Defence Standard 00-970 Part 13

Issue 11 Date: 13 Jul 2015

_______________________________________
Design and Airworthiness
Requirements for Service Aircraft

Part 13: Military Common Fit

Equipment
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 DEF STAN 00-970 PART 13/11

REVISION NOTE

This standard is raised to Issue 11 to incorporate changes:

Note. Major revisions to this Part of the Defence Standard are noted in Part 0, Section 6.

HISTORICAL RECORD

This standard supersedes the following:

Defence Standard 00-970 Part 13 Issue 10 dated 30 January 2015.

Design and Airworthiness Requirements for Service Aircraft

Defence Standard 00-970 Part 13 Issue 9 dated 11 July 2014.

Defence Standard 00-970 Part 13 Issue 8 dated 10 January 2014.
Defence Standard 00-970 Part 13 Issue 7 dated 05 July 2013.
Defence Standard 00-970 Part 13 Issue 6 dated 07 January 2013.
Defence Standard 00-970 Part 13 Issue 5 dated 06 July 2012.
Defence Standard 00-970 Part 13 Issue 4 dated 31 October 2011.
Defence Standard 00-970 Part 13 Issue 3 dated 31 January 2011.
Defence Standard 00-970 Part 13 Issue 2 dated 15 January 2010.
Defence Standard 00-970 Part 13 Issue 1 dated 12 December 2007.

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 DEF STAN 00-970 PART 13/11

DESIGN AND AIRWORTHINESS REQUIREMENTS FOR SERVICE AIRCRAFT

MILITARY COMMON FIT EQUIPMENT

PREFACE

(a) This Part of the Defence Standard provides requirements for Airworthiness and Design
Certification for the design, development and testing of military common fit equipment for use on
UK Military aircraft in all classes of airspace. The requirements stated herein shall be applied by
the Ministry of Defence (MOD) and the contractor as agreed and defined in the contract.

(b) This document has been produced on behalf of the Military Aviation Authority Executive
Board (MEB) by the Military Aviation Authority (MAA), MAA Technical Group, MOD Abbey Wood.

(c) The appropriate Parts of this document are to be used, when called up in the Contract, for all
future designs, and whenever practicable for amendments to existing designs. If any difficulty
arises which prevents application of this document, DSA-MAA-Cert-ADS1 shall be informed so that
a remedy may be sought: e-mail: DSA-MAA-Cert-ADSGroup@mod.uk

(d) Where the requirements of other Standards are considered applicable, the relevant chapters
and/or clauses are cross-referenced by this Part of the Defence Standard.

(e) Any enquiries regarding this document in relation to an invitation to tender or a contract in
which it is incorporated are to be addressed to the relevant MOD Project Team Leader (PTL)
named in the invitation to tender or contract.

(f) Please address any enquiries regarding this standard, whether in relation to an invitation to
tender or to a contract in which it is incorporated, to the responsible technical or supervising
authority named in the invitation to tender or contract.

(g) Compliance with this Defence Standard shall not in itself relieve any person from any legal
obligations imposed upon them. Project Leaders are to ensure that equipment procured from
outside of the European Union (EU) meets or exceeds those legal requirements mandated within
the EU (See MAA 01 Chapter 1 and the RA1000 Series).

(h) This standard has been devised solely for the use of the Ministry of Defence (MOD) and its
contractors in the execution of contracts for the MOD. To the extent permitted by law, the MOD
hereby excludes all liability whatsoever and howsoever arising (including, but without limitation,
liability resulting from negligence) for any loss or damage however caused when the standard is
used for any other purpose.

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CONTENTS

Description Page

Preface iii
Contents iv

SECTION 0 GENERAL REQUIREMENTS

0.1 Scope 1
0.2 Warning 1
0.3 Normative References 1
0.4 Definitions 2
0.5 Abbreviations 2

SECTION 1 COMMON FIT EQUIPMENT

1.1 Navigation 4
1.2 Communication Systems 18
1.3 Data Recording Systems 20
1.4 Oxygen Systems 24
1.5 Ice Protection 71
1.6 Survivability and Recovery 87
1.7 Safety Related Programmable Elements 104

SECTION 2 COMMON FIT EQUIPMENT - LEAFLETS

0 References 2
1 View and Clear Vision - Standards of Rain 5
2 View and Clear Vision - Methods of Rain Clearance from Windscreens 6
3 Oxygen Systems - Physiological Requirements for Oxygen Systems 8
4 Oxygen Systems - Pressure Losses in Oxygen Delivery Systems 28
5 Oxygen Systems - Tests on Liquid Oxygen Systems 31
6 Ice Protection - Precautions to Prevent Waste Water Leaving Aeroplanes as Ice 33
7 Ice Protection - Icing Conditions 34
8 Ice Protection - Ice Protection Systems 40

SECTION 3 MILITARY SPECIFIC SYSTEMS

3.1 Armament Installations – General 2

3.2 Armament Control Systems 4
3.3 Gun Installations 15
3.4 Installation of Explosive Devices 19
3.5 Air to Air Refuelling 20
3.6 Arresting Hooks 43
3.7 Installations for Emergency Recovery from Stall and Spin 52
3.8 Target Towing Installations 61
3.9 Reduction of Vulnerability to Battle Damage 76
3.10 Protection of Aircrew against Conventional weapons 79
3.11 Protection from the Effects of Nuclear Explosions, Laser weapons,
Chemical and Biological Warfare Agents 82
3.12 Aircrew Equipment 91
3.13 Brake Parachute Installations 92
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3.14 Integration of Stores 96

3.15 Not Issued 101
3.16 Defensive Aids Systems (DAS) 102
Tables 115

SECTION 4 MILITARY SPECIFIC SYSTEMS - LEAFLETS

1 References 2
2 Armament Installations – Weapon Release and Fuzing 11
3 Armament Installations – Jettison Systems 15
4 Armament Installations – The Effect of Firing Air weapons on the Behaviour of
Turbine Engine Aircraft 17
5 Gun Installations – General Recommendations 20
6 Gun Installations – Gun Gas Concentrations 24
7 Gun Installations – Gun Blast: The Effect of Gun Firing on Turbine engines 25
8 Installation of Explosive Devices – General Recommendations 27
9 In-Flight Refuelling Systems – General Recommendations 30
10 Arresting Hooks for Land-Based Aeroplanes 47
11 Installations for Emergency Recovery from Stall and Spin – General
Information and Recommendations 60
12 Installations for Emergency Recovery from Stall and Spin – Parachute
Installations 64
13 Installations for Emergency Recovery from Stall and Spin – Rocket Installations 67
14 Target Towing Installations – Definitions and Glossary 69
15 Target Towing Installations – General and Operational Requirements 72
16 Target Towing Installations – Aerodynamic and Flying Qualities 74
17 Target Towing Installations – Loading and Shedding 76
18 Target Towing Installations – Cockpit Controls and Indicators 78
19 Reduction of Vulnerability to Battle Damage – General requirements 81
20 Protection of Aircrew against Conventional Weapons – General requirements 89
21 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents – Definitions 91
22 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents – Nuclear Weapon Effects on Aircraft 98
23 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents – General Recommendations – Chemical and
Biological Warfare Agents 107
24 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents – LASER Weapon Effects on Aircraft 117
25 Brake Parachute Installations – Safety and Strength Recommendations 125
26 Integration of Stores – Integration Methodology 129
27 Integration of Stores – Description of Design Considerations & Loading Actions 131

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DESIGN AND AIRWORTHINESS REQUIREMENTS FOR SERVICE AIRCRAFT

PART 13 - MILITARY COMMON FIT EQUIPMENT

SECTION 0 GENERAL REQUIREMENTS

0.1 Scope

This standard specifies requirements for Equipment common to more than one aircraft
type.

0.2 Warning

The Ministry of Defence (MOD), like its contractors, is subject to both United Kingdom and
European laws regarding Health and Safety at Work. Many Defence Standards set out
processes and procedures that could be injurious to health if adequate precautions are not
taken. Adherence to those processes and procedures in no way absolves users from
complying with legal requirements relating to Health and Safety at Work.

Note: Where a design to the requirements of this document may result in an adverse
environmental impact the MOD PTL shall be advised.

0.3 Normative References

0.3.1 The publications shown in Part 0 are referred to in the text of this standard.

Note: Def Stan’s can be downloaded free of charge from the DStan web site by visiting
http://dstan.uwh.diif.r.mil.uk for those with rli access or https://www.dstan.mod.uk for all
other users. All referenced standards were correct at the time of publication of this standard
(see 0.3.2, 0.3.3 & 0.3.4 below for further guidance), if you are having difficulty obtaining
any referenced standard please contact the DStan Helpdesk in the first instance.

0.3.2 Reference in this Standard to any normative references means in any Invitation to
Tender or contract the edition and all amendments current at the date of such tender or
contract unless a specific edition is indicated. Care should be taken when referring out to
specific portions of other standards to ensure that they remain easily identifiable where
subsequent amendments and supersession’s might be made. For some standards the most
recent editions shall always apply due to safety and regulatory requirements.

0.3.3 In consideration of clause 0.3.1 above, users shall be fully aware of the issue,
amendment status and application of all normative references, particularly when forming part
of an Invitation to Tender or contract. Correct application of standards is as defined in the ITT
or contract.

0.3.4 DStan can advise regarding where to obtain normative referenced documents.
Requests for such information can be made to the DStan Helpdesk. Details of how to contact
the helpdesk are shown on the outside rear cover of Defence Standards.

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0.4 Definitions

0.4.1 Definitions are contained in Part 0 of this standard and within the MAP, MAA 02.

0.5 Abbreviations

0.5.1 Abbreviations are contained in Part 0 of this standard.

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SECTION 1 COMMON FIT EQUIPMENT

1 This section specifies the requirements relating to safety equipment and systems embodied
on service aircraft. Requirements are provided to cover the following:

1.1 Navigation 4
1.2 Communication Systems 18
1.3 Data Recording Systems 20
1.4 Oxygen Systems 24
1.5 Ice Protection 71
1.6 Survivability and Recovery 87
1.7 Safety Related Programmable Elements 104

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SECTION 1 COMMON FIT EQUIPMENT

REQUIREMENT COMPLIANCE GUIDANCE

1.1 NAVIGATION
1.1.1 AIRCRAFT OPERATING LIGHTS - EXTERNAL LIGHTS
A - Fixed Wing
TAXI LAMPS
1.1.1.1 Taxying lamps shall be fitted on all
aeroplanes and shall be such as to illuminate
clearly any obstruction to the passage of the
aeroplane at a distance of 4 lengths of the
aeroplane ahead, and at any point from dead
ahead to 8 metres outside each wing tip.
LANDING LAMPS
1.1.1.2 All aeroplanes shall be provided The functions of taxying and landing lamps may
with sufficient landing illumination aids to permit a be combined in one unit e.g. a dual filament
safe landing in those modes and on those lamp.
surfaces (major runways, minor runways, unlit
roads, grass strips) for which night landings are
envisaged.
EXTERNAL LIGHTING CIRCUITS
1.1.1.3 On all aeroplanes required to This should normally include illumination of the
operate at night, the external lighting circuits shall sides of the fin and upper surfaces of the wing.
be controlled by a single master switch. All
aeroplanes shall be equipped with sufficient
illumination for night formation flying.
1.1.1.4 All external lighting shall be An ON/OFF switch located on the throttle is a
dimmable. desirable feature.
1.1.1.5 There should be no possibility of
downward recognition lights becoming obscured
with mud during take-off.
1.1.1.6 All pilot training aeroplanes shall
have an external light to indicate when the
undercarriage is locked down. The light shall be
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REQUIREMENT COMPLIANCE GUIDANCE

o
visible from ahead to 120 on either side of the
flight path to an observer, level with or below the
aeroplane. It shall be amber with an intensity of
between 4 and 25 candelas and shall not interfere
with the pilot’s night vision. In addition, on basic
training aircraft an automated warning tone, not
audible to the pilot, shall be added to all radio
transmissions when the undercarriage is locked
down.
LIGHTS
1.1.1.7 All military aeroplanes shall be The navigation system shall include anti-collision This content is similar to CS 25.1383 - CS
provided with navigation light systems, which will lights as well as the wingtip and taillights or their 25.1403. It contains requirements for aeroplane
provide illumination completely around the normal equivalent unless otherwise agreed with the navigation and anti collision lights based on an
plan of flight of the aeroplane. Project Team Leader (PTL) international agreement within NATO STANAG
3224. The coloured wingtip and white taillights or
their equivalent, provide direction of flight
information to pilots of other aircraft in the vicinity.
The anti-collision light system provides a signal,
which generally permits aeroplanes to be seen at
greater distances than aeroplanes provided only
with the wingtip and tail lights.
ANTI-COLLISION LIGHT SYSTEM
LOCATION
1.1.1.8 The anti-collision light(s) shall be
located so that the emitted light shall not be
detrimental to the crews' vision and will not detract
from the conspicuity of the navigation lights.
COLOUR
1.1.1.9 Red and/or Aviation White (while
operating) FIELD OF COVERAGE.
1.1.1.10 The system shall consist of such The field coverage shall extend in all directions
lights as will afford coverage of all areas around within 30° above and 30° below the horizontal
the aeroplane. plane of the aeroplane; except that obstructed
visibility totalling not more than 0.03 steradian
shall be permissible within a solid angle of 0.15
steradian centred about the longitudinal axis in

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REQUIREMENT COMPLIANCE GUIDANCE

the rearward direction.
FLASHING CHARACTERISTICS
1.1.1.11 The arrangement of the system, The flash frequency for any single light source
such as the number of light sources, beam width, shall not be less than 40 cycles per minute. The
speed of rotation, etc., shall be such as to give flash frequency shall not be more than 100
optimum flash frequency of 90 cycles per minute. cycles per minute except when the system
includes overlaps created by more than one light
source. In overlaps, the effective flash frequency
shall not exceed 180 cycles per minute. The
effective flash frequency is to be established as
that frequency at which the aeroplanes complete
anti-collision light system is observed from a
reasonable distance.
LIGHT INTENSITY
1.1.1.12 The minimum effective intensities in If a higher intensity is desired, a colourless glass Ie = effective intensity (candelas), and is the
all vertical planes, measured with the red filter, may be used. In this case, the value of the maximised value of the right-hand side of this
shall be in accordance with Guidance Table. effective intensity of the white light must be at equation.
least four (4) times higher than the minimum It = instantaneous intensity as a function of time.
intensity of the red light (Guidance table). The t2 - t1 = flash duration (seconds).
following relation shall be Note: The maximum value of Ie is obtained when
assumed: t2 and t1 are so chosen that the effective intensity
is equal to the instantaneous intensity at t2 and
t1.

