(1) Fire/Failure Indications for Both Engines. Both engines should be closed down and the appropriate fire/failure drills carried out.

(2) Fire/Failure Indications for One Engine. The failed engine should be closed down and the appropriate fire/failure drills ·carried out. Con­sideration should then be given to the following:

(a) If accompanied by obvious signs of engine structural failure ( eg marked vibration, explosion or ruptured aircraft skin) and circumstances permit, the adjacent engine of the pair should be shut down as a precaution.

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(b) If there are no obvious signs of structural failure, the other engine should be left at cruise power and monitored carefully for symptoms of malfunction. It is emphasised that there is no benefit to be gained in this situation from throttling the adjacent engine to flight idle.

Relighting in Flight 6. If two engines fail simultaneously, relight the out­board first.

1. Relighting is progressively more certain with reduction in altitude and should always be possible at heights below 35,000 feet, and at airspeeds of 0·9M or less if practicable (preferably below 200 knots). 'Hot relights' may be achieved at any height, provided that relight action is taken within a few seconds of flame­out; hot relights are not recommended if the cause of flame-out is not known, except in the case of a multi­engine flame-out.

8. a. When attempting a 'hot relight', immediately on noticing that the engine has flamed out, press the relight button and progressively close the throttle until a successful relight is achieved. This is indicated by a steady rise of both JPT and RPM.

b. Excessive play in tl1e t hrottle/HP cock control runs may result in engine RPM stabilising at approximately 40% and JPT at 200 to 250°C, with the throttle lever at the idling gate. Provided that no other abnormal symptoms are present, the throttle may be opened slightly to assist acceleration. How­ever, if the RPM or JPT fluctuate or any other unusual symptoms occur, the HP cock must be shut immediately.

c. If a rapid rise in JPT occurs without an accom­panying rise in RPM, a state of over-fuelling may exist. The throttle must be closed, if necessary to behind the idling gate, and the relight button released. (If the JPT rises to 600°C the throttle must be closed immediately to the HP cock SHUT position.) When the JPT falls, again press the relight button and slowly open the throttle until a successful relight is achieved.

d. If a successful relight is not achieved within 30 seconds, the throttle must be closed to the HP cock SHUT position and the procedure for relighting a cold engine carried out. (See Flight Reference Cards.)

9. a. If, during a cold relight, a rapid rise in JPT occurs, without an accompanying rise in RPM, follow the procedure in para 8c.

b. If no relight is obtained after 30 seconds, release the button and SHUT the HP cock. A further attempt can be made at a lower altitude, after allow­ing the engine a minimum of 3 minutes to drain out.

c. Do not make more than three attempts to relight any one engine in the same sortie.

d. With Olympus 301 engines, severe buffeting can occur with an inboard engine windmilling. Relight­ing under such circumstances can be impossible. Buffeting can be reduced by decreasing airspeed and/or reducing RPM on the adjacent outboard engine. e. Relighting is facilitated by throttling back to idling RPM the adjacent engine of the pair.

10. After the engine has been relit, inform the AEO.

11. Whenever an engine is relit in flight, a JPT /RPM comparison check between the engines must be carried out as detailed in Part 3, Chapter 2, paras 6 and 7.

General 12. In asymmetric flight conditions, above 200 knots, there is little difficulty in controlling the aircraft, even with two engines failed on the same side. At and below 200 knots, care is required to co-ordinate aileron and rudder movements to reduce the effects of adverse yaw caused by aileron drag and to maintain accurate control of the aircraft.

13. If asymmetric, with bomb bay loads of 21,000 lb and above, only gentle manoeuvres are Ito be carried out; rapid and/or full movement of the controls is not to be made.

14. When flying with one inboard engine windmilling, some airframe buffet may be experienced at speeds above 0 ·85M. In aircraft powered by Olympus Mk 200 series engines, ·this buffet is very pronounced at around 40,000 feet and enough to shake the instrument panels vigorously as speed is increased above 0·91M. At altitudes around 50,000 feet the buffet is less marked and changes little with change of mach number. Aircraft powered by Olympus Mk 301 engines are affected in a similar manner, although me :buffet is even more pronounced and becomes most severe if the adjacent outboard engine is set at high power. When flying at 0·85M at 30,000 feet, the vibration becomes excessive if the RPM of the outboard engine are increased to 88 % , while at 48,000 feetj0·85M severe vibration is felt with this engine set at 100% RPM. As height is increased above 50,000 feet, the vibration becomes much less marked and at all altitudes a setting of 80% in the outboard engines reduces the vibration to negligible proportions.

Asymmetric Landing and Overshoot 15. a. One Engine Inoperative. The technique anCl

procedure used for approach using three engines are the same as for a normal approach, except that the power settings axe increased by approximately 3%. The visual committal height (VCH) for a three engine approach is 150 feet.

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b. Two Engines on One Side Inoperative. The minimum approach speed during a two engine asymmetric approach down to VCH must not be below the calculated approach speed for the weight or 145 knots, whichever is higher. High drag airbrake should not be selected tmtil a decision to land has been made at the VCH of 200 feet above touch-down

16. If the thrust available is marginal, it is recom­mended that a landing without airbrakes is carried out.

17. When an asymmetric overshoot is made it must be initiated at or above the VCH and the approach

~ speed. The following procedure will allow an asym­metric overshoot on two engines to be made at any AUW within the WAT limits:

a. Level the wings

b. Increase the power required, counteracting yaw with rudder.

c. Airbrakes IN

d. Select landing gear up if required. To avoid tmdue stress on the landing gear door assemblies, only raise clle landing gear if any of the following apply: ~

(1) Rear crew must escape (2) At weights above 140,000 lb (3) Leaving the circuit ( 4) It is essential for safe control (5) The AAPP operation must be considered

e. Allow the aircraft to accelerate to pattern speed before climbing away. Rudder must be applied to counteract the yaw and up to 5° of bank may be applied towards the live engines; coarse use of aileron must be avoided.