MINIMUM EFFECTIVE INTENSITIES FOR ANTI-

COLLISION LIGHTS
Angle Effective Intensity Ie
Above and (Candelas)
Below Imin < 0.30 Imin > 0.30
Horizontal Imax Imax
Plane
0° to 5° 100 See Clause
5° to 10° 60 1.1.1.13.
10° to 20° 20
20° to 30° 10

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REQUIREMENT COMPLIANCE GUIDANCE

Imin = minimum intensity during "OFF" period
Imax = maximum intensity during "ON" period
Note: It is desirable to increase the angle above
and below the horizontal plane to 600. Between
200 and 600 the minimum effective intensity shall
be 10 candelas.
FLASH FREQUENCY Vs EFFECTIVE INTENSITY
1.1.1.13 The rise and decay characteristics of high current As an example, if the flash frequency is 45 cycles
lamps flashed by electrical means are such that per minute (a decrease from 90 cycles per minute
the intensity may not decay during the "OFF" of 50 percent), the effective intensity requirements
period to an acceptable level of less than 0.30 of Clause 1.1.1.12 shall be increased by 100
times the peak intensity. In such cases the flash percent.
frequency may be reduced to obtain an adequate
decay provided that the effective light intensity
(see Clause 1.1.1.12. above) is increased by
twice the percentage of flash frequency reduction
below 90 cycles per minute.
NAVIGATION LIGHT SYSTEM
WING LIGHTS (SIDE EXTREMITY LIGHTS)
1.1.1.14 Location - The wing lights shall be Such lights may be installed on other than the
spaced laterally and as far as practicable on the wing tip (for example on swept wing aeroplanes).
extremities of the wings. Supplementary lights may be installed in any
location as necessary to meet minimum light
distribution requirements. Each light as installed
shall show unbroken light in accordance with
Figs. 1 and 2.
1.1.1.15 Colour - The wing lights shall be
Aviation Red for the left wing and Aviation Green
for the right wing.
1.1.1.16 Candlepower - Candlepower
requirements shall be as shown in Figs. 1 and 2.
TAIL LIGHTS (AFT EXTREMITY LIGHTS)
1.1.1.17 Location - The tail light shall be Such lights may be installed on other than the
located as near as practicable to the rear aeroplane tail (for example aft wing tip on swept
extremity of the fuselage. wing aeroplanes). Supplementary lights may be
installed if necessary to meet the minimum
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REQUIREMENT COMPLIANCE GUIDANCE

distribution requirements of Figs. 1 and 2. On
aeroplanes where jet exhaust or wing
configuration would limit the distribution, tail lights
may be installed on the wing trailing edge.
1.1.1.18 Colour - Aviation White
1.1.1.19 Candlepower - The candlepower
requirements shall be as shown in Figs. 1 and 2.
1.1.1.20 Aeroplanes not provided with anti-
collision lights shall be furnished with flashing
navigation lights. The flash rate for the navigation
lights shall be 85 ±15 flashes per minute with an
"on” to "off” ratio of between 3:1 and 1.857:1. The
complete system, wingtip and tail lights, shall be
flashed simultaneously.
1.1.1.21 Naval aeroplanes shall be fitted
with flashing navigation lights controlled by a
FLASHING/OFF/STEADY switch and by a
BRIGHT/DIM switch. They shall, where possible,
be fitted with a flashing white light(s) visible in all
directions and be capable of independent
operation.
COLOURS
1.1.1.22 Chromaticity - Colour reference in Chromaticity co-ordinates are as follows; See also Fig. 3
this document shall have the applicable
International Commission on Illumination (CIE) (a) Aviation Red:
chromaticity co-ordinates. y is not greater than 0.335
z is not greater than 0.020

(b) Aviation Green:

x is not greater than 0.440 - 0.320y
x is not greater than y - 0.170
y is not less than 0.390 - 0.170x

(c) Aviation White:

x is not less than 0.300
x is not greater than 0.540

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REQUIREMENT COMPLIANCE GUIDANCE

y - y0 is not numerically greater than 0.01.
y0 is the y
co-ordinate of the Plankian radiator for
which "x0 = x"

(d) Red for Anti-collision Lights:

y is not greater than 0.350
z is not greater than 0.020
1.1.1.23 Colour Limits - Fig. 3 graphically
represents limit co-ordinates of Aviation colours.

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FIG 3

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REQUIREMENT COMPLIANCE GUIDANCE

B - Rotorcraft
1.1.1 For Operating Lights Refer to IR
OPS CAT.IDE.H.115
1.1.2. CLEAR VIEW AND VISION
A – Fixed Wing
1.1.2.1. A clear vision for piloting, search,
navigation, bomb and weapon aiming shall be
available under all weather conditions. Sufficient
area on transparent panels shall remain clear for
these purposes, in external conditions of snow,
rain and ice, and in internal conditions of mist and
ice.
RAIN REMOVAL FROM WINDSCREENS
1.1.2.2. Adequate vision through the As a minimum requirement, the system shall: See Leaflets 1 and 2.
windscreen(s) shall be provided for the aeroplane
roles envisaged in all rain conditions up to heavy (a) clear an area of the windscreen
rain. which will allow the pilot(s) to see from 10°
below the horizon to 4° above it and to 7°
on either side of the aircraft design eye
position when making an approach and
landing, and

(b) Provide clearance from zero

forward speed up to 2.0 times the normal
approach speed, unless otherwise stated in
the Aeroplane Specification.
Where an aeroplane in the landing
approach condition has a downward view
exceeding 10° the aim should be to
provide, in rain conditions the same
downward view as obtained in clear air
(see Part 1 Section 4 Clauses 4.17.8 to
4.17.11).
1.1.2.3. The system shall be under the (a) When switched on, the system shall be
control of the pilot, and shall be capable of being automatically controlled so that, below 10,000 ft
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REQUIREMENT COMPLIANCE GUIDANCE

used as soon as the engine is started. the full effect of the system shall be used for rain
removal. Above this height, full advantage shall
be taken of the de-icing and demisting
potentialities of the system.

(b) On multi-engine aeroplanes, failure of any

one engine shall not impair the performance of
the system.

(c) The Contractor shall demonstrate, by

ground tests before the aeroplane flies, and by
flight tests at an early stage in the development
of the aeroplane, the effectiveness of the system
installed.
PROTECTION OF WINDSCREENS FROM ICE
1.1.2.4. See Part 13 Section 1.5 Ice Protection.
OPTICALLY TRANSPARENT COMPONENTS
1.1.2.5. Transparencies shall be designed (a) Where applicable, those considered shall Note: Transient conditions can arise from speed
to meet all possible combinations of steady and include pressure, aerodynamic and inertia loads, and other effects or from combinations of them.
transient conditions, including thermal conditions effects produced by aeroplane systems, erosion, Where speed effects are involved flight at speeds
that can occur within the specified flight envelope hailstones, bird strike and handling. Steady above the maximum in level flight may, be
of the aeroplane or on the ground. conditions are defined as those flight or ground regarded as transient.
conditions enduring for five minutes or more. All
others are transient.

(b) The following points shall be considered

in design and represented during approval tests:

(1) outside air temperatures over the

full range specified in Part 1 Section 7
Clause 7.2;
(2) solar radiation,
(3) kinetic heating,
(4) cabin heating and cooling,

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REQUIREMENT COMPLIANCE GUIDANCE

(5) de-icing and demisting systems,
(6) heat runaway due to thermostat
failure,
(7) rain dispersal systems including
windscreen wipers,
(8) edge effects from mounting,
(9) altitude effects,
(10) erosion,
(11) those special requirements
detailed in Part 1 Section 4 Clause 4.18.8
which are called up in the Aeroplane
Specification.
B - Rotorcraft
1.1.2 Refer to Part 7, Leaflet 101-00,
Leaflet 104, Leaflet 715-00, Leaflet 1006.
1.1.3. EQUIPMENT FOR OPERATIONS REQUIRING A RADIO COMMUNICATION AND/OR RADIO NAVIGATION SYSTEM
A – Fixed Wing
1.1.3 N/A for Fixed Wing at this time.
B - Rotorcraft
1.1.3 Refer to IR OPS CAT.IDE.H.325
1.1.4. DAY VISUAL FLIGHT REFERENCE OPERATIONS – FLIGHT NAVIGATIONAL INSTRUMENTS AND ASSOCIATED EQUIPMENT
A – Fixed Wing
1.1.4 Refer to IR OPS CAT.IDE.A.125
B - Rotorcraft
1.1.4 Refer to IR OPS CAT.IDE.H.125
1.1.5. INTRUMENTED FLIGHT REFERENCE OPERATIONS (IFR) OR NIGHT VISUAL FLIGHT REFERENCE OPERATIONS (VFR) – FLIGHT
NAVIGATIONAL INSTRUMENTS AND ASSOCIATED EQUIPMENT
A – Fixed Wing
1.1.5 Refer to IR OPS CAT.IDE.A.130
B - Rotorcraft
1.1.5 Refer to IR OPS CAT.IDE.H.130

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REQUIREMENT COMPLIANCE GUIDANCE

1.1.6. ADDITIONAL EQUIPMENT FOR SINGLE PILOT OPERATIONS UNDER INSTRUMENTED FLIGHT REFERENCE
A – Fixed Wing
1.1.6 Refer to IR OPS CAT.IDE.A.135
B - Rotorcraft
1.1.6 Refer to IR OPS CAT.IDE.H.135
1.1.7. ALTITUDE ALERTING SYSTEM
A – Fixed Wing
1.1.7 Refer to IR OPS CAT.IDE.A.140
B - Rotorcraft
1.1.7 Refer to IR OPS CAT.IDE.H.145
1.1.8. GROUND PROXIMITY WARNING SYSTEM AND TERRAIN AWARENESS WARNING SYSTEM
A – Fixed Wing
1.1.8 Refer to IR OPS CAT.IDE.A.150
B - Rotorcraft
1.1.8 N/A to Rotorcraft at this time.
1.1.9. AIRBORNE COLLISION AVOIDANCE SYSTEM
A – Fixed Wing
1.1.9 For Military Transport Type Aircraft
Refer to IR OPS CAT.IDE.A.155
B - Rotorcraft
1.1.9 N/A to Rotorcraft at this time.
1.1.10. AIRBORNE WEATHER RADAR EQUIPMENT
A – Fixed Wing
1.1.10 Refer to IR OPS CAT.IDE.A.160
B - Rotorcraft
1.1.10 Refer to IR OPS CAT.IDE.H.160
1.1.11. COMMUNICATION AND NAVIGATION EQUIPMENT FOR OPERATIONS UNDER IFR, OR UNDER VFR OVER ROUTES NOT
NAVIGATED BY REFERENCE TO VISUAL LANDMARKS
A – Fixed Wing
1.1.11 Refer to IR OPS CAT.IDE.A.345
B - Rotorcraft
1.1.11 Refer to IR OPS CAT.IDE.H.345
1.1.12. ADDITIONAL NAVIGATION EQUIPMENT FOR OPERATIONS IN MINIMUM NAVIGATION PERFORMANCE SPECIFICATION AIRSPACE

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A – Fixed Wing
1.1.12 Refer to IR OPS SPA.MNPS.105
B - Rotorcraft
1.1.12 N/A to Rotorcraft at this time.
1.1.13. EQUIPMENT FOR OPERATION IN DEFINED AIRSPACE WITH REDUCED VERTICAL SEPARATION MINIMA (RVSM)
A – Fixed Wing
1.1.13 Refer to IR OPS SPA.RSVM.110
B - Rotorcraft
1.1.13 N/A to Rotorcraft at this time.
1.1.14. TRANSPONDER EQUIPMENT
A – Fixed Wing
1.1.14 Refer to IR OPS CAT.IDE.A.350
B - Rotorcraft
1.1.14 Refer to IR OPS CAT.IDE.H.350
1.1.15. CIRCUIT PROTECTION DEVICES
A – Fixed Wing
1.1.15 Refer to IR OPS CAT.IDE.A.110
B - Rotorcraft
1.1.15 Refer to Part 7, Leaflet 706-00 and
Leaflet 707-1
1.1.16. STOWAGE OF NAVIGATION MAPS AND CHARTS (ANO SCHEDULE 4 ARTICLE A2)
A – Fixed Wing
1.1.16.1. A map stowage, which shall also The position shall be determined in Consultation
include a separate compartment for Aircrew with the Project Team Leader.
Manuals/Pilot's Notes and Flight Reference Cards
(FRCs), shall be provided for each pilot.
B - Rotorcraft
1.1.16 Refer to Part 7, Section 7, Leaflet
712, Para 12.2

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1.2 COMMUNICATION SYSTEMS
1.2.1. RADIO EQUIPMENT
A – Fixed Wing
1.2.1 Refer to IR OPS CAT.IDE.A.285
B - Rotorcraft
1.2.1 Refer to IR OPS CAT.IDE.H.330
1.2.2. AUDIO SELECTOR PANEL
A – Fixed Wing
1.2.2 Refer to IR OPS CAT.IDE.A.335
B - Rotorcraft
1.2.2 Refer to IR OPS CAT.IDE.H.335
1.2.3. RADIO EQUIPMENT FOR OPERATIONS UNDER VFR OVER ROUTES NAVIGATED BY REFERENCE TO VISUAL LANDMARKS
A – Fixed Wing
1.2.3 Refer to IR OPS CAT.IDE.A.340
B - Rotorcraft
1.2.3 Refer to IR OPS CAT.IDE.H.340
1.2.4. FLIGHT CREW INTERCOMMUNICATION SYSTEM
A – Fixed Wing
1.2.4.1. Intercommunication shall be
provided between all operational stations inside
the aeroplane, and also to a point(s) on the
outside of the aeroplane for use by ground
personnel.
B - Rotorcraft
1.2.4 Refer to Part 7, Leaflet 707-1
1.2.5. CREW MEMBER INTERCOMMUNICATION SYSTEM
A – Fixed Wing
1.2.5.1. Intercommunication shall be
provided between all operational stations inside
the aeroplane, and also to a point(s) on the
outside of the aeroplane for use by ground
personnel.
B - Rotorcraft

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1.2.5 Refer to Part 7, Leaflet 707-1
1.2.6. PUBLIC ADDRESS SYSTEM
A – Fixed Wing
1.2.6 Refer to IR OPS CAT.IDE.A.180
B - Rotorcraft
1.2.6 Refer to IR OPS CAT.IDE.H.180