~ At normal landing weights, the power should be in­'Creased initially to 93% RPM (full power can be used if required). ~

WARNING : The throttles must be opened carefully so that increase in power is simultaneous with throttle movement. Too rapid throttle movement which ent'!lils a lag in power response must be avoided.

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PART 3

CHAPTER 5- OPERATING IN ICING CONDITIONS

Contents

On rthe Ground Climb Level FHght Descent

Note: The ice protection systems are of the anti­icing type and should therefore be switched on before icing conditions are encountered. See also Part 1, Chapter 11.

On the Ground 1. Engine Anti-Icing. Engine anti-icing must be used on the ground in conditions of cold damp weather when the temperature is +3°C or below, and the humidity is more than 90%, or the visibility is less than 1000 metres. It must be switched on after the engines have been started and switched off after take­off when clear of icing conditions. 2. With the system in use the following actions are required:

a. Olympus Mk 301. When holding, increase RPM momentarily rto maximum continuous every 3 minutes. No special action is required during ·taxying. b. Olympus Mk 200 Series. When holding, increase RPM to maximum continuous for 1 minute every 10 minutes. No special action is required during taxying but RPM is to be increased to maximum continuous for 1 minute at the end of the taxying period, if the taxying period exceeds 10 minutes. The engines should be simi1arly opened up imme­diately prior to commencing the take-off run.

3. Airframe Anti-Icing. The airframe anti-icing system must not be used while ground-running or taxying. If required, it may be switched on imme­diately before starting the ·take-off run when the engines have stabilised at 80%. A momentary pause should be made to ensure the 'temperatures are rising before the brakes are released. 4. Windscreen Heating. Select the windscreen heat­ing to LOW before starting and to MEDIUM before take-off. (To prevent cracking, HIGH is to be selected when ambient temperature is + 30°C or higher,

~ reverting to MEDIUM at altitude). If a windscreen overheat is indicated, isolate the affected windscreen. After take-off it may be switched on again. ~

Climb 5. During a climb with the anti-icing in operation, to minimise the risk of range losses from ice accretion

Para 1 5 6 9

to the rear of the heated leading edges, it is recom­mended that speed be reduced to 250 knots, without reducing the engine RPM, until the aircraft is clear of the icing layer.

Level Flight 6. Engine Anti-Icing

a. In icing conditions, engine RPM should exceed the following:

200 Series 301

85% above 10,000 feet At all altitudes:

78% in temperatures below 68% in temperatures below minus 12•c below 10,000 minus 14•c feet

68% in temperatures above 60% in temperatures above minus 12•c below 10,000 minus 14•c feet

200 series engines run at lower speeds must be opened up to maximum continuous RPM for 1 minute after every 10 minutes, while 301 engines run at lower speeds must be opened up momentarily to maximum continuous RPM after every 3 minutes. If the engines are n01: opened up to cle1lr them of ice after the required periods, large pie~ of ice may break away to enter the engine with consequent risk of flame-out.

b. In order to obtain maximum engine protection, it is necessary to have both the airframe and engine systems on.

c. The 301 engines are particularly prone to ice damage and rthe 'anti-iicing must therefore always be on before and during flight in cloud.

7. Airframe Anti-Icing a. The ECM air-intake system should be selected ON whenever the wing anti-icing system is in use.

b. During level flight below 10,000 feet, at any speed, ·the airframe system provides adequate pro­tection if the outside air temperature does not fall below an indicated temperature of 0°C.

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c. If the system is switched on at high altitudes the wing !temperatures are unlikely to read more than approximately 80° to 100°C. At low altitudes, pro­viding engine RPM exceed 75%, the temperatureS should stabilise within the range of 130° to 150°C although there may be minor variations with changing flight conditions. With engine speeds below 75% the hot air supply may be insufficient to maintain the required temperature and if necessary, power should be increased. If the temperatures still can­not be maintained in AUTO, select MANUAL and INC until 140°C is reached but, assuming a fully serviceable system, AUTO should give a higher temperature than MANUAL as the cold air valve is opened fully when this selection is made.

d. The wing anti-icing gauges must be monitored at least every 15 minutes and if either temperature exceeds 170°C, the affected side must be switched off (except as in para 7e). If this does not reduce the temperature, select MANUAL and DEC. If this fails shut one or both ENGINE AIR switches on that side, ensuring that air is still available for cabin conditioning. The single controlling fuse for both the HA V and the CA V should also be checked.

e. With any part of the anti-icing system maifunc­tioning, flight in icing conditions is to be avoided. If this is not possible, the system may be kept on until icing conditions have been cleared, provided that the temperature can be reduced to 170°C and that temperatures are monitored continuously.

f. Wi.th two engines failed on the same side, some degree of airframe ice protection may be obtained by setting the engine stan master switch ON, in

order to open the air crossfeed, and selecting AUTO on the failed side.

g. With the system switched off, the co-pilot must monitor the temperature gauges to detect any un­acceptable hot air leaks. Indications will depend on flight conditions; in very low ambient temperatures providing the system has not recently been in use, any reading above 0°C indicates a leak. In wanner conditions a rise of 30°C above ambient denotes a leak. If a leak does become apparent the system should be switched on and controlled manually.

8. Windscreen Heating. If MEDIUM heat is not enough, switch to HIGH. If, subsequently, the overheat warning is given, isolate the affected windscreen.