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1.3 DATA RECORDING SYSTEMS
A – Fixed Wing
1.3.1 COCKPIT VOICE AND FLIGHT This content provides requirements for Cockpit
DATA RECORDERS (CV/FDR). Voice/Flight Data Recorders (CVR/FDR) and
Health and Usage Monitoring Systems (HUMS)
and is similar to CS 25.1457 and CS 25.1459
1.3.1.1. All aircraft shall be fitted with a The CVR and FDR shall be compliant with Consideration should be given to the potential
Cockpit Voice (CV) Recorder and a Flight Data European Organisation for Civil Aviation requirement to inhibit recording and, particularly,
Recorder (FDR) as itemised in Table A below. Equipment (EUROCAE) specification ED-112, transmission of information which is often a
Minimum Operational Performance Specification feature of Commercial Off the Shelf (COTS)
for Crash-Protected Airborne Recorder Systems. CV/FDR & HUMS systems but which may conflict
with military operational requirements.
1.3.1.2. The letters "FDR" shall be painted Paint marking using heat resistant paint should Marking the external surface of the panel close to
on the external surface of the panel covering the be used. The marking should be as large and where the recording device is mounted is
structure to which the recording device is bright as practicable; see Part 1, Section 7 intended to make it easier to locate the FDR in
mounted. Clause 7.4 for operational colours and markings. the event of a crash.
Where the recording device is mounted on a
combat ac, paint that changes colour on heating
may be used. The marking should be as close
as practicable to the recorder location.
1.3.2. HEALTH AND USAGE MONITORING SYSTEM (HUMS)
1.3.2.1. All aircraft shall be fitted with an (a) The Project Team Leader (PTL) may See Part 0 for definition of HUMS.
appropriate HUMS. determine that the design and role of an aircraft
(e.g. glider or basic primary trainer) do not merit MOD preference is for post-flight ground based
fitment. processing but event/exceedence monitoring
could form part of the airborne system to provide
(b) The system requirements should be immediate alerts to maintenance organisations.
stated in the aircraft specification. The HUMS
should be an integrated system with sufficient
capacity and flexibility to enable the airborne unit
to be reconfigurable via the ground support
system. It may include FDR and CVR
functionality but should include structural loads
monitoring (See Part 1 Section 3 Leaflet 38),
engine condition inspection and monitoring

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capability (see Part 1 Section 4 Clause 4.4.6)
and transmission health monitoring as
appropriate.
1.3.3. SONAR LOCATING BEACONS (SLB) AND EMERGENCY LOCATOR TRANSMITTERS (ELT)
1.3.3.1. At least two SLB shall be installed Each SLB shall be installed in such a position The requirement should not constrain the
in an aircraft such that they are afforded free that the sonic energy is best transferred to the installation of deployable SLBs which may offer
contact with the water on immersion and that they surroundings. The location of the SLB on the increased survivability and retrieval probabilities
do not separate from the recording device or airframe shall be as agreed with the Project over conventional installations.
airframe when subjected to impact shock. Team Leader (PTL) and any installation conflicts
and difficulties resolved.
1.3.3.2. SLBs shall meet the requirements (a) SLBs operating at 9.5 kHz shall meet the
of current CAA or military specifications and, if requirements of ETSO C200 and SAE AS 6254.
necessary, the requirements of Clauses 1.3.3.3. to
1.3.3.6. (b) SLBs operating at 37.5 kHz shall meet
the requirements of ETSO C121b, and SAE
AS8045 or Civil Aviation Authority (CAA)
specification 12.
1.3.3.3. For aircraft fitted with HUMS which Consideration should be given to integrating the
have parameters relevant to post crash analysis SLB within the CSMU or combined CV/FDR as
recorded on a crash survivable Memory Unit appropriate.
(CSMU):

(a) An SLB shall be securely attached

to the CSMU and shall operate at a
frequency of 37.5 kHz.

(b) A second SLB shall be secured to

the airframe and shall operate at a
frequency of 9.5 kHz.
1.3.3.4. For aircraft fitted with FDR and
CVR in separate crash protected units:

(a) SLBs shall be securely attached to

both the FDR and the CVR and shall
operate at a frequency of 37.5 kHz.

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(b) A separate SLB shall be secured to
the airframe and shall operate at a
frequency of 9.5 kHz.
1.3.3.5. For aircraft not fitted with HUMS or
CVR/FDR:

(a) An SLB shall be secured to the

airframe in the area of the cockpit and shall
operate at a frequency of 37.5 kHz.

(b) A second SLB shall be secured to

the airframe at an alternative location and
shall operate at a frequency of 9.5 kHz.
1.3.3.6. For Aircraft fitted with a deployable
FDR/CVR:

(a) An ELT, which complies with the

requirements of EUROCAE Documents ED-
112 and ED-62 shall be fitted to the
FDR/CVR unit in place of an SLB.

(b) An SLB shall be secured to the

airframe and shall operate at a frequency of
9.5 kHz or as stated in the aircraft
specification.
1.3.3.7. The installation of ELTs in an
aircraft shall comply with the requirements in
EUROCAE Minimum Operational Performance
Specification ED62.
B - Rotorcraft
1.3.1 COCKPIT VOICE AND FLIGHT This content provides requirements for Cockpit
DATA RECORDERS (CV/FDR). Voice/Flight Data Recorders (CVR/FDR) and
Health and Usage Monitoring Systems (HUMS)
and is similar to CS 29.1457 and CS 29.1459.
1.3.1.1 All rotorcraft shall be fitted with a As 1.3.1.1 Fixed Wing As 1.3.1.1 Fixed Wing
Cockpit Voice Recorder (CVR) and Flight Data

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Recorder (FDR) as itemised in table B below.
1.3.1.2 The letters "FDR" shall be painted As 1.3.1.2 Fixed Wing As 1.3.1.2 Fixed Wing
on the external surface of the panel covering the
structure to which the recording device is
mounted.

Fixed Wing FDR/CVR Requirements

1 2
MCTOM (X Kg) MO FDR CVR
ICAO Type II A FDR or
<9 EUROCAE Class 3 CVR
X<5700
Or Combined FDR/CVR
>9 ICAO Type II A FDR EUROCAE Class 3
X>5700 N/A EUROCAE FDR Class A EUROCAE Class 1

Table A Fixed Wing FDR/CVR Fitment Requirements

Rotary Wing FDR/CVR Requirements

MCTOM (Kg) FDR CVR
EUROCAE FDR Class B or
X<3175 EUROCAE Class 3 CVR or
Combined FDR/CVR
X>3175 EUROCAE FDR Class B EUROCAE Class 3

Table B Rotary Wing FDR/CVR Fitment Requirements

1. MCTOM - Maximum Certified Take-off Mass.

2. MO – Maximum Occupancy (includes crew).

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1.4 OXYGEN SYSTEMS
A – Fixed Wing
General Background
This document defines the physiological requirements for, and the design and installation of, aircraft breathing systems employing stored gaseous oxygen
(GOX), stored liquid oxygen (LOX) or an on-board oxygen generating (OBOG) system as the breathing gas source. Further requirements are defined in
the referenced documents, but in all cases this document shall take precedence.
For a detailed discussion of the physiological requirements for oxygen systems reference should be made to LEAFLET 3. There follows here an
introductory summary.
An aircraft oxygen system must provide protection against hypoxia and the inhalation of toxic smoke and fumes in the event of contamination of the cabin
atmosphere with these materials. In some highly agile aircraft the oxygen system also provides protection against the effects of high +Gz accelerations
[see leaflet 3, section 6].
Physiological Effects of Altitude
The physiological effects of exposure to altitude fall into four categories:

• effects due to the low partial pressure of oxygen in the ambient air (hypoxia).
• effects related to gas containing cavities of the body.
• decompression sickness (condition produced by evolution of bubbles of nitrogen and other gases in the tissues).
• effects due to low ambient temperature.
• ebullism (vaporisation of tissue fluids at altitudes above 63,000 feet).

The primary function of an aircraft breathing system is to prevent hypoxia in the face of the fall of the absolute pressure of the environment to which the
crew are exposed. In providing protection against hypoxia the oxygen system must not impair the escape of expanding gases from the lungs associated
with rapid ascent or rapid decompression or replacement of gases into the middle ear and nasal sinuses during descent.
The concentration of nitrogen in the gas delivered by an oxygen system affects the incidence of decompression sickness following decompression of the
cabin and exposure of the crew to altitudes above 18,000 to 22,000 feet. In certain operational roles an oxygen system may be required to provide 100%
oxygen at low altitude in order to provide protection against decompression sickness at high altitude. Whilst an aircraft oxygen system does not provide
protection against the effects of low environmental temperatures it must perform satisfactorily at the low temperatures which can occur following
decompression of the cabin or escape at high altitude”.
There are three techniques employed to provide protection against hypoxia at altitude for aircraft personnel:

(a) Pressurisation of the cabin with air so that the absolute pressure of the cabin (cabin altitude) is higher than the absolute pressure of the
atmosphere at the height at which the aircraft is flying (aircraft altitude). Normally the cabin altitude increases with increasing aircraft altitude. The
maximum cabin altitude may be set so that it does not exceed the value (5,000 to 8,000 feet) at which it is acceptable for the crew and passengers to
breathe air. Alternatively, as in high performance military combat aircraft; the maximum cabin pressure differential may be limited to 35 to 40 kPa (5
to 6 psi) in order to reduce the weight of the cabin structure and the potential for a sudden failure of the pressure cabin to injure the crew.

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(b) Increasing the concentration of oxygen in the gas breathed by the occupants on ascent to altitude so as to maintain the partial pressure of
oxygen (PAO2) in the lung (alveolar) gas greater than 103 mmHg in order to prevent hypoxia. This technique is widely used in military combat aircraft
in association with limited pressurisation of the cabin. Sea level equivalent oxygen delivery to the body can be maintained by breathing up to 100%
oxygen at altitudes up to 34,000 feet. Above this altitude the alveolar PAO2, even breathing 100% oxygen, falls below that associated with breathing
air at sea level i.e. 103 mmHg.

(c) Increasing the absolute pressure at which 100% oxygen is delivered to the respiratory tract above the absolute pressure of the immediate
environment (pressure breathing) allows the partial pressure of oxygen (PAO2) in the lung (alveolar) gas to be maintained. Typically such pressure
breathing is introduced at 38,000 to 40,000 feet. This may be accomplished by means of specialised mask/headgear with or without counter-
pressure garments, or by a partial or full pressure suit system. This technique is employed to prevent severe hypoxia in the emergency situations of
loss of cabin pressurisation or escape at altitudes above 38,000 to 40,000 feet).

Successful aircraft life support systems achieve their aims by employing each of these three methods, generally in the order given above, to achieve the
most suitable system in terms of performance, cost and size.
The physiological requirements and hence the design of aircraft breathing systems depend upon the performance of the aircraft and the nature of the
duties of its occupants, so that the requirements of the pilot of a highly agile high altitude aircraft differ markedly from those of passengers following
decompression of a transport aircraft.
In practice, typical aircraft breathing systems consist of the following essential components:

(a) A source of breathing gas. This may be from a pre-charged gaseous or liquid oxygen storage system, from candles producing oxygen as a
result of chemical reaction or from an onboard oxygen generator (OBOG). OBOG-based systems produce oxygen-enriched breathing gas from
pressurised engine bleed air and/or from an auxiliary source of pressurised air.

(b) A pressure demand breathing regulator. This controls the flow of breathing gas in response to the respiratory needs of the aircrew.

(c) An oronasal mask. This is connected by hoses and connectors to the outlet of the regulator.

(d) An emergency or backup supply. This typically consists of bottled 100% gaseous oxygen which may be aircraft or seat-mounted to suit the
operational need.
This document considers general and physiological aspects common to all aircraft breathing systems, as well as issues specific to the source of breathing
gas, the type of system, and the type of aircraft for which it is intended. Aspects of system testing are also covered herein.
Definitions:
Use of “shall” and “should” within this document shall be in accordance with the following definitions:-
The word “shall” in the text expresses a mandatory requirement. Departure from such a requirement is not permissible without formal written agreement
between all affected parties.
The word “should” in the text expresses a recommendation or advice on implementing a mandatory requirement. It is expected that such
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recommendations or advice will be followed unless good reasons are stated and agreed for not doing so.
Reference Documents:
Where not specified, the issue of the referenced documents shall be that current at the issue of this Def Stan.
Symbols and Abbreviations:
AEA Aircrew Equipment Assembly
AGV Anti-G Valve
ATPD Ambient Temperature and Pressure Dry gas
BC Biological & Chemical
BOS Back-up Oxygen System
EO Emergency Oxygen
GOX Gaseous Oxygen
+Gz Head wards acceleration (also called positive G or “eyeballs down”)
ICD Interface Control Document
LOX Liquid Oxygen
MSOC Molecular Sieve Oxygen Concentrator
NBC Nuclear, Biological and Chemical
NTP Normal Temperature and Pressure (15°C at 1 Atmosphere)
OBOG On-Board Oxygen Generator
OBOGS On-Board Oxygen Generating System(s)
PAO2 Partial pressure of oxygen in the alveolar gas
PBA Pressure Breathing at Altitude
PBG Pressure Breathing with G
PSA Pressure Swing Adsorption
RAF CAM Royal Air Force Centre of Aviation Medicine
This section is similar to CS 25.1439 – CS
25.1453. It defines the physiological requirements
for, and the design and installation of, aeroplane
breathing systems employing stored gaseous
oxygen, stored liquid oxygen or an on-board
Molecular Sieve Oxygen Concentrator (MSOC) as
the breathing gas source. Further references can
be found in ASCC standards 25/34 and 15/14 and
STANAGs 3865, 7106, 3499 and 3198.

NB “The UK has ratified STANAG 7106 but is

not in full compliance. The agreement specifies a

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limit of 25 ppm of methane in liquid breathing
oxygen whereas the UK limit is 50 ppm.”

An aeroplane breathing system consists of the

following essential components:

(a) A source of breathing gas which

may either be from a pre-charged gaseous
or liquid oxygen storage system or from an
oxygen concentrator which produces
breathing gas from engine bleed air and/or
auxiliary source of pressurised air.

(b) A pressure demand regulator,

which controls the flow of breathing gas in
response to the respiratory needs of the
aircrew.

(c) An oronasal mask connected by

hoses and connectors to the outlet of the
regulator.

(d) An emergency supply of gaseous

oxygen.
1.4.1 PHYSIOLOGICAL REQUIREMENTS - GENERAL
1.4.1.1. Breathing systems for aircrew shall: This requirement relates to STANAG 3865

(a) Prevent significant hypoxia whilst

the cabin is pressurised without inducing
acceleration atelectasis or delayed optic
barotraumas.

(b) Prevent significant hypoxia

following decompression of the cabin to
altitudes of up to the maximum cabin
altitude, which can occur in flight.

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(c) Meet respiratory demands without
imposing excessive resistance to breathing.

(d) Provide, when required, safety

pressure in the mask cavity to prevent
inboard leakage of environmental air.

(e) Provide pressure breathing

automatically (including where necessary, a
gas supply for inflation of counter pressure
garments etc., for high altitude protection),
when the cabin altitude exceeds 38000 feet
and, when selected manually, at ground
level (press-to-test facility).

(f) Prevent the pressure in the mask,

generated either by trapped gas on rapid
decompression or by a failure of the demand
mechanism, from exceeding acceptable
physiological limits.

(g) Not produce significant oscillations

of pressure within the mask cavity.

(h) Not impose excessive re-breathing

of expired gas.

(i) Provide, where required, pressure

breathing as a means to enhance tolerance
to +Gz acceleration.
1.4.2. PHYSIOLOGICAL REQUIREMENTS - SPECIFIC
CONCENTRATION OF OXYGEN IN THE INSPIRED GAS
1.4.2.1. Cabin Pressurised. The minimum RESPIRATORY DEMAND This requirement relates to STANAG 3865
concentration of oxygen in the inspired gas when
the cabin is pressurised shall be sufficient to: The performance of the oxygen system shall The curves describing the minimum concentration
meet the requirements at individual pulmonary of oxygen required to prevent hypoxia on rapid
(a) Maintain an alveolar oxygen ventilations (defined as flow averaged over 30 decompression will vary according to the cabin
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tension of at least 13.33 kPa (100 mmHg). seconds) between 5 and 50 L (ATPD) min-1 and pressurisation schedule. Details of the method of
peak inspiratory and expiratory flow demands of construction, with examples, are given in Clause
(b) Prevent the alveolar oxygen up to 200 L (ATPD) min-1 with maximum rates of 1.4 but a typical example of a physiological
tension falling below 4.0 kPa (30 mmHg) on change of 10 L (ATPD) sec-2 at peak flows of 90 specification for oxygen concentration with cabin
rapid decompression when 100% oxygen is L (ATPD) min-1 and 20 L (ATPD) sec-2 at peak altitude is shown graphically at Leaflet 3 Fig. 1.
delivered to the mask immediately after the instantaneous flows of 200 L (ATPD) min-1. NOTE: ATPD - Atmospheric Temperature and
decompression (94 - 99% oxygen in the NOTE: These parameters can only be set up Pressure Dry gas - is the volume of gas
case of a system employing a MSOC) as nominal values as the rates of change of flow expressed as dry gas at the prevailing
will be modified by the delivery characteristics of atmospheric pressure and temperature. In the
the regulator under test. context of this Clause, the atmospheric pressure
It is accepted that there may be short periods is the absolute pressure of the gas within the
with an MSOC system when the concentration of mask cavity and the temperature is constant at
oxygen in the inspired gas given above may not +15°C).
be achievable, for example, low engine power
settings associated with idle descent or engine
idle on the ground (see Clause 1.4.2.5).
1.4.2.2. For steady-state cabin altitudes, For aircrew, the relationship between the See LEAFLET 3, Physiological Requirements,
and changes in cabin altitude resulting from concentration of oxygen in the gas delivered to Section 4 “Composition of Inspired Gas”.
changes in aircraft altitude, the concentration of the mask and the cabin altitude shall be such Note that the characteristic curve implied by the
oxygen in the inspired gas shall be maintained at that:- compliance conditions (showing minimum
a level sufficient to prevent hypoxia. required breathing gas oxygen concentration
(a) The partial pressure of oxygen in versus cabin altitude) will be specific to each
the alveolar gas (PAO2) is maintained at or application, since it depends on (a) the aircraft’s
above 13.33 kPa (103 mmHg) at cabin cabin pressurisation schedule, (b) the aircraft’s
altitudes up to 34,000 feet. maximum operating altitude, and (c) the delivery
pressure of the breathing regulator (which may
(b) 100 % oxygen is delivered above vary with cabin altitude).
a cabin altitude of 34,000 feet. The curve may be generated using the Alveolar
Gas Equation (LEAFLET 3, paragraph 4.4) for
For passengers the concentration of oxygen in each altitude in the range of interest.
the inspired gas shall be such as to maintain a
minimum alveolar PAO2 of 7.33 kPa (55 mmHg)
at altitudes between 10,000 and 18,500 feet and
6.00 kPa (45 mmHg) at altitudes above 18,500
feet.
1.4.2.3. For aircraft with pressurised cabins, For aircraft with low cabin volume giving a risk of The limit of 0.6 L ATPD of gas which can be
on decompression of the cabin to a cabin altitude realistic worst-case decompression times of less inspired from the commencement of a low cabin
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above 30,000 feet the alveolar oxygen tension than 3s (typically single and dual-seat aircraft) volume decompression before the concentration
shall not fall below 4.00 kPa (30 mmHg). the concentration of oxygen in the gas delivered of oxygen gas delivered to the inlet of the mask
to the mask cavity shall rise to above 99.5% rises to 99.5% determines the maximum volume
when no more than 0.6 L ATPD of gas has been of the low pressure delivery system from the
inspired after the commencement of the outlet of the breathing regulator to the mask.
decompression. In systems using OBOGS as the primary gas
For aircraft with higher cabin volumes and source, in which the concentration of oxygen in
correspondingly slower realistic worst-case the product gas supplied to the regulator prior to
decompressions the concentration of oxygen in decompression is less than 100%, account should
the gas delivered to the mask cavity shall rise to also be taken of the quantity (expressed as
above 99.5% within a few seconds from the start volume ATPD) of product gas between the point
of decompression. The design aim for this time at which 100% oxygen is delivered into the
shall be 2 seconds or less. system and the demand valve of the regulator. A
In all cases the prevention of alveolar oxygen purging arrangement, typically initiated
tensions of less than 4.00 kPa (30 mmHg) during automatically by the fall in pressure during
the decompression shall be achieved by decompression, may be necessary to ensure that
maintaining oxygen concentration of the inspired 100% oxygen reaches the furthest breathing
gas during pressurised flight above that required regulator from the gas source in an acceptable
in paragraph 0. above in accordance with the time.
equation in LEAFLET 3, paragraph 4.7
1.4.2.4. The volume of dead space external The effective respiratory dead space of an Re-breathing of expired gas (which will be present
to the respiratory tract shall be minimised. oronasal mask or pressure helmet shall not in any dead space) lowers alveolar oxygen
exceed 0.2 L ATPD. tension, raises alveolar CO2 tension, and
increases pulmonary ventilation.
1.4.2.5. For agile aircraft, for pressurised For agile (high +Gz) aircraft the maximum The compliance limits are a compromise taking
cabin altitudes at which the aircraft is capable of concentration of oxygen in the inspired gas shall into account cabin pressurisation, the low
sustaining Gz accelerations greater than +3 G, the not exceed 60% at cabin altitudes up to 15,000 probability of sustained high +Gz manoeuvres at
concentration of oxygen in the inspired gas shall feet rising to 75% at a cabin altitude of 20,000 high altitude, and the technical difficulty and cost
not be so high as to induce acceleration feet. associated with designing systems with tight
atelectasis. control (max-min) limits on oxygen concentration.
The limits to maximum oxygen concentration do
not apply following decompression at altitude
since in this event prevention of hypoxia is of
more immediate concern.
1.4.2.6. The concentration of oxygen in the It is generally desirable for the maximum See Guidance notes in Clause 1.4.2.5 above.
inspired gas with the cabin pressurised shall not concentration of oxygen in the inspired gas not to
be so high as to induce delayed otitic barotraumas exceed 60% at cabin altitudes from Sea Level to
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(“oxygen ear”). 15,000 feet.
1.4.2.7. The adverse effects of smoke or For systems using air dilution regulators a means On shutting the air dilution regulator air inlet
toxic fumes in the inspired gas shall be mitigated. shall be provided of shutting the air inlet orifice orifice the breathing gas is obtained entirely from
during flight, including manual selection below the primary gas source, giving protection against
12,000 feet cabin altitude. contaminants in the cabin air. Clearly if the
For systems using OBOGS as the primary gas primary source is stored gas its duration will be
source a means shall be provided of manually significantly reduced in this mode.
selecting a breathing gas supply of 100% oxygen Systems using OBOGS as the primary gas
during flight. source with non-dilution regulators offer inherent
For larger aircraft one or more portable oxygen protection against smoke and fumes originating
sets containing 100% oxygen shall be provided within the cabin since the breathing gas originates
so as to allow the aircrew to tackle a cabin fire. externally (typically from engine bleed air).
Depending on its principle of operation the
OBOGS may also offer some protection against
smoke and fumes originating within the engine or
external to the aircraft (the OBOG may act as a
filter).
Typically, the secondary (backup) store of 100%
oxygen in OBOGS-based systems is relatively
small, so the duration of manually-selected 100%
oxygen may be low, and since the main functions
of such a back-up store is to provide protection
against hypoxia in the event of a failure of the
MSOC to provide a supply of product gas
containing an adequate concentration of oxygen
or on decompression of the cabin, its use for
smoke and fumes protection may limit the safe
maximum aircraft altitude for the remainder of the
mission.
Breathing Demands
1.4.2.8. The oxygen system shall meet the For aircrew, the oxygen system shall be capable See LEAFLET 3, Physiological Requirements,
requirements of this Standard when supplying the of supplying sustained pulmonary ventilations in Para 2.7 and 2.8 “Minimum and Maximum
range of sustained individual pulmonary the range of 5.0 L.min-1 ATPD to 50 L.min-1 Pulmonary Ventilation”.
ventilations likely to be demanded under all flight ATPD per crewmember. Pulmonary ventilation is equivalent to average
conditions. For passengers, the oxygen system shall be inspiratory or expiratory flow. Sustained
capable of supplying sustained pulmonary pulmonary ventilation is defined as a value
ventilations in the range of 5.0 L.min-1 ATPD to averaged over 30 seconds or longer.
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30 L.min-1 ATPD per person.
1.4.2.9. The oxygen system shall meet the For single-seat aircraft, the oxygen system shall See LEAFLET 3, Physiological Requirements,
requirements of this Standard when supplying the be capable of meeting peak inspiratory or paragraph 2.11 and 2.13.
maximum instantaneous respiratory flows (peak expiratory flows of 3.33 L.s-1 (200 L.min-1) For a two-seat aircraft, the system peak flow
flows) likely to be demanded under all flight ATPD. The design aim shall be to meet peak requirement will be (0.85 x 2 x 3.33 = ) 5.66 L.s-1
conditions. flows of 4.17 L.s-1 (250 L.min-1) ATPD. (340 L.min-1) ATPD, with a design aim of (0.85 x
For two-seat or multi-seat aircraft (in which the 2 x 4.17 =) 7.09 L.s-1 (425 L.min-1) ATPD.
breathing patterns of the crew are unlikely to It should be noted that these requirements should
coincide exactly in time), the oxygen system shall be met even in the presence of other demands on
be capable of meeting the following percentage the system (e.g. garment inflation flows, demist
of the peak inspiratory or expiratory flows which flows).
could be demanded by all crewmembers
breathing exactly in phase:

85% for aircraft with 2 seats

75% for aircraft with 3 seats
65% for aircraft with 4 or more seats
1.4.2.10. The oxygen system shall not The oxygen system shall be capable of meeting See LEAFLET 3, Physiological Requirements,
impose significant impedance to respiration when the requirements of this section (and in particular Para 2.12.
delivering the maximum rate of change of those for mask cavity pressures) when supplying The rates of change of flow are defined as being
respiratory flow likely to be demanded under all rates of change of flow of at least 10 L.s-2 ATPD between zero flow and 90% of the relevant peak
flight conditions. at a peak flow of 1.5 L.s-1 (90 L.min-1) ATPD, flow.
and 20 L.s-2 ATPD at a peak flow of 3.33 L.s-1 Impedance to respiration is in practice defined in
(200 L.min-1) ATPD. terms of the relationship between mask cavity
pressures and respiratory demands (flows). See
TABLE 1 below in Clause 1.4.2.13
Mask Cavity Temperature & Pressure
1.4.2.11. The temperature of the inspired gas The temperature of the breathing gas at the It is probable that heat transfer between the cabin
shall not induce discomfort or distraction. mask should be within ±10 °C of the cabin environment and the breathing gas distribution
ambient temperature under both normal and ducts will accomplish this.
emergency operating conditions. A temperature control system is unlikely to be
necessary.
1.4.2.12. The oxygen system shall prevent The mean mask cavity pressure shall not fall Leakage of cabin ambient air into the mask cavity
significant ingress of cabin ambient air through a below cabin ambient pressure at inspiratory peak will increase the risk (a) of hypoxia due to dilution
leak in the mask by generating a safety pressure flows of up to at least 1.2 L/s (72 L/min) ATPD. of the breathing gas, and (b) from toxic fumes or
in the mask cavity. This characteristic is essential at cabin altitudes BC warfare agents.