Descent 9. When icing conditions are forecast for the descent, select and maintain 70% power on all engines and descend at 250 knots using HIGH DRAG airbrake, with all anti-icing systems on. If required, an increase in the rate of descent above flight level 150 can be achieved by increasing the speed to 330 knots. At the bottom of the descent, if severe icing is expected, hold­off should be made above or below the icing layer, as necessary, depending on the height of the layer. It must be emphasised that 70% RPM is a compromise and may not ensure full engine and airframe protection under all conditions.

10. After a descent through icing conditions the air­brakes may accumulate significant amounts of ice. If an accumulation of ice is suspected the airbrakes must not be selected from HIGH DRAG to MEDIUM DRAG or MEDIUM DRAG to IN as ice may damage the aircraft structure.

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AP lOIB-1902-15

PART 4

EMERGENCY PROCEDURES

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PART 4

CHAPTER I-EMERGENCIES

Contents

Crash Landing Para

1

Landing with Undercarriage in Abnormal Positions

Abandoning the Aircraft ...

3

5

Ditching

Aborted Take-Off I>.rocedures

Emergency Evacuation on the Ground

Engine Failure Above Decision Speed

Index of Emergency Drills and Malfunctioning Procedures

10

12

13

14

15

Crash Landing

1. The following considerations are recommended if a crash landing becomes necessary :

a. Before Landing (1) Reduce weight as much as is practicable.

(2) Have the nav /radar make the ejection seats safe. The pip pin for the canopy jettison gun must not be removed.

(3) If crash landing on an airfield, request foam on the runway as early as possible.

(4) If, in the opinion of the captain, there may be a danger of the navigators and AEO being trapped in the aircraft after the landing, they should be ordered to abandon the aircraft. Ensure that the undercarriage is in the up position.

(5) A check should be made for obstructions, bearing in mind the direction of expected swing, ie one mainwheel up condition. (6) If crash vehicles are available check their positions. (7) If possible, check that a ~ong ladder is available to expedite the crew's escape. (8) Ensure that all unnecessary navigational and electrical equipment is switched off. (9) Uncover the rear cabin windows prior to crash landing.

b. Approach (1) Make a normal approach with the under­carriage up or down according to circumstances. The advantages of reducing impact load with the undercarriage down, however, should be carefully considered.

(2) Jettison the canopy and close the HP cocks just before touchdown.

2. The recommended Crash Landing Drill is given in the Flight Reference Cards.

Landing with Undercarriage in Abnormal Positions

3. General a. If, after using the emergency system, only one leg is lowered, it is recommended that the aircraft is abandoned. In other cases, if a landing is oon­sidered feasible, then the general principle is that all crew stay with the aircraft. Teclmiques for landing ar-e given in para 4.

' field equipped with foam-laying apparatus. en landing with one main unit unl~ke the foam strip should be laid along_the· -slc:1e of the runway ,AN A~ that the wing ti eered to strike. The foam acts as ncant and so delays the start of the

loop, which imposes a heavy strain upon the

c. The possibility of major damage is also reduced if, after touchdown, the unsupported wing or nose is lowered at a controlled rate while the flying oon­trols are still effective, rather than be allowed to drop on to the runway.

4. Landing Techniques a. Belly Landing. If, after the use of the emer­gency system, all units of the undercarriage remain retracted, it is recommended that the aircraft be belly-landed, as follows:

(1) Reduce weight as much as practicable and switch off all unnecessary equipment. (2) Have the nav jradar make the ejection seats safe. The pip-pin for the canopy jettison gun must not be removed. (3) Ensure that the bomb doors and entrance door are closed. ( 4) Eooure all Joose objects are stowed, that all crew have t!heir harnesses tig'ht and locked, p~ rective helmens on, with 100% and emer-gency oxygen selected and flowing correctly.

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(5) Make a normal approach. ( 6) Jettison the canopy while still on the approach. (7) Make a normal landing, keeping the wings level and the rate of descent to a minimum. (8) Close the HP cocks. As soon as possible, operate the fire extinguishers and switch off all electrics.

b. Nosewheel Up, Both Mainwheels Down (1) Move the CG as far aft as possible, within permitted limits of the available fuel.

(2) Carry out a low overshoot, to check wind conditions.

(3) Ensure all loose dbjects are stowed, mat all crew have their harnesses tighit and locked, pro­tective helmets on, with 100% and emeTgency oxygen selected and flowing correctly.

( 4) Carry out ill normal circuit. Open vhe entrance door. Jettison the canopy on completion of final turn. Switch on the landing lamps at night.

(5) Touch down normally at the correct speed.

( 6) When firmly on the main wheels, stream the tail parachute (crosswind permitting) and cut the outboard engines.

(7) AEO to switch off Nos 1 and 4 alternators.

(8) Hold the nose up until speed drops to 80 knots, runway length permitting.

(9) While devaror control is still available, lower the nose on to the ground.

(10) As soon as the nose touches, cut the remain­ing engines; the co-pilot switches off all • ~ ·LP cocks; the AEO switches OFF all alternators and operates the battery isolating switch.

(11) When the nose is ·firmly on me ground, apply the brakes gently and evenly.

(12) When the aircraft stops, tihe co-pilot 'leaves first, followed by the nav jradllr, AEO, nav /plotter and 1st pilot in ·that order. If 6th and 7th seat mentbers are carried they should leave the air­craft after the co-pilot in the order 6th, then 7th.

c. Nosewheel Down, One Mainwheel Down (1) Move the CG as far aft and as far away from the failed mainwheel as possible. (2) Carry out at least one low overshoot, to check wind conditions. (3) Ensure all loose objects .are stowed, that all crew have their !harnesses tight and 1ocked, pro­tective helmets on, with 100% and emergency oxygen selected and flowing ·correotily. ( 4) Carry out a normal circuit and jettison the canopy on completion of the final tum. Switch on the landing lamps at night.

(5) Touch down normally at the correct speed. (6) On touchdown, cut the outer engines and lower the nosewheel onto the ground. (7) AEO switches off Nos 1 and 4 alternators. (8) Hold clre wing up, using aileron, rudder and nosewheel steering. (9) Before control effectiveness is lost, l.ower the wing and cut the remaining engines ; hold the aircraft straight for as long as possible. The co­pilot switches off the • ~ LP cocks; the AEO switches off aU the alternators and operates the battery isolating switch. (10) When the aircraft 'Stops, the co-pi:Iot leaves first followed by the nav /radar, AEO, nav /plotter, and 1st pilot in that order. If 6th and 7th seat members are carried ·they should leave the aircraft after the co-pilot in the order 6th, '!!ben 7th.

d. Nosewheel Down, Both Mainwheels Up. In these circumstances, it is recommended that the air­craft be abandoned, as it is considered that the hazards for the rear crew escaping past the nose­wheel are less than the danger to the whole crew of the nose section breaking off and the main fuse­lage overrunning the cabin. If, for any reason, a landing is imperative, the following technique is recommended:

(1) Reduce weight to the minimum practicable. (2) Insert the pins in the ejection seats but not in the canopy jettison gun. Switch off all unnecessary equipment. Ensure aU loose objects are stowed, that a1l crew have their harness'es tight and locked, protective helmets on, wivh 100% and emergency oxygen selected and :flowing correotly. (3) Make a normal approach. Jettison the canopy at the end of the final tum. Switch on the landing lamps at night. ( 4) Make a normal landing, keeping the wings level ; avoid a high nose-up attitude and land with minimum drift. Do not stream the brake para­chute. (5) As soon as the nosewheel drops to the ground, cut all engines and switch off all services.

Abandoning the Aircraft

5. General a. Ejections may be initiated in straight and level flight, at any height from ground level upwards. However, runway ejections should only be made when the speed of the aircraft is above 90 knots. At low altitude the aircraft should be straight and level or climbing; any significant rate of descent or nose-up antitude at the instant of ejection reduces the seat performance.

b. Rear crew members can leave the aircraft down to a minimum height of 250 feet at a maximum speed of 250 knots. Whenever possible, speed

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should be reduced to 200 knots and the aircraft be in a shallow climb with undercarriage raised, prior to escape.

6. To minimise the possibility of1 injury; by air blast or by loss of equipment, it is recommended that, if circumstances permit, speed is reduced as much as possible before attempting to escape. Escape is made easier if no personal survival pack is worn.

7. The following amp'lifies the drills in the Flight Reference Cards:

a. Escape for Rear Crew Members-General (1) Whenever possible, altitude should be reduced to below 40,000 feet. Above 40,000 feet the air­craft should only be abandoned in an extreme emea:­gency. Crew members should initiate the demand emergency oxygen set and then disconnect from

~ the aircraft system. Post-Mod 2393 a combined static line, oxygen hose and micjtel release coupling is fitted for rear crew members. Speed~ should be reduced as much as possible. It is most important that the exit is made by sliding cleanly down the door, in a bunched-up attitude. Ground tests also show that, by using the technique described in para 7 c, rear crew members can clear an extended nose-wheel on escape at speeds up to 180 knots.

(2) Move the abandon aircraft switch rearwards to the EMERGENCY position and confirm on the

~ intercom. Post-SEM 027, operation of the abandon aircraft switch automatically illuminates the cabin light. ~

(3) Navjplotter and AEO operate the door opening switches. Above 20,000 feet, ·the normal door opening lever should only be operated in addition to and simultaneously with the switches when escape in the minimum time is essential. Below 20,000 feet, the switches and door opening lever may both be used. Whichever method has been used to open the door, the first rear crew member to reach the door should ensure the door opeiung lever is in the gated EMERGENCY position. Ensure that static lines are connected. ( 4) When giving the order to abandon the air­craft, the pilot should normally indicate to the rear crew members that the static lines are to be used. However, below 1000 feet and 200 knots be should order the manual overrides to be used. (5) The nav ;radar's last action before sliding down the door must be to ensure that his oxygen hose passes behind his PSP. (6) Navigators and AEO leave the aircraft in ;the order, navjradar, AEO and navjplotter. If an experienced 6th seat crew member is carried he will be first to leave the aircraft. In the case of an inexperienced 6th member he will leave after

the nav /radar. When an inexperienced 7th mem­ber is carried, the 6th member must be an experienced aircrew member or a crew chief. The 7th member is not to be given any task other man leave the aircroft when instructed. The order of abandoning will be 6th, 7th, nav /radar, AEO and nav /plotter. The co-pilot, if possible, should watch the rear crew members leave the aircraft and inform the 1st pilot when the nav /plotter has left. The crew gone lights indicate to the pilot that the rear crew members have left the aircraft.

b. Escape for Rear Crew Members-Undercarriage Raised

(1) Sit on the floor at the front end of the door aperture facing aft.