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above 12,000 feet, and highly desirable at other Excessive safety pressure will produce an
cabin altitudes. undesirably high resistance to expiration.
In systems which automatically provide safety At cabin altitudes above 38,000 feet the
pressure above 12,000 feet cabin altitude it shall requirement for positive pressure breathing will
be possible to manually select safety pressure at supersede that for safety pressure.
lower cabin altitudes.
1.4.2.13. At cabin altitudes between 0 and The minimum pressure, maximum pressure and In certain aircraft installations using OBOGS as
38,000 feet the relationship between mask cavity total change of pressure in the mask cavity over the primary gas source the pressures at which air
pressures and respiratory demands (flows) shall the respiratory cycle (“pressure swing”) for the is supplied to the OBOG at the low engine power
be such as to reflect minimal impedance to specified demand flows at altitudes up to 38,000 settings associated with idle descent and engine
respiration. feet shall not exceed the limits listed in TABLE 1 idle on the ground may be insufficient to provide
below. flow of product gas to an adequate pressure.
For systems using OBOGS as the primary gas To meet this requirement it will be necessary to
source, the mask cavity pressures and pressure advise the airframe and engine manufacturers of
swings (which are an indication of the impedance the need for a minimum air supply pressure.
to respiration imposed by the breathing system) The ideal minimum supply pressure at the inlet to
shall not normally exceed the limits listed in the the breathing regulator to attain a peak flow of
table below. 200 L/min is 100 kPag (15 psig).
At the worst-case (lowest) OBOGS supply
pressure (engine idle power setting) the
breathing system should be capable of meeting
individual pulmonary ventilations of at least 25
L/min ATPD and peak inspiratory demands of at
least 1.5 L/s (90 L/min) ATPD with mask cavity
pressure swings not exceeding 1.5 kPa (6.0 in
wg).
TABLE 1 - LIMITS TO MASK CAVITY PRESSURE AT ALTITUDES BETWEEN 0 AND 38 000 FEET

Peak Inspiratory and Expiratory Mask Cavity Pressure kPa (in wg)
Flows Limits to
(litres ATPD per sec)
Without Safety Pressure Minimum Maximum Maximum Swing
0.5 -0.38 (-1.5) +0.38(+1.5) 0.50 (2.0)
1.5 -0.55 (-2.2) +0.65(+2.6) 0.85 (3.4)
2.5 -1.12 (-4.5) +1.00(+4.0) 1.75 (7.0)
3.3 -1.90 (-7.6) +1.50(+6.0) 3.00 (12.0)

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With Safety Pressure
0.5 +0.02 (+0.1) +0.75(+3.0) 0.50 (2.0)
1.5 -0.20 (-0.8) +0.95(+3.8) 0.85 (3.4)
2.5 -0.90 (-3.5) +1.25(+5.0) 1.75 (7.0)
3.3 -1.75 (-7.0) +1.65(+6.6) 3.00 (12.0)
1.4.2.14. At cabin pressure altitudes between The mean mask cavity pressure (averaged over The alveolar oxygen tension (PAO2) when 100%
38,000 and 50,000 feet pressure breathing with the respiratory cycle) shall increase linearly with oxygen is breathed at ambient pressure falls
mask alone shall be used. The mean mask cavity fall of cabin ambient pressure from 20.6 kPa below 7.3 kPa (55 mmHg) on exposure to
pressure shall increase linearly with fall of cabin (38,000 feet) to 11.6 kPa (50,000 feet). The altitudes above 40,000 feet, this value of PAO2
ambient pressure so as to increase the alveolar mean mask cavity pressure shall be within the being that below which tasks requiring complex
oxygen tension sufficiently to provide additional range +0.1 to +1.0 kPa (0.75 to 7.5 mmHg) at hand-eye coordination are affected (it is
short term protection against hypoxia. 18.8 kPa ambient (40,000 feet), and within 4.0 to equivalent to breathing air at about 10,000 feet
4.5 kPa (30.0 to 33.8 mmHg) at 11.6 kPa altitude).
ambient (50,000 feet). Continuous positive pressure breathing with
Within this range of cabin ambient pressures 100% oxygen is employed above 40,000 feet to
(altitudes) the mask cavity pressure swings shall provide short duration protection against hypoxia
not exceed the maxima specified in TABLE 1 on decompression of the cabin and during
above in Clause 1.4.2.13 subsequent emergency descent in the aircraft or
following escape at high altitude.
Several partial pressure assemblies each with its
own relationship between mask pressure and
altitude can be employed to provide this
protection as denoted in the following
Requirements in this section.
1.4.2.15. Pressure breathing with a mask The mean mask cavity pressure shall increase The trunk counter pressure may be applied by a
together with counter pressure applied to the trunk linearly with fall of environmental pressure pressure jerkin alone or a chest counter pressure
and lower limbs shall be used to provide short between 38,000 and 60,000 feet. garment (pressure waistcoat) and the abdominal
duration hypoxia protection at altitudes up to The mean mask cavity pressure shall be within bladder of G trousers (see Clause 1.4.2.18). The
60,000 feet. The mean pressure in the mask the limits +0.1 to +1.0 kPag (0.75 to 7.5 mmHg) lower limb counter pressure is applied by the G
cavity and counter pressure garment(s) shall at 20.6 kPa (38,000 feet) and within +9.1 to 9.6 trousers. The bladder of the respiratory counter
increase linearly with fall of cabin ambient kPag (68.0 to 72.0 mmHg) at 7.2 kPa (60,000 pressure garment is connected into the breathing
pressure. feet). gas delivery demand regulator and the pressure
The mask pressure swings at various peak breathing mask.
respiratory flow rates shall not exceed the
maxima specified in TABLE 1 above in Clause
1.4.1.14

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The difference between the pressures in the
mask cavity and respiratory counter pressure
garment shall at no time exceed 0.5 kPag (3.8
mmHg).
The rate of inflation of the respiratory counter
pressure garment on decompression
(decompression time of 0.1s) shall be such that
the pressure in the mask cavity and garment
does not fall below 16 kPa absolute (120 mmHg
absolute) for longer than 2s, and the final
breathing pressure is to be attained within a
further 2s.
The lower limb garment shall be inflated to the
same time criteria. The final pressure in the lower
limb garment shall be 1 to 2 times the breathing
pressure depending upon the types of trunk and
lower limb garments employed.
Lower limb garment pressure shall be maintained
until completion of the ejection sequence and
descent to below pressure breathing altitudes if
ejection is initiated whilst it is applied.
1.4.2.16. Pressure breathing employing a The mean mask pressure and the pressure in the As in Clause 1.4.2.15 above, these partial
pressure helmet with counter pressure to the trunk counter pressure garments shall be within pressure assemblies employ either a pressure
trunk, to the lower limbs and in some applications the limits 18.8 to 20.0 kPa absolute (141 to 150 jerkin or a pressure waistcoat and the abdominal
to the upper limbs shall be used to provide mmHg absolute) at all altitudes above 40,000 bladder of G trousers to provide respiratory
protection against hypoxia at altitudes between feet. The limits for the changes of mask pressure counter pressure.
40,000 and 100,000 feet. during the respiratory cycle, the differences of
pressure between mask cavity and the bladder of
the respiratory counter pressure garment, the
rate of inflation of the counter pressure garments
and the pressure in the G trousers are identical
to those specified in Clause 1.4.2.15 above.
1.4.2.17. For all types of pressure breathing The mean mask pressure over the respiratory
system, deviation of mean mask pressure from its cycle during pressure breathing at altitudes over
nominal value shall be minimised. 40,000 feet shall be within 0.27 kPa (2.0 mmHg)
of its nominal value.
1.4.2.18. Pressure breathing with a mask The required relationship between the mean Pressure breathing with respiratory counter-
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together with counter pressure applied to the trunk mask cavity pressure and +Gz acceleration in a pressure to the chest and counter-pressure to the
and lower limbs can be used to provide protection breathing system providing PBG at accelerations abdomen and lower limbs by G trousers is
against the adverse effects of +Gz accelerations up to +9 G is for the mask pressure to increase employed to enhance the tolerance of aircrew to
of between +4 and +9 G. The mean pressure in linearly with increasing acceleration from safety sustained +Gz accelerations (PBG). The G
the mask cavity and counter pressure garment(s) pressure (0.1 to 1.0 kPag) at +4 G to 8.0 / 8.7 trousers must be inflated to the appropriate
shall increase linearly with acceleration in this kPag at +9 G. The total change of mask pressure pressure whenever pressure breathing is
range so as to maintain adequate blood pressure during the respiratory cycle shall not exceed the provided. Typically, this is achieved by inflating
in the head. limits specified in TABLE 1 above in Clause the G trousers in direct response to +Gz
1.4.2.13 and the difference between the mask acceleration by means of a moving mass in a
pressure and the pressure in the respiratory pneumatic Anti-G Valve (AGV) supplying air to
Counter pressure garment shall not exceed 0.5 the trousers (although electronically controlled
kPa. systems using an accelerometer are also
The rate of inflation of the respiratory counter available), and using the pressure in the G
pressure garment is to be such that the pressure trousers to control PBG by means of a pneumatic
in the mask cavity and garment do not lag more or electronic signal.
than 0.5 seconds behind the rise of pressure in The bladder of the respiratory counter pressure
the G trousers. When the level of +Gz garment is connected into the hose between the
acceleration is reduced the pressures in the pressure demand regulator and the mask so that
mask cavity and garment are not to lag more the garment is inflated to the PBG pressure.
than 0.5 seconds behind the fall in pressure in In the UK systems using a full-coverage lower
the G trousers. limb protection garment, the latter is typically
See also Clause 1.4.2.5 inflated at +2g to 10 kPag with a pressure rising
linearly with acceleration to 70 kPag at 9G
1.4.2.19. Mask cavity pressures in excess of The rise of mask cavity pressure induced by The performance of a breathing system on rapid
the requirements above (which may be induced by realistic head movement (mask hose pumping) or decompression is to be initially tested using a
short-term effects, transient events or equipment the maximum rate of climb of the aircraft shall not suitable lung simulator and (if appropriate for the
failure) shall be limited to levels which do not exceed 0.25 kPa (1.0 in wg). application) decompression times down to 0.1 s.
cause distress or discomfort to the crew member. The transient increase of mask pressure which is Refer to LEAFLET 3 and Flying Personnel
produced by rapid decompression of the cabin in Research Committee Report No. 1150 (1961).
0.1 seconds shall be less than 13.3 kPag (53.4 in
wg)and shall not exceed 5.5 kPag (22.0 in wg)
for longer than 50 ms.
The mask cavity pressure produced by the
regulator demand valve failing in the open
position shall not exceed 5.5 kPag (22 in wg).
1.4.2.20. There shall be minimal oscillatory The double amplitude of any oscillation of Mechanical tests of the oscillatory behaviour of a
activity in the mask cavity when the oxygen pressure within the mask cavity that lasts 0.25 breathing gas system are made by the sudden
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system is used – such activity is disconcerting for seconds or longer shall not exceed 0.06 kPa application of various flow rates to the complete
the crew member, may induce hyperventilation (0.25 in wg). system through several different levels of
and can interfere with verbal communications. resistance that simulate the range of impedances
of the human respiratory tract to oscillatory
behaviour.
Test of the absence of unacceptable oscillatory
behaviour must be made using human subjects.
The subjects impose a variety of respiratory
demands (rest, and light exercise with and without
speech) on the complete system (or at a minimum
the breathing gas demand regulator, low pressure
delivery system and oronasal mask) at ground
level and at selected altitudes.
Refer to LEAFLET 3 Para 3.9 and Clause
1.4.7.11 below.
Pressure Breathing Mask Seal
1.4.2.21. The mask-to-face seal of an The oronasal mask shall seal to the face both at The donning times and sealing properties of the
oronasal mask used for the delivery of pressure altitudes above 40,000 feet and on exposure to oronasal mask are to be established using human
breathing (PBA and/or PBG) shall be sufficient to +Gz accelerations of up to +9 G such that the subjects who have been trained in the use of the
minimise outward leakage so as not to distract / outward leakage does not exceed 0.1 L.s-1 equipment. Assessments are to be performed in
discomfort the crew member and waste breathing ATPD at a mask pressure of 4 kPag (30 mmHg) tests at ground level, on rapid decompression in a
gas. and 0.25 L.s-1 ATPD at a mask pressure of 9.3 Hypobaric Chamber and on exposure to +Gz
kPag (70 mmHg). accelerations in a man-carrying centrifuge.
1.4.2.22. The required standard of mask seal When the mask is already in place on the face Subjects with a variety of facial shapes and sizes
shall be achieved rapidly at the onset of pressure the required standard of mask seal shall be that span those of the aircrew population are to
breathing (PBA and/or PBG) so that the maximum achieved either automatically or by a simple be employed in these assessments. Where rapid
protection may be afforded to the crew member. manual operation. donning of the mask is required in the event of an
With automatic mask tensioning the specific emergency in flight, realistic tests of mask
standard of mask seal is to be achieved within donning times are to be conducted at each crew
0.5 seconds of the increase in mask pressure. station of the aircraft in which the equipment is to
When manual mask tightening is employed the be employed.
increase in the tension should be applied within 2 System configurations requiring mask donning in
seconds of the increase in mask pressure. the event of an emergency are normally only used
When the mask is not worn routinely and only with pressure breathing levels up to 4 kPag (30
donned in the event of an emergency the storage mmHg). Where higher levels are potentially used
and method of attachment of the mask to the the mask is worn at all times.
headgear (headset or helmet) shall be such that Refer to LEAFLET 3 and ASCC documents
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the mask can be placed on the face within 5 61/101/2B and 61/101/5A.
seconds of the commencement of an emergency
such as decompression of the cabin. The mask
shall be attached to the headgear and the
specified standard of seal achieved within a
further 5 s.
Press–to-test Facility
1.4.2.23. A facility whereby pressure Systems without counter pressure Garments The performance of this facility is to be such that
breathing may be obtained by the operation of a Operation of the regulator press-to-test facility the user can perform several respiratory cycles
manual control shall be provided, together with shall raise the mean mask cavity pressure to with the mask pressure raised.
indication of zero or continuous flow, to enable the within the limit +3.5 to +4.5 kPa. The total change Consideration shall be given to the provision of
user to test the integrity of the breathing gas and of mask cavity pressure during the respiratory pressure sensors in the mask supply system to
the pressure breathing system up to and including cycle should not exceed 0.75 kPa when the peak verify that the required pressure is achieved
the mask-to-face seal. inspiratory and expiratory flows are 0.5 l/s ATPD. during test.
Where either manual, or automatic, mask
Systems with counter pressure garments tensioning is provided to allow sealing under the
A single press-to-test facility shall be provided to maximum mask cavity pressure, consideration
inflate the anti-G trousers and raise the mask shall be given to activating the tensioning system
cavity pressure. during press-to-test.
Operation of the regulator press-to-test facility A ratio of G trouser pressure to breathing
shall raise the mask cavity pressure to a value pressure of 1:1 to 1.5:1 is required for use with a
not exceeding the maximum mask cavity chest counter pressure garment.
pressure at the highest aircraft operating altitude. The presence of a flow indicator (typically a
The magnitude of the mask pressure produced blinker showing white for flow and black for no
by operation of press-to-test when respiratory flow) allows for leak detection in the components
counter pressure garments are worn should not downstream of the regulator (including the mask)
in any case exceed 7.5 kPag (30 in wg). during press-to-test with breath held, and gives
The total change of mask cavity pressure during visual indication of normal regulator function
the respiratory cycle should not exceed 0.75 kPa during the breathing cycle.
when the peak inspiratory and expiratory flows
are 0.5 l/s ATPD.
The pressure within the anti-G trousers should be
raised to 1.5 to 2.0 times the breathing pressure
in the mask.
The press-to-test facility shall increase the mask
cavity pressure and anti-G pressure
progressively to the maximum specified.
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The press-to-test facility shall be capable of
being manually operated by the aircrew when
seated in the ejection seat with the harness fitted,
fully locked, and tightened.
Allowable Further Increases of Mask Pressure
1.4.2.24. During routine and emergency Any rise in rise of mask pressure including an See LEAFLET 3, Further increases of Mask
conditions, any rises in the pressure in the mask increase produced by a high continuous flow Pressure, Para 3.7
cavity above the values seen during breathing in failure of a demand valve should not exceed 5.5
the steady state shall be kept to a minimum. kPag (22 in wg).
1.4.3. Common System Requirements
Demand Regulators – General Requirements
1.4.3.1. The regulator shall supply The high pressure of the supply gas shall be The overall design should be tailored to the
physiologically acceptable breathing gas to the controlled by the demand regulator to provide a position where the demand regulator is mounted
mask in response to the reduction in mask flow of gas at pressures that comply with within the aircraft system. The three basic
pressure on inhalation. physiological objectives of Table I at Clause mounting positions are:
1.4.2.13, in response to inhalation demand by the
user. (a) on the ejection seat.