(2) Grasping the handle at the bottom of me centre seat, swing fOl'WW'd onto the door and slide down it. At speeds above 200 knots it is advisable to adopt and hold a bunched-up attitude to minimise the possibility of injury from limb flailing. Below 200 knots an extended attitude with the legs straight out and rigid probably gives an easier exit. An upward pull with the arms is necessary to ensure that the PSP is lifted clear of the door edge.

c. Escape for Rear Crew Members-Undercarriage Lowered. If the undercarriage cannot be raised, the following technique is recommended:

(1) Grasping the handle at the bottom of the centre seat, swing the legs onto the door facing aft. Slide down the door with the legs apart until the feet can be braced against the door­operating jacks. An upward pull with the arms is necessary to ensure that the PSP is lifted clear of the door edge. (2) Releasing the grip, lean forward with bent knees and grasp the right-hand (port) jack with •both hands, as low as possible, thumbs upper­most, right hand on top.

(3) Withdrawing both feet inwards from the jacks, keeping the knees bent and the body close to the port jack, swing down and round the port jack and over the port side at the bottom of the door. Release the hold on the jack as the body swings completely dear. Try to maintain a com­pact position with the arms dose to the body after letting go. Keeping close to the port jack decreases the risk of the PSP fouling the star­board jack.

d. Escape for Rear Crew Members at Low Altitudes

Note: Whenever possible, convert speed to height. If it is necessary to abandon the aircraft at very low altitude (below 1000 feet), reduction of the time interval between the moment at which the order to

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abandon aircraft is given and the moment at which the parachute deploys can be of overriding import­ance and the following points should be borne in mind:

(1) The time taken to open the door can be reduced to a minimum by operating either rear crew switch immediately then, if necessary, operating the manual door control. (2) The static line arms a barostat, which then withdraws the parachute pack pins after a delay of 2 seconds. Therefore, whether a static line is connected or not, the parachute release handle should be pulled as soon as possible after clearing the door.

8. The pilots should escape, using their ejection seats, after the rear crew members have escaped.

9. If the ejection seat automatic system fails after ejection, proceed as follows:

a. When forward speed is sufficiently low, discard the face screen. b. Pull the manual separation lever outwards and then up. c. Fall clear of the seat and pull the rip-cord handle.

Ditching 10. Model tests indicate that the ditching qualities of the aircraft are good there being no tendency to nose

~ under after impact. a. The following considerations and actions amplify the drill given in the FRC:

(1) Assessment of Sea S tate. Whenever possible, fly low over the water and study its surface before ditching. It is important to establish correctly the direction of the swell and of the surface wind. (2) Direction of Approach. The aircraft should always be ditched into wind if the surface of the water is smooth or there is a very long swell. However, ditching into the swell or large waves should be avoided because of the danger of nosing under. In practice a direction of approach which is a compromise between swell, wave and wind direction, may be the best choice. (3) Judgment of Height. As judgment of height over \Vater can be difficult, the Alt 7 or Alt 6 should be used if possible. The landing lamps should also be used at night. (4) Fuel Weight. Fuel weight should be reduced as much as practicable prior to ditching. Excess fuel may be used to position the aircraft in a more favourable location, eg closer to ships or land, but it is essential that the ditching is carried out while engine power is still available.

b. Ditching Drill (1) Ensure all loose objects are stowed, that all crew have their harnesses locked and tight, pro-

tective helmets on. 100% and emergency oxygen selected and flowing correctly. Uncover the rear cabin windows. (2) Have ejection seat pins replaced. (3) Disconnect PSP and lanyards, leg restraint, emergency oxygen and parachute harnesses as appropriate to crew position. (4) On the approach, stow the fuel console and jettison the canopy.

c. Touchdown (1) Touchdown should be made in a tail-down attitude at the lowest practicable speed commen­surate with the minimum rate of descent. Touching down at high speed and low angle of attack should be avoided due to the likelihood of the aircraft bouncing and the probable collapse of the bomb­aimer's blister with subsequent flooding of the cabin. In any event the control column should be held hard back after impact. ~

11. The recommended Ditching Drill is given in the Flight Reference Cards.

Aborted Take-Off Procedures

12. a. If an emergency occurs before decision speed, take-off is to be aborted in accordance wi~h the FRC drill. The following emergencies constitute mandatory reasons for abandon!ing take-off, unless otherwise authorised.

(1) Engine failure. (2) Any fire warning light coming on. (3) Double alternator failure. ( 4) PFC failure (main warning not accompanied by a white reminder MI).

b. The Captain is to warn the crew "Aborting".

c. The pilot flying the aircraft is to : (1) Close the throttles. (2) Select HIGH DRAG airbrake. (3) Apply maximum continuous braking at NMBS or below.

d. The non-flying pilot is to: (1) Stream the tail braking parachute (75 knots to 145 knots). (2) Carry out engine failure/fire drills as ordered.

e. The AEO calls "Aborting, Aborting" on the frequency in use.

f. The navfplotter calls airspeeds down to 50 knots.

Emergency Evacuation on the Ground

13. The following considerations for evacuating the aircraft in an emergency on the ground amplify the drills given in the Flight Reference Cards:

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a. Undercarriage Position. The direction of exit depends on the emergency and whether the under­carriage is raised or has collapsed during the emergency. If the nosewheel has collapsed, it may be possible to leave the aircraft through the door ; if, however, all the undercarriage legs are retracted, exit will have to be via the canopy aperture. The route from the canopy to leave the aircraft will depend upon the condition of the aircraft and whether a fire exists.

b. Canopy Jet.tison. It will be necessary to leave via the c':lnopy aperture if exit through the door is not possible. When the canopy is jettisoned, if the aircraft is stationary, the possibility exists that the canopy will fall back on to the cockpit and may injure one or other of the pilots.

c. Ejection Seat Pins. If exit through the entrance door is feasible, the pilots should replace the seat pan firing handle safety pin prior to leaving their seats. If the exit has to be made via the canopy aperture, and time permits, the main gun sear pins are to be inserted to make the ejection seats safe.

d. Crew Ladder. If speed of exit is essential and exit through the door is feasible the crew should slide down the door and clear the vicinity of the aircraft as quickly as possible. Only replace the door ladder when the degree of emergency allows.

e. Battery. If conditions permit the AEO should switch off the aircraft battery prior to evacuating the aircraft. It should be borne in mind that with the battery off all cabin lighting is lost. However, if the battery was left on, when it is safe and if it is feasible, the AEO should return to the aircraft to switch off the battery in order to make the aircraft electrically safe.

Engine Failure Above Decision Speed 14. a. If engine failure or other serious emergency

occurs above decision speed, the take-off is normally to be continued and the drill recommended in Part 3, Chapter 5 para 1 and 2 followed.

b. When the aircraft is safely airborne, the pilot flying the! aircraft is to close the HP cock(s) of the affected engine(s), simultaneously ordering the non· flying pilot to select undercarriage up. The flying pilot is then to order the non-flying pilot to carry out the drill 'l'equired, eg "Engine Failure Drill, No 3 Engine".

c. The non-flying pilot is to complete, from memory, the engine fire/failure drill as ordered: the immediate actions listed in the FRC as far as 'Fuel Pumps' are to be completed.

d. The AEO declares an emergency on the fre­quency in use.

e. Once the aircraft is fully under control (pattern speed attained) the FRC checks for engine fai1uref fire are to be completed.

f. A tum on to the downwind leg is not to be initiated below 1000 feet above ground level, and until pattern speed is attained. Bank is to be restricted to a maximum of 25° in the tum. It is recom­mended that an instrument palltem is flown following double engine failure.

g. When the aircraft is established on the downwind leg, the 'Resetting' checks are to be carried out (if required) followed by the 'Pre-landing' checks. The undercarriage is not to be lowered until approaching the glide path.

Index of Emergency Drills and Malfunctioning Procedures 15. In addition to the emergency drills and malfunc­tioning procedures contained in the Flight Reference Cards, Table 1 shows where, in the Aircrew Manual, the subject is covered.

Table 1. Index of Emergency Drills and Malfunctioning Procedures

Malfunctioning or Emergency Procedure Part Chapter Paragraph

Autopilot: Malfunctioning 1 12 84 to 87 Use after engine failure 1 12 82

Brake parachute: Unselected streaming 1 7 81

Canopy: 1 13 9 Jettisoning 1 9 47

Descent: Rapid descent 3 2 41

Electrical system: Alternator failures ... 1 4 32 to 37 Vital, essential and non-essential loads 1 4 Table 4

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' Malfunctioning or Emergency ProcedurB Part Chapter Paragraph

Emergency exits: Entrance door and static lines 1 9 43

Entrance door: Flight with entrance door open 3 2 36

Engines: Failure to start 3 1 21 to 28 Failure on take-off 3 41 1 Failure in Bight 3 41 3 Relighting 3 41 6

Fire protection systems: Operation of engine fire extinguishers 1 6 4 Operation of AAPP fire extinguisher 1 6 7 Operation of fuel tank fire extinguishers 1 6 8

Flying controls: PFC failure 1 7 17 Artificial feel failure 1 7 37 Trim failures ... 1 7 46 Yaw damper failures 1 7 53 Pitch damper failures 1 7 58 Automach trimmer failures 1 7 63 Airbrake failures 1 7 76

Fuel system: Suspect booster pump ... 1 8 78

~ Non-return valve failure l 8 83~ Sequence timer failure 1 8 ~84 Booster pump running continuously at full speed 1 8 85 Leaking tanks ... 1 8 86 Leaking fuel line 1 8 87 Amplifier faults 1 8 88 Effect of electrical load shedding on fuel system 1 8 89~

Landing emergencies : Landing with hydraulic failure 3 3 33 Brake parachute failure 3 3 31 Undercarriage emergency operation ... 1 10 37 Asymmetric landing and overshoot 3 4 15 to 17

Military Bight system : MFS failures 1 12 38, 48 to 53

Oxygen: Failure indications 1 2 49 Regulator failure 1 2 50

Pressurisation: CA U overspeed 1 1 27 Loss of cabin pressure 1 1 28 Pressurisation failure, oxygen drills 1 2 48 Emergency decompression 1 1 29

TFR malfunctions 1 15 32

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AP lOlB-1902-15

PART 5

ILLUSTRATIONS

List of Illustrations

Fig.

COCKPIT, GENERAL VIEW ... 1 - 1st PILOT'S INSTRUMENT PANEL ... 2 v

CENTRE INSTRUMENT PANEL 3

CO-PILOT'S INSTRUMENT PANEL 4

PORT CONSOLE (6P) ... 5

RETRACTABLE CONSOLE (SP) 6

STARBOARD CONSOLE (7P) ... 7

FUEL CONTENTS PANEL (2P) AND THROTTLE QUADRANT 8

ALTERNATOR CONTROL PANEL (lOP) 9

SECONDARY SUPPLIES PANEL (SOP) 10

AAPP CONTROL PANEL (70P) ... 11

RADIO SUPPLY PANEL (12P) 12

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5-l Fig. 1 Cockpit, General View

Nosewheel steering engage button 10 E2B compass

2 Elevator and aileron feel relief switch 11 Tl'R head-up display

3 Aileron and elevator trim switch 12 Aileron and elevator trim switch

4 Press-to-transmit swi rch 13 P ress-to-transmit switch

5 TFR head-up display 14. E levator and aileron feel relief switch

6 Tail clearance warning lights 15 Undercarriage emergency lowering control

7 E2B compass 16 Airb r:akes selector switch

8 RAT release handle 17 Pilot's bomb release control

9 Engine fire warning lights and extinguisher buttons (four) 18 Brakes toe buttons (four)

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AP 101B-1902-15 Illustrations