(b) on the side console.

(c) worn by the aircrew.

1.4.3.2. The regulator shall provide a The regulator shall provide a positive pressure It is desirable for safety pressure to be present at
means to eliminate inward leakage of ambient air (safety pressure) such that the pressure in the all cabin altitudes for protection against smoke
into the mask and its connecting tube at cabin mask cavity does not fall below cabin ambient and fumes. A typical characteristic maintains the
altitudes above 12,000 feet. pressure at peak inspiratory flow rates of up to mean mask cavity pressure between 0.25 and
1.2 L/s (72 L/min) ATPD at cabin altitudes above 0.50 kPag (1.0 and 2.0 in wg) at cabin altitudes
12,000 feet (and preferably at all other altitudes). up to 38,000 feet.
If safety pressure above 12,000 feet is provided
automatically there shall also be a means of
selecting it manually at other cabin altitudes.
Demand Regulators – Specific Requirements
1.4.3.3. Specific additional physiological or Specific mission and system requirements shall Discussion with the Project Team Leader,
system compatibility requirements shall be met by be analyzed to determine whether any of the Defence Staff and aeromedical specialists may be
incorporating into the design one or more of the following features are necessary. required to determine the necessity or otherwise
features described in the following Clauses. of the features described below.
Air Dilution

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1.4.3.4. The oxygen concentration of the This type of regulator (airmix) shall progressively The correct ratio of cabin air to oxygen may be
gas supplied to the aircrew shall be controlled to reduce the air dilution of the 100% oxygen supply achieved automatically by use of a barometric
levels physiologically appropriate to the cabin gas with increasing cabin altitude. Undiluted aneroid capsule.
altitude by means of dilution of a 100% oxygen 100% oxygen shall be delivered at cabin altitudes To control the high supply pressures associated
supply with cabin-ambient air. A means of of 34,000 feet and above. with supplies from stored oxygen sources
overriding the dilution by manually selecting 100% Manual selection of 100% oxygen supply shall (gaseous or liquid), regulators of this type may
oxygen supply shall be provided. prevent any ingress of air to the breathing gas. require an inbuilt pressure reducing valve (PRV)
upstream of the demand valve. Such multi-stage
devices can operate over a wide range of supply
pressures.
In addition to providing oxygen when required the
manual selection facility allows smoke and fumes
in the cabin to be excluded from the breathing
gas.
Pressure Breathing with Altitude (PBA)
1.4.3.5. The alveolar partial pressure of The pressure in the mask cavity shall be in The correct pressure schedule may be achieved
oxygen shall maintained at physiologically accordance with Clause 1.4.2.18. for cabin automatically by use of a barometric aneroid
acceptable levels by increasing the pressure in the altitudes above 38,000 feet. capsule.
mask cavity linearly with fall of cabin ambient
pressure for cabin altitudes above 38,000 feet.
Pressure Breathing with G (PBG)
1.4.3.6. The mean pressure in the mask The pressure in the mask cavity and counter When PBA and PBG conditions occur
cavity and counter pressure garment(s) shall pressure garment(s) shall be in accordance with simultaneously, the higher of the two pressure
increase linearly with +Gz acceleration in the Clause 1.4.2.18 at +Gz accelerations above +4 schedule requirements shall be maintained in the
range +4 to +9 G so as to maintain adequate G. mask cavity.
blood pressure in the head.
Compatibility with MSOC (OBOG) Supplied Systems
1.4.3.7. If used in a MSOC-supplied The regulator shall meet the requirements of A single stage demand or pressure demand
(OBOG) system the regulator shall be capable of Clause 1.4.4.6 at the relatively low outlet regulator may satisfy requirements for MSOC
meeting physiological requirements at the typical pressures generated by a MSOC. (OBOG) systems. Distribution/delivery pipe back
low supply pressures generated by MSOCs. pressures are naturally more critical for low
pressure systems.
System Installation Requirements
1.4.3.8. The need for filters shall be Where required efficient filters, designed to Where filters are fitted they should be readily
minimised by design. minimise pressure losses (especially in MSOC accessible to permit replacement and inspection.
systems) shall be provided to prevent the ingress

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of foreign matter. Filter elements shall be readily
removable for cleaning or replacement.
1.4.3.9. In systems employing an MSOC, a Test shall be made of the filter element and of Filters should be readily accessible to permit
particulate filter shall be incorporated downstream the MSOC package to demonstrate compliance. replacement and inspection.
of the molecular sieve beds. The filter element shall be rated at no greater
than 1 micron (< 1 μm).
1.4.3.10. The mountings shall be able to The acceleration and the factors of safety to be
withstand all loads induced by the acceleration the applied shall be determined from design data
aircraft is capable of generating with a factor of stated for a particular project by the airframe
safety. constructor. Pipes connected to valved cylinders
shall be arranged to minimise the strain on
soldered joints should the nipple not line up
correctly with the valve outlet connection.
1.4.3.11. Pressure losses shall be Tests shall be made to establish the satisfactory Guiding principles on the design of oxygen
minimised. technical operation of the system (see also MAP systems to minimise pressure losses are given in
RA 5200 series). Tests, based on guidance from LEAFLET 4. A recommended schedule of tests
the Royal Air Force Centre of Aviation Medicine for liquid oxygen system is given in LEAFLET 5.
(RAF CAM) shall also be made to ensure the
satisfactory physiological operation of the
system.
System Warnings
1.4.3.12. Warnings and indications shall be (a) A means shall be provided in the supply The warning system should where possible or
provided at the crew stations. line to the regulator to give warning of a failure in feasible be provided with a sensor or sensors to
The warning of oxygen failure shall be provided on a gaseous oxygen system. monitor breathing gas quality.
the Standard Warning System as required by Def The oxygen flow indicating device may be the
Stan 00- 970 Section 4, Clause 4.19.59. (b) An indication of the presence or absence “blinker” on panel mounted demand oxygen
There shall be a positive indication, visible to each of breathing flow shall be displayed to each user. demand regulators or an independently mounted
crew member, of oxygen flow to his/her mask. “blinker” when other forms of demand oxygen
An indication of regulator inlet pressure or a low (c) The contents of back-up and/or regulators are used.
pressure warning device shall be provided, visible emergency oxygen systems shall be displayed. An oxygen flow indicating device may not be
to each crew member in single and two seat required on portable equipment for emergency
aeroplanes and (as a minimum) to selected crew (d) MSOC systems shall be provided with a use (e.g. fire-fighting).
members in multi-seat aeroplanes. low oxygen partial pressure sensing device or Individual low pressure warnings may not be
It shall be possible for selected crew members to equivalent oxygen monitor. required in systems where all the breathing gas is
determine the combined contents of the gaseous derived from a common source (e.g. an MSOC).
oxygen storage cylinder or the contents of each (e) The oxygen monitor shall provide a status In this instance a single system pressure warning

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liquid oxygen converter at any time during the signal to the aircraft. may be sufficient.
flight. The low oxygen warning level shall be set to
ensure that the aircrew alveolar tension does not
fall below the values specified in section 1.4.2
1.4.3.13. Pre-flight Test Verification of breathing and anti-g gas
It shall be possible for the pilot to check the connections to the pilot should be conducted
integrity of connections and the absence of leaks, prior to take-off by means of activating the press-
as well as the functioning of the system on the to-test. In MSOC systems the oxygen monitor
ground before take-off. shall conduct necessary built-in-test functions to
Standby oxygen supplies (back-up oxygen and verify correct operation.
emergency oxygen) shall not be used for ground
testing.
1.4.4. Breathing Gas Sources
Stored Gaseous (GOX) or Liquid (LOX) Oxygen
Quantity Required

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1.4.4.1. The quantity of oxygen carried shall The quantity provided shall be sufficient to meet The rate of consumption of stored 100% oxygen
be sufficient to provide a mixture of breathing air the requirements under the more severe of the may be calculated from the appropriate curves of
and oxygen of acceptable quality, or 100% following conditions: the oxygen requirements for aircrew graph, in
oxygen, as required by the cabin altitude and the which allowance has been made for:
breathing demands of each crewmember. (a) The maximum endurance of the
The equipment shall be capable of supplying the aircraft in the most critical case (for oxygen)  Type of aircraft
crew member(s) with breathing gas of appropriate in the roles specified, including any role in  Number of aircrew
composition and quality, at satisfactory pressure which the aircraft is refuelled in flight.  Mask leakage
and flow, under all flight conditions and at all  Variations in work rate
altitudes to which the aircraft is required to (b) A return to base after loss of cabin  Variations in regulators
operate. pressure at half the maximum range of the  Individual variation.
The quality and composition of the breathing gas aircraft.
in the above requirements shall be determined by In both cases, account shall be taken of the The rate of consumption of 100% oxygen for use
reference to Clause 1.4.5.4 possible need to provide increased with an air-mix regulator is shown for a typical
Physiological Objectives. breathing pressure automatically to the panel mounted air-mix regulator in curves 4, 5,
Calculations covering the amount of stored mask in the event of a loss of cabin and 6.
oxygen shall be submitted to the Project Team pressure at high altitude and to inflate As the Defence Staff may require the continuous
and agreed prior to final storage vessel selection. pressure garments if worn. use of 100% oxygen, the quantity of oxygen
required should be calculated for both the use of
If the critical case is that of a ferry role, the air-mix and of 100% oxygen and the results
additional oxygen required beyond that for the reported to the Project Team Leader so that a
operational role might be carried in readily decision may be reached in discussion with the
removable containers for which stowage shall be Defence Staff.
provided.

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OXYGEN REQUIREMENTS FOR AIRCREW

1.4.4.2. The number of oxygen containers The quantity of oxygen available from standard High Pressure Cylinders are normally used for
to be provided shall be calculated based on the containers and liquid oxygen converters is given primary, backup or emergency gas sources and
capacities of standard containers. in the tables below. typically store gas at pressures of 12,410 to 3,790
Where crew and passengers draw from the same kPag (1,800 to 2,000 psig).
supply, a quantity shall be reserved exclusively for CAPACITY OF STANDARD CONTAINERS Low Pressure Cylinders are sometimes used in
the crew. Ideally crew and passenger supplies Note: Normal Temperature and Pressure (NTP) portable oxygen sets and typically store gas at
should be independent. is 15°C and 101.325 kPa absolute. pressures of 2,760 to 3,450 kPag (400 to 500
If two or more gaseous oxygen cylinders (or psig).
converters) are carried, the system shall be so a. Gaseous Oxygen
designed that no single failure in an oxygen Nominal cylinder Litres of gas
cylinder, its attached parts and its immediately size (litres) available at NTP
adjacent delivery lines or manifold will lead to loss 70 66

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of more than half the remaining oxygen. See also 200 190
Clause 1.4.4.6 below. 400 385
In multi-crew aircraft, the remaining cylinder(s) 750 710
shall be capable of supplying oxygen at the 1400 1330
maximum flow demand to all crewmembers for a 2250 2130
minimum period of time specified by the Project
Team.
Notes:
1. Allowance has been made for the
minimum bottle capacity.
2. The quantity of oxygen available
applies to a bottle charged at 15°C.

b. Liquid Oxygen
Available gaseous Yield (litres NTP)

Capacity Immediately 12 24
(Litres) (10 mins) hours hours
after Filling After After
Filling Filling
3.5 2800 2660 2520
5 4000 3800 3600
10 8000 7600 7200
25 20000 19000 18000

Notes:
1. The available yields are based
upon a conversion factor of 800 litres (NTP)
gas per 1 litre liquid oxygen.
This figure incorporates a safety margin for
gauging and filling and also for the gas
unavailable for use within stabilisation
systems.
2. 5 % is assumed to be lost in 12
hours and 10% in 24 hours.
Installation & Operation – General

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1.4.4.3. The equipment shall operate in The equipment shall support the use of inflatable
normal flight as defined by the mission roles of the garments where appropriate (See Clause
aircraft; during and after any in-flight emergency; 1.4.2.15) including those associated with
and as appropriate after escape from the aircraft. pressure breathing as protection against high
+Gz acceleration and/or altitude.
The installation shall also be compatible with and
support the use of nuclear, biological and
chemical (NBC) protective equipment and
clothing, as required by the aircraft specification.
1.4.4.4. The equipment shall be protected An over pressure relief device shall be connected
by an over pressure relief device. to the high pressure storage vessel to protect the
equipment from over pressurisation. The device
shall protect the equipment at all times,
especially during charging and at times when the
equipment is exposed to elevated temperatures.
The pressure at which an over pressure relief
device is activated shall not exceed the proof
pressure of the equipment that it is protecting.
Gas vented from an over pressure relief device
should not be vented into an enclosed bay.
Where equipment is mounted in an enclosed bay
the gas shall be vented overboard.
1.4.4.5. Charging procedures to account for The pressure to which a high pressure storage
ambient temperature and pressure. vessel is charged shall take account of the
ambient temperature and pressure at which it is
being charged and to which it will be subjected.
Compensation shall be made for variations in
temperature and pressure so that the vessel is
charged with the equivalent quantity of gas as it
would be at standard temperature and pressure
conditions.
1.4.4.6. Loss of gas shall be prevented in Where multiple cylinders or liquid oxygen
the event of localised damage or during servicing converters are employed, check valves and/or
operations. isolation valves shall be installed in the
distribution lines.
1.4.4.7. Provision shall be made for the A contents indicator and/or low-level warning For GOX systems, low level contents indication

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crew to be aware of low contents in the storage device shall be provided. The indication system may be provided by a pressure gauge (upstream
medium. shall not be altitude sensitive. of any pressure reducing valve) mounted directly
on the unit for applications where the unit is
visible in flight, or by:

(a) Remote pressure gauge visible to

the crewmember.

(b) An audible or visual warning from

the aircraft system provided with an input
from the unit by either a pressure switch or
pressure transducer.