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5-1 Fig. 2 1st Pilot's Instrument Panel

Auto-throttle airsoeed detector indicator (if fitted) 10 ILS marker light

2 MFS annunciator (port compass) 11 Mk. 7B radio altimeter

3 Oxygen flow indicator 12 Director horizon

4 White floodlighting switch l3 Standby artificial horizon and slip indicator

5 Windscreen wiper switch 14 Beam compass

6 Auto-throttle comparator lights (inoperative if fitted) 15 Anti-dazzle lamp switch

7 TFR video light 16 Rudder pedal adjustment control

8 TFR warning light 17 Bombing indicator

9 TFR failure light 18 Crew escape lights (four)

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AP 101B-1902-15 Illustrations

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28

5-l Fig. 3 Centre Instrument Panel

Hydraulic triple pressure gauge 15 Bomb doors (or tin) position indicator

2 Hydraulic power pack switch 16 Canopy unlocked indicaror

3 Oil pressure gauges (four) 17 Entrance door unlocked indicator

4 RPM gauges (four) 18 Port pressure-head heater indicatOr

5 Engine governor control indicator (inoperative) 19 Main warning light

6 JPT gauges (four) 20 Control smfaces position indicator

7 Tail parachute swirch

8 Accelerometer

9 Main warning light

10 PFC warning indicator

11 Artificial feel indicator

12 Auto-stabiliser indicator

13 Airbrakes position indicator

14 Alternator failure warning light

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21 MFS selector

22 Auto-pilot trim indicator

23 Fuel pressure indicators (four)

24 Tacan indicator

25 Undercarriage control burtons

26 CG check bur.ron

27 CG indicator

28 Undercarriage position indicator

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5

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24 23 22 21

5-l Fig. 4 Co-pilot's Instrument Panel

ADF indicator 13 Windscreen wiper switch

2 Beam compass 14 Windscreen demist switch

3 Director horizon 15 Windscreen de-ice switch

4 TFR fail light 16 Fuel flow indicator (individual engines)

5 Radio altimeter limit lights 17 Oxygen flow indicator

6 ILS marker light 18 White flood lighting switch

7 Bomb bay tanks fire warning light 19 Flowmeter reset switch

8 Wing/fuselage tank fire warning light 20 Co-pilot's station box

9 TFR warning light 21 MFS annunciator (starboard compass)

10 MFS/TFR switch 22 Flowmeter total flow indicaror

11 'fFR video light 23 Co-pilot's slip indicator

12 Windscreen overheat warning light 24 Rudder pedal adjustment control

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ll

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37 36 35 34 33 32 3.1 30 29 28 27

5- l Fig. 5 Port Console (6P)

Air ventilated suits flow control 20 Store safety lock and warning lights

Air ventilated suits temperature control 21 Stowage for manual release key

Ration heater 22 Bomb doors normal control

Start master switch 23 Bomb doors emergency control

Rapid/normal start selector switch 24 Bomb jettison switch

Air crossfeed indicator 25 U/V lighting dimmer switch

Ignition switch 26 Oxygen regulator

Rapid start pushbutton 27 Radio altimeter controller

Gyro hold-off pushbutton 28 Artificial feel lock switch and indicawr light

Individual start pushbuttons (four) 29 Artificial feel warning cancel buttons

RT1 tone switch 30 Mach trimmer reset (two) and master control buttons

Audio warning test button 31 Pitch damper control buttons (four)

Canopy locking indicator 32 PFC stop buttons (ten)

Canopy jettison lever 33 Comparator reset button

TFR controller 34 Yaw damper selector switch

ILS/Tacan/ ADF audio switch 35 PFC and artificial feel start buttons (three)

RT2 tone switch 36 1st pilot's station box

Aerial changeover switch 37 RT2 controller

Abandon aircraft and emergency decompression switches

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AP 101B-1902-15 Illustrations

5-1 Page 5

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13 14 IS 16 17 18 19 20

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42 4 1 40 39 38 37 36 35

5-l Fig. 6 Retractable Console (5P)

1 Bomb bay fuel pressure indicator 2 Bomb bay tanks pressurisation switch (inoperative) 3 Bomb bay /main tank selector switch 4 Crossfeed cock switches (three) 5 Crossfeed cock position indica tors (three) 6 No. 6 tank (port) pump switch and contents pushbutton 7 No. 7 tank (port) pump switch and contents pushbutton 8 No. 4 tank (port) pump switch and contents pushbutton 9 No. 5 tank (port) pump switch and contents pushbutton

10 No. 3 tank (port) pump switch and contents pushbutton 11 Port auto/manual switches (two) 12 No. 2 tank (port) pump switch and contents pushbutton 13 Port fuel CG/transfer switch 14 Auto-pilot pitch control 15 Auto-pilot turn control 16 No. 1 tank (port) pump switch and contents pushbutton 17 Flight refuelling lateral CG control switch 18 Emergency trim control 19 No. 1 tank (stbd) pump switch and contents pushbutton 20 Starboard fuel CG/transfer switch 21 No. 2 tank (stbd) pump switch and contents pushbutton