For LOX systems, use of pressure gauges will not

provide indication of contents. A capacitance
probe in the liquid with an external electrical
bridge circuit can be successfully used to do so,
and to provide a signal to the aircraft system to
give an audible or visual warning of low contents.
1.4.4.8. Equipment mountings shall be able The acceleration and the factors of safety to be
to withstand all loads, including a factor of safety, applied shall be determined from design data
induced by the acceleration the aircraft is capable stated for a particular project by the airframe
of generating. constructor.
1.4.4.9. Mechanical and pneumatic Stresses shall be reduced by keeping the
stresses in high-pressure pipe work shall be diameter of high-pressure lines to the minimum
minimised by design. conducive with the required mass flows and
tolerable back-pressures.
Installation & Operation – GOX Specific
1.4.4.10. The installation shall operate All the specified physiological requirements shall The flow/pressure characteristic of pressure
effectively when depletion of contents has be met when the supply pressure at the storage reducing valves in the installation will be a
significantly reduced the cylinder supply pressure. cylinder has fallen to 1,551 kPa (225 psig), and controlling factor.
for emergency inflation of pressure garments
when the pressure has fallen to 2,068 kPa (300
psig).
1.4.4.11. Gaseous oxygen cylinders shall The position of the external charging valve shall Wherever possible, cylinder stowage shall be
comply with the requirements of Part 1 Section be such that exposure to contamination is arranged such that the cylinders are vertical, with

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6.14 and BS 4N 100. Valved cylinders shall be minimised and that access by service personnel the outlet uppermost, when the aircraft is in
fitted in all Naval and RAF Aircraft when three allows for easy and safe operation of the straight and level flight.
cylinders or fewer are carried. charging procedure without undue damage being Pipes connected to valved cylinders shall be
Adequate access shall be provided to enable caused. arranged to minimise the strain on soldered joints
valved cylinders to be changed without removing Compatibility with ground crew charging should the nipple not line up correctly with the
other items of equipment. This is to be achievable procedures should be verified by the Project valve outlet connection.
under all military operational environments without Team.
risk to the integrity of the system or the quality of
the breathing gas supplies.
Installation & Operation – LOX Specific
1.4.4.12. Provision shall be made for the
rapid stabilisation of liquid oxygen converters.
1.4.4.13. At least two independent means of i.e. a primary relief valve and a totally separate If the converter is installed in a closed aircraft bay
preventing a hazardous rise of pressure in the second relief valve or bursting disc. consideration shall be given to independently
converters shall be provided. The primary pressure relief valve shall be vented venting the secondary relief valve (or bursting
overboard. The primary and secondary relief disc) overboard.
valves shall be designed to minimise the
possibility of being rendered inoperative by ice
formation.
1.4.4.14. Overboard vents shall be designed The overboard vents shall be well away from fuel
so as to minimise ignition and ice accretion risks. vents and other outlets where a risk of ignition
could be increased.
Overboard vent outlets shall be so positioned to
minimise the possibility of ice formation in flight.
1.4.4.15. Provision shall be made to reduce Consideration shall also be given to locating
as far as possible the effects of condensation on oxygen connections and vents with their ports
the converter and pipelines, and for the safe downward to prevent water ingress when the
collection and disposal of condensed moisture. converter is not connected to the airframe
system.
Consideration shall be given to locating suitable
isolation valves within the aircraft system.
1.4.4.16. Isolation valves shall be installed in The Project Team shall define the position of
the distribution lines to address safety issues and isolation valves in the distribution lines to address
to facilitate servicing. safety issues and to facilitate servicing.
Gas Generating Systems
1.4.4.17. The breathing gas generating The requirements of Clauses 1.4.2.8 and 1.4.2.9 Gas generating systems for aircraft are generally

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system shall deliver a sufficient flow to meet the shall be met under all conditions of crew termed Onboard Oxygen Generator Systems
total breathing gas requirements of all aircrew workload (breathing demand) and engine throttle (OBOGS).
(and passengers, if applicable) at physiologically setting (which may affect the pressure and A Molecular Sieve Oxygen Concentrator (MSOC)
acceptable oxygen concentration levels, temperature of the pneumatic supply to the gas is a type of Onboard Oxygen Generator (OBOG)
irrespective of aircraft altitude or flight condition generating system). Pipe work downstream of that uses a molecular sieve to produce oxygen
(except for flight conditions for which a separate the MSOC shall be kept to a minimum and enriched breathing gas from conditioned engine
and independent supply of breathing gas has designed to minimise pressure losses. Higher air bleed air by a pressure swing adsorption
been specifically included). inlet pressures can be reduced by fitting a technique.
In certain aircraft installations the pressure at pressure regulating valve to the inlet of the MSOC product gas oxygen concentration varies
which air is supplied to the MSOC at the low MSOC. inversely with product demand flow. Generally,
engine power settings associated with idle Other (non-breathing) demands may have to be higher bleed air supply pressures and lower vent
descent and ground idle conditions may be met from the gas generating system (e.g. ambient pressure promote better MSOC
insufficient to provide a flow of product gas to an garment inflation, demist flows to visors). The performance. The maximum acceptable oxygen
adequate pressure. In these circumstances a mean load represented by such auxiliary concentration obtainable from the conventional
deterioration of performance with respect to demands shall be taken into account when pressure swing adsorption (PSA) MSOC is 94 %
Clause 1.4.2.8 may be acceptable. In the worst calculating the total flow required from the gas with the balance comprising argon.
case (engine idle setting) the breathing system generating system (which dictates the size and/or The philosophy employed in the UK is to use a
shall be capable of meeting individual pulmonary number of gas generation modules in the backup supply of 100% oxygen at cockpit
ventilations of at least 25 L.min-1 ATPD and peak design). altitudes of 27,000 feet or above.
inspiratory demands of at least 90 L.min-1 ATPD To meet the minimum performance requirement The performance of the MSOC may be assessed
with the total change of mask pressure not at engine idle it will be necessary to advise the and verified during build by drawing steady
exceeding 1.5 kPa (6”wg). airframe and engine manufacturers of the product demand flows.
minimum required pressure of the engine bleed
air supply.
1.4.4.18. Aircrew (and passengers, if A back-up supply of gaseous oxygen shall be The back-up source shall be of sufficient quantity
applicable) shall be physiologically protected in provided for use during temporary loss of supply to meet the breathing demands of aircraft
the event of a failure in the MSOC or its air supply. pressure to the MSOC or failure of the MSOC to personnel under representative workloads during
provide physiologically acceptable breathing gas a descent to a cabin altitude at which it is
as required by Clause 1.4.2.8 physiologically acceptable to breathe air at
ambient pressure (typically 10,000 feet).
1.4.4.19. Backpressure in the vent path of If vent ducts are required they shall be sized and This requirement applies in particular to systems
MSOC beds shall be minimised. routed to offer minimal impedance to vent flows, in which vent ducts are fitted to the MSOC. In
such that the MSOC vent port references true many installations the MSOC is located in an
aircraft ambient pressure. unpressurised bay and the beds may be freely
vented into the compartment.
When the on-line bed becomes saturated with
nitrogen, the bleed air is switched to another bed
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to ensure a continuous delivery of product gas.
The off-line bed is then rapidly de-pressurised to
aircraft ambient pressure and the regeneration
step completed by flowing a proportion of the
product gas back through the vented bed. If a
vent duct is necessary to pipe vent gas to aircraft
altitude, then any backpressure will reduce MSOC
performance.
1.4.4.20. The installation shall be designed Consideration shall be given to the installation of An MSOC system acts to remove contaminants
such that contamination (e.g. by oil and fuel filters to prevent particulate matter (a) entering which may be present, not least because ~80% of
vapours or particulate matter) in the compressed the MSOC, and (b) being carried to the breathing the supply air is vented overboard as nitrogen-rich
air (engine bleed air) supply to the MSOC is mask(s). exhaust, but also because the molecular sieve
minimised and preferably eliminated. It will be necessary to advise the airframe and beds act as efficient filters. However, in “filtering”
engine manufacturers of the need to provide a any contaminants from the gas stream the
clean, uncontaminated engine bleed air supply. efficiency of the MSOC in terms of oxygen
enrichment (nitrogen removal) may be adversely
affected. Also, after in-service exposure to a
harsh operating environment there is a risk of
small particles of sieve material being carried in
the breathing gas unless outlet filtering is present
in the MSOC unit.
1.4.4.21. The installation shall be designed A water separator/extractor device shall be The molecular sieve materials used in MSOC
such that the frequency and quantity of free water installed if analysis of the operating envelope of applications typically have an affinity with
entering an MSOC is minimised and preferably the system indicates that free water is likely to substances other than nitrogen, water being
eliminated. evolve at the MSOC inlet. The system shall be principal amongst these (indeed such sieves are
designed such that the operating temperature also used in desiccant applications). The
and pressure of the water separator/extractor contaminant substances block the receptor sites
relative to those of the MSOC itself are in the molecular lattice, thus reducing the sieve’s
conducive to the extraction of water from the effectiveness at nitrogen adsorption.
MSOC air supply. It is also highly desirable to minimise the quantity
The bleed air or other air supply system shall be of water vapour in the supply gas to the MSOC
fitted with adequate and sufficient drains to because of the possibility of condensation under
prevent the collection of condensate in the pipe certain temperature conditions (see Clause
work or other parts of the supply system. 1.4.4.22 below).
Consideration shall be given to the provision of at
least one Ubend in the supply pipe work with a
drain at its lowest point.
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1.4.4.22. The installation shall be designed The climate/environment in which the aircraft will It is preferable that during operation the MSOC
such that temperature differences between the operate shall be considered and if necessary temperature is kept stable, and not allowed to
MSOC beds and the ambient environment do not heating or lagging of the MSOC or its cool below the ambient or air supply temperature.
give rise to water condensation in the beds. compartment shall form part of the installation. MSOC performance is generally optimised at an
air supply and ambient temperature of 20°C.
Performance can be expected to change with
decreasing or increasing temperatures from this
point.
See also the requirement in Clause 1.4.2.11
1.4.4.23. Air Management. An interface control document shall be generated The Project Team should ensure this interface is
The management of the air supply to the MSOC between the air management and the MSOC defined and adhered to.
shall be such that all contaminants harmful to suppliers.
molecular sieve are minimised or eliminated.
1.4.5. Types of System
Primary Breathing Systems
Protection Against Hypoxia
1.4.5.1. A breathing system shall: The breathing system shall comply with the The two general categories for oxygen equipment
physiological objectives of Clause 1.4.2.8 are fixed and portable. These two types of
(a) Prevent significant hypoxia whilst equipment can be of the continuous flow, diluter
the cabin is pressurised without inducing demand or pressure demand variety depending
acceleration atelectasis or delayed otitic on the application.
barotraumas. Continuous oxygen flow systems dilute the flow of
oxygen with cabin air at low cabin altitudes and
(b) Prevent significant hypoxia high respiratory demands. They generally provide
following decompression of the cabin to adequate protection up to 25,000 feet but may be
altitudes of up to the maximum cabin altitude used as emergency protection for brief exposures
that can occur in flight. (The cabin pressure to altitudes as high as 43,000 feet typically
altitude can be higher than that of the following the decompression of civil airliners.
aircraft due to aerodynamic suction following Demand flow equipment requires a pressurised
a defect in the cabin structure). gas supply source (generally pure oxygen or
OBOG product gas). A slight reduction of
(c) Prevent, when required, greater pressure in the mask induces oxygen (breathing
than 2% admixture of cabin air with the gas gas) flow into the mask. A rise of mask pressure
from the aircraft oxygen store or onboard stops the flow of gas from the regulator. The
oxygen generating system by providing, oxygen can be diluted with the appropriate
when required, safety pressure in the mask proportions of air at all altitudes below 34,000

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cavity to prevent inward leakage of feet.
environmental air.
Delivery Pressure
1.4.5.2. A breathing system shall: The breathing system shall comply with the Pressure demand equipment is similar to demand
physiological objectives of Clause 1.4.2.10 flow equipment but provides additional protection
(a) Meet respiratory demands without above 38,000 to 40,000 feet by progressively
imposing excessive resistance to breathing. increasing the breathing delivery pressure of
100% oxygen.
(b) Provide increased pressure to the Pressure breathing is also used to provide
mask (and a gas supply for inflation of protection against the effects of +Gz acceleration
counter pressure garments where fitted) for at levels of +4 G or above.
high altitude protection when the cabin
altitude exceeds 38,000 to 40,000 feet and,
when selected manually, at ground level
(press-to-test facility).

(c) To prevent the pressure in the

mask, generated either by trapped gas on
rapid decompression or by a failure of the
demand mechanism, from exceeding
acceptable physiological limits.

(d) Not produce significant oscillations

of pressure within the mask cavity.

(e) To provide, where required,

pressure breathing as a means to enhance
tolerance to +Gz acceleration.
Minimal Dead Space
1.4.5.3. A breathing system shall not The dead space of the mask volume shall not
impose excessive rebreathing of expired gas. exceed 0.2 litres.
Breathing Gas Composition
1.4.5.4. A breathing system shall: The composition of the breathing gas shall be in This applies to all methods for the generating the
accordance with one or more of the following breathing gas supplies, including pre-charged
(a) Provide pure breathing gas standards: gaseous or liquid oxygen storage systems or
supplies which are free of any hazardous or Gaseous and Liquid oxygen systems: onboard oxygen generator (OBOG) type systems.

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odorous chemicals or compounds under all STANAG 7106 Characteristics of Gaseous In the former cases, this requires careful control
conditions of use within the normal operating Breathing Oxygen, Liquid Breathing Oxygen and and monitoring of the process used to charge the
envelope of the aircraft. The hazard may Supply Pressures, Hoses and Replenishments system: in the latter case this means that typical
arise either from exposure to the chemical or Couplings: and untypical chemicals or compounds in the air
compound or from the effects of the STANAG 7187 On Board Oxygen Generating supply to the OBOG should not degrade the
chemical on the ability of the breathing Systems (OBOGS) Performance Standards. ability of the system to produce pure breathing
system to function correctly, over both short MSOC systems: gas of an acceptable oxygen concentration over
and long term periods of operation. ADV-pub-61/101/10 The Minimum Quality the lifetime of the breathing system.
Examples of chemicals or compounds that Criteria for On-Board Generated Oxygen. Filter devices placed in the breathing gas supply
may produce a hazard to the operator of the Verification shall be by testing in a manner have a finite capacity for contaminants, and
breathing gas supply or to the breathing representative of the way in which the aircraft failure to cleanse the breathing gas may not be
system includes fuel vapours, other and breathing system are operated. detectable other than by failure of the breathing
hydrocarbon species and water vapour. In peacetime use, breathing gas of acceptable gas system itself or the exposure of the user to
quality shall be provided without the need for hazardous or odorous chemicals or compounds.
(b) Not be degenerated by exposure, additional filtration systems such as carbon The capability to provide a secure source of
either over short or longer term periods of based devices placed in the breathing gas breathing gas supplies, and gas supplies for other
usage, to typical or untypical chemicals or supply. purposes associated with life support is of
compounds which might be present in the In wartime use, where other types of chemical or fundamental importance, and where possible or
environment where the breathing gas other hazards may be present in the feasible a means to monitor the quality of the
system is used, or where gas supplies are environment, the incorporation of a lifed filter breathing gas used for other life support or safety
derived, including all associated pipe work, element may be considered if the breathing gas critical functions should be considered.
connectors, regulators and supporting system cannot provide supplies which are at all
equipment. times free of any of these other types of chemical
or other hazards.
NBC Equipment Compatibility
1.4.5.5. When the aircraft specification The system shall provide breathing gas of an This applies to all methods for generating gas
requires protection against NBC hazards the acceptable quality when exposed to appropriate supplies, including pre-charged gaseous or liquid
installation shall be compatible with and support NBC hazards. Verification by analogy may be oxygen storage system or onboard oxygen
the use of NBC protective equipment and clothing, acceptable where the design system is similar. generator (OBOG) type systems. The system
as defined in the aircraft specification. Guidance criteria for acceptability shall be given should be capable of providing breathing and
The installation shall be capable of preventing the by the Project Team Leader. other relevant gas supplies which are free of NBC
passage or ingress of NBC hazards in any hazards in either liquid, gaseous, vapour or
unacceptable concentration into the breathing gas aerosol forms, preferably without the fitment of
supplies or into any gas supplies which are additional filtration devices to the breathing
provided for other life support or safety critical equipment or system.
functions, such as respirator demist supplies, The system may form part of an integrated NBC
including in the event of limited equipment failure. defensive system (Def Stan 08-11). Any specific
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requirements for the interfacing of the oxygen
system and its functionality into the NBC
defensive system shall be defined by the Project
Team Leader.
Operational Circumstances
1.4.5.6. The breathing system shall operate The breathing system shall comply with the
under the following circumstances: physiological objectives of Clauses 1.4.2.8 to
1.4.2.10 and Clause 1.4.4.3
(a) Normal flight,

(b) During and after any in-flight

emergency.

(c) Emergency descent whilst in the

aircraft.

(d) As appropriate after escape from

the aircraft.
Garment Compatibility
1.4.5.7. The breathing system shall support The breathing system shall comply with the See LEAFLET 3.
the use of inflatable garments where appropriate, physiological objectives of paragraph 6 of Leaflet
such as those associated with pressure breathing 3.
as protection against high +Gz acceleration and /
or high altitude.
Pressure Losses
1.4.5.8. Pressure losses in aircraft pipe Pressure loss calculations shall be supplied and See LEAFLET 4. Coordination shall be the
work shall be minimised. agreed prior to aircraft layout, and verified by test responsibility of the Project Team Leader.
(ICD required).
Materials
1.4.5.9. The materials used in the Non metallic materials shall comply with BS Particular care should be taken in the selection
construction of the system shall not be liable to 4N100. and usage of non-metallic materials, which should
ignition, attack or corrosion by contact with pure be suitably tested to demonstrate their
oxygen or oxygen enriched gas. compatibility with oxygen at representative
Materials used shall not produce toxic gases or working pressures and temperatures.
noticeable odours.
Materials and finishes used shall not cause

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dermatitis or other skin irritation.
1.4.5.10. When the aircraft specification The system shall resist the effects associated This applies to the complete breathing gas
requires protection against NBC hazards the with contamination and be capable of system including pipe work, regulators and other
materials used in the construction shall resist the decontamination. devices associated with the function of the system
potentially damaging and degrading effects which Verification of NBC compatibility shall be by which may feasibly be exposed to such hazards.
are associated with certain types of NBC hazards. controlled exposure to appropriate NBC hazards The equipment should be chemically hardened.
or as directed by the Project Team Leader. Consideration should be given to providing an
ability to maintain the equipment where hazards
are present.
Layout
1.4.5.11. The system shall be so arranged The breathing system shall comply with the
that local changes of pressure within the cabin for physiological objectives of Clause 1.4.2.1
any reason (e.g. loss of cabin canopy) shall not
cause malfunction of the system itself or
associated equipment.
1.4.5.12. The location of gas cylinders and The airframe shall be designed where practicable Def Stan 08-41 and Def Stan 08-11 provide useful
molecular sieve oxygen concentrators, and their to locate the gas cylinders and molecular sieve guidance. Any aircraft bays containing oxygen
associated pipe work, shall be such as to afford oxygen concentrators, and their associated pipe equipment (including MSOC type systems) should
the best protection possible from enemy action work, to afford the best protection possible from ideally be dry bays which are designed to prevent
including where NBC hazards are employed. enemy action including NBC hazards. the ingress of contamination. These bays should
Consideration shall also be given to the location of ideally also be designed to be free draining and
components to minimise the likelihood of damage compatible with decontamination using hot water
during a survivable crash. and detergent.
Icing Precautions
1.4.5.13. The system shall be constructed Environmental testing of the equipment shall Particular care is needed with regard to filters,
such that internal or external ice accretion will not ensure that any ice formation within the ports and moving parts of reducing valves.
seriously affect performance. equipment does not seriously affect its
performance.
Low points within aircraft pipe work that may trap
water shall be provided with a means of
drainage.
Freedom from Leaks
1.4.5.14. With the exception of designed Leakages from individual items of equipment When non-operational, the system integrity shall
leakages for functional purposes the system shall shall be minimised, shall not significantly degrade be such that no diffusion or permeation of
be permanently gas tight at its working pressures system performance and shall not allow the hazardous chemicals or compounds into the
over the full temperature range (operating and ingress of any external contaminants. equipment will take place.

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non-operating).
Servicing
1.4.5.15. Recharging the main gaseous and Recharging shall not take longer than five During charging operations the rate at which a
liquid oxygen systems shall be from a point or minutes unless otherwise agreed with the Project cylinder is charged shall not cause the
points outside the aircraft. The recharging point(s) Team Leader. temperature of the cylinder to exceed 40°C
shall be protected from rain, and the ingress of The position of the external charging valve shall (104°F)
any harmful contaminants, and not cause be such that exposure to contamination is
interference with other servicing. The recharging minimised and that access by service personnel
points shall be designed to allow safe recharging allows for easy and safe operation of the
in the presence of NBC hazards by personnel charging procedure without undue damage being
wearing individual protective equipment. caused.
Compatibility with ground crew charging
procedures shall be verified by the Project Team.
The charging connection shall comply with
STANAG 7106.
1.4.5.16. The risk of overcharging shall be A clear indication of system contents shall be Overpressure or spillage hazards should be
minimised. available to the servicing crew at the point of avoided.
charging.
1.4.5.17. Dismantling of system components Charging of back-up and emergency oxygen Every disconnection/reconnection operation is a
during servicing operations shall be minimised. cylinders in situ shall be possible without the potential source of leakage, contamination or
removal of any equipment. The replacement of malfunction.
MSOC service items (e.g. filters) should be
possible without the removal of the MSOC from
the aircraft.
Backup Oxygen Systems (BOS)
1.4.5.18. A back-up supply of gaseous 100% The BOS shall supply the aircrew with oxygen in The emergency and back-up oxygen supply
breathing oxygen shall be provided for use during: the event that the primary breathing system is requirement may be satisfied by a single oxygen
malfunctioning or unavailable, or in the event of supply mounted on the ejection seat; otherwise
(a) Loss of supply pressure to the cabin decompression, and shall comply with the the BOS may be mounted on the airframe.
MSOC. physiological requirements of Table 1 at Clause The main requirement for the design of a BOS is
1.4.2.13 to provide a source of 100% breathing oxygen
(b) Failure of the MSOC to deliver during an in flight failure of the MSOC.
product gas. The longest anticipated duration of BOS supply
shall include the time taken following a loss of
(c) Partial pressure of oxygen in the engine power at high altitude in a glide descent,
product gas falling below an acceptable and for attempts at re-start. The stored breathing

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level. oxygen shall meet the requirements of STANAG
7106. A warning system shall alert the
(d) In the event of decompression of crewmember when a minimum volume of oxygen
the cabin. remains. With a minimum of oxygen remaining
the aircraft should immediately descend to a
cabin altitude not exceeding 10,000 ft.
1.4.5.19. The BOS shall be of sufficient The BOS shall be shown to have sufficient stored The volume of oxygen required for the defined
quantity to protect the crew members for the oxygen capacity to meet the specified worst-case worst-case scenario shall be calculated at an
longest anticipated duration of MSOC failure; that crewmember usage scenario by calculations early stage of the project.
is, sufficient for the time taken following a loss of during the system design stage. The calculations shall take into account the
engine power at high altitude to glide descent, and Aircraft system testing in an altitude chamber maximum time required to remain at specific
for attempts at engine re-start. should be carried out on the equipment to show altitudes, aircraft descent rates and crewmember
In the event this course of action fails to restore that it will meet the oxygen supply requirements pulmonary ventilation requirements for the
MSOC pressure, the BOS shall be capable of and duration. Man rating testing, as part of the required workload.
supporting a descent to below 10,000 feet altitude. aircraft system, should be carried out on the
equipment to show that it will meet physiological
requirements.
1.4.5.20. The selection and de-selection of If a mechanical (non-electrical) switch or lever is Automatic selection may be achieved by either a
the BOS shall be possible by both automatic and provided for selection of the BOS the force release valve on the equipment or, in the case of
manual means. required to operate it shall not be greater than a live system, by a remote selection valve.
13.6 kg (30 lb) or less than 6.8 kg (15 lb). For Manual selection may be achieved by activation
seat-mounted systems, all crew positions shall of the ejection seat Emergency Oxygen control, or
be considered. by activation of another control accessible to the
crewmember.
1.4.5.21. Where applicable, the BOS shall Design calculations, aircraft system altitude In addition, when the aircrew equipment
also support the use of pressure garments and testing and man rating testing should be carried assemblies (AEA) for the aircraft calls for an
NBC protective equipment. out on the equipment to show that it will meet the partial pressure garments, the BOS is to be
oxygen supply requirements and duration when capable of inflating the chest / trunk garment
used with pressure garments and NBC protective bladders and should also support aircrew NBC
equipment. equipment where necessary.
1.4.5.22. The BOS may be combined with Design calculations, aircraft system altitude The requirements for Emergency Oxygen may be
the Emergency Oxygen system, subject to testing and man rating testing should be carried combined with the requirements for Back-up
sufficient quantity of breathing oxygen being out on the equipment to show that it will meet the Oxygen Supply, as long as a sufficient volume of
provided to accommodate the needs of both. oxygen supply requirements and duration for oxygen is provided and the unit is ejection seat
both Back-Up and Emergency use. mounted.
1.4.5.23. The BOS shall be selected Aircraft system testing in an altitude chamber The aircraft systems will typically select the BOS

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automatically if the cabin altitude exceeds its shall be carried out on the equipment to show on detection via a transducer of cabin pressure
normal maximum (taking into account tolerances compliance with the automatic selection below a preset value.
in the cabin pressure control system). requirements.
1.4.5.24. The BOS shall have a minimal Qualification testing shall be carried out on the Leakage shall be minimised to prevent excessive
leakage rate. equipment to show that it will meet the specified topping up of the BOS.
requirements.
1.4.5.25. The BOS shall be fitted with Qualification testing shall be carried out on the Pressure gauges, pressure switches and
suitable means to indicate to the crew the quantity equipment to show that it will meet the specified pressure transducers shall be fitted to the
of oxygen stored. requirements. equipment to allow remote indication of contents,
as appropriate to the system requirements.
1.4.5.26. The Back-up Oxygen Supply shall Qualification testing shall be carried out on the Storage pressures of 1,800 psig are commonly
contain a stored source of high pressure breathing equipment to show that it will meet the specified used for Back-up Oxygen Systems. Other
oxygen. requirements. pressures may be considered if advantageous.
1.4.5.27. The Back-up Oxygen Supply shall Qualification testing shall be carried out on the The Back-up Oxygen System shall be fitted with
contain the means to allow an outlet of breathing equipment to show that it will meet the specified an integral pressure-reducing valve to allow the
oxygen a suitable reduced pressure and flow. requirements. unit to supply a suitable regulated pressure and
flow to the aircraft system.
The flow shall be sufficient to allow the
crewmembers physiological requirements to be
met.
1.4.5.28. The Back-up Oxygen Supply shall Qualification testing shall be carried out on the A bursting disc shall be fitted to protect the high
be fitted with pressure relieving means capable of equipment to show that it will meet the specified pressure oxygen storage. Consideration shall be
protecting the equipment and restricting the requirements. given to discharging the oxygen overboard when
maximum outlet pressure in the event of an the equipment is airframe mounted.
aircraft fire or failure of the equipment. A pressure relief valve, with sufficient flow
capacity, shall be fitted to the reduced pressure
outlet.
1.4.5.29. The Back-up Oxygen Supply shall Qualification testing shall be carried out on the A charging valve to STANAG 7106, or alternative
be capable of being replenished either whilst equipment to show that it will meet the specified as specified, shall be fitted to allow replenishment
installed in the aircraft, or on the ejection seat or requirements. of the stored oxygen.
after removal to a ground, oxygen charging facility.
1.4.5.30. The Back-up Oxygen Supply shall Qualification testing shall be carried out on the See Def Stan 00-970 Part 1 Section 6.14 for
be fitted with a high pressure oxygen storage cylinder to show that it will meet the specified Pressurised Gas Storage Vessels specifications.
cylinder. requirements. Fragmentation testing to the requirements of
LEAFLET 33 shall be carried out if the energy
level defined by LEAFLET 30 requires it.

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Emergency Oxygen (EO) Systems
1.4.5.31. In general, provision shall be made The emergency oxygen set shall be capable of See Table 4 of LEAFLET 3.
for a separate gaseous emergency oxygen supply preventing hypoxia under all foreseeable
to be available either automatically on escape emergency conditions. As a minimum, the The emergency and back-up oxygen supply
from the aircraft or on manual selection by the following circumstances shall be considered: requirement may be satisfied by a single oxygen
crewmember during flight. supply mounted on the ejection seat.
In particular, an Emergency Oxygen Supply shall (a) aircrew outside the aircraft after
be fitted to the ejection