22 Starboard auto/manual switches (two) 23 No. 3 tank (stbd) pump switch and contents pushbutton 24 No. 5 tank (stbd) pump switch and contents pushbutton 25 No. 4 tank (stbd) pump switch and contents pushbutton 26 No. 7 tank (stbd) pump switch and contents pushbutton 27 No. 6 tank (stbd) pump switch and contents pushbutton 28 Fuel flow engine selector pushbuttons (four) 29 Bomb bay/ main rank selector switch 30 Bomb bay fuel pressure indicator 31 Pump switches (two) for forward or A bomb bay tank 32 Pump switches (two) for rear orE bomb bay tank 33 Auto-pilot bomb switch 34 Auto-pilot lAS/altitude lock switch 35 Auto-pilot engage switch 36 Auto-pilot in indicator 37 Auto-land prime switch (inoperative) 38 Auto-pilot ready indicator 39 Channel engage switches (three) 40 Auto-pilot power switch 41 Auto-pilot glide switch 42 Auto-pilor track switch

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4

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2

Cabin temperature selector

2 Temperature control valve position indicator

3 Oxygen regulator

4 T hrottle detent isolation switch (if firred)

5 Cabin pressure selector

6 Engine air switches (four)

7 Cabin air switches (two)

8 AAPP air bleed indicator

9 AAPP cabin air bleed (only operative with Mk. 10101)

10 Canopy jettison lever

11 Canopy locking indicator

12 Probe lighting dimmer switches

13 Nitrogen purge switch

14 Tank pressurisation switch

15 Airframe ami-icing auto-manual switches (three)

16 Tank pressurisation indicators

17 Anti-icing temperature gauges (three)

18 Engine anti-icing switches (two)

19 Ration heater 20 AVS master switch

5-1 Fig. 7 Starboard Console (7P)

21 Windscreen overheat indicators

22 AVS temperature control

23 Anti-icing manual heat control switches (three)

24 AVS flow control

25 ECM monitor I alarm control

AP 101B-1902-15 Illustrations

26 Navigation lights/anti-collision lights steady / flash switch

27 Air-to-air refuelling master switch

28 Starboard landing/taxi lamp switch

29 Pon landing/taxi lamp switch

30 Air-to-air refuelling indicator

31 Identification light steady /morse switch

32 External lights master switch

33 Air-to-air refuelling pressure gauge

34 Flood flow switch. (inoperative)

35 Pressure headf,\00 heater switch

36 Ram air valve switch

37 ufv lighting dimmer switch

38 Ram air valve position indicator

39 Cold air unit overspeed indicator

40 Cabin temperature control switch

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5-1 Fig. 8 Fuel Contents Panel (2P) and Throttle Quadrant

JPT limiter switch 10 No. 4 engine group fuel contents gauge

RPM governor switch 11 HRS/MFS ~witch

Parking brake lever 12 Auto-throule engage pushbutton (inoperative)

No. 1 engine group fuel contents gauge 13 Airbrakes normal/emergency switch

No. 2 engine group fuel contents gauge 14 Throttle friction control

Cabin altimeter 15 Co-pilot's auto-throttle cut-out button (inoperative)

Rudder feel relief pushbutton 16 Relight buttons (four)

Rudder trim switch 17 Throttle levers (four)

No. 3 engine group fuel contents gauge 18 1st pilot's auto-throttle cut-our button (inoperative)

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AP 101B-1902-15 Illustrations

5-l Fig. 9 Alternator Control Panel ( lOP)

Alternator on/off switches (four) ll

Alternator KW /KVAR meters (four) 12

Alternator warning lights (four) 13

Alternator synchronising magnetic indicators (four) 14

Synchronising busbar voltmeter 15

RAT synchronising magnetic indicator 16

Incoming alternator voltmeter 17

RAT test pushbutton 18

Alternator failure red warning light 19

Alternator selector switch 20

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Extra supplies trip button

AAPP synchronising magnetic indicator

AAPP test button

Incoming alternator frequency meter

Synchronising busbar frequency meter

AAPP synchronising button

Alternator isolating buttons (four)

Alternator reset buttons (four)

KVAR reading selector button

Non-essential supplies reset switch

5-l Page 9

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7 a 9 10

S-1 Fig. 10 Secondary Supplies Panel (SOP)

Battery isolation switch

2 Banery magnetic indicator

3 Load-shed magnetic indicators

4 Port TRU switch

5 Port ·rnu magnetic indicator

6 Starboard TRU magnetic indicator

7 Starboard TRU switch

8 Port main transformer switch

j

9 Port main transformer magnetic indicator

replaced by

ammeters

post-Mod. 2195

10 Starboard main transformer magnetic indicator

11 Starboard main transformer switch

12 Crew's ration heater switch

13 Pilot~' ration heater switch

14 oc voltmeter

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5-l Fig. 11 AAPP Control Panel (70P)

Fuel level magnetic indicator

2 JPT gauge

3 csou oil temperature gauge for Nos. 1 and 2 units

4 csou oil temperature gauge for Nos. 3 and 4 units

S AAPP starter button

6 AAPP oil pressure gauge

7 Master switch

8 Oxygen and relight switch

9 Fire warning light test button

10 F ire warning light and extinguisher button

II Ignition isolation switch

12 liP cock override switch

13 LP cock switch

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AP 101 B-1902-1 5 Illustrations

5-l Page 11

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3 4 5 6

12 II 10

5-l Fig. 12 Radio Supply Panel (12P)

No. 1 frequency changer failure indicator

2 Radio altimeter master switch

3 NBS transformer switch

4 NBC supplies switch

S H2S supplies switch

6 GPI switch

7 Oscillator tuning switch

8 No. 2 frequency changer failure indicator

9 NBC emergency DC supplies switch

10 No. 2 frequency changer start and stop buttons

11 Frequency changer emergency changeover switch

12 No. 1 frequency changer start and ~top buttons

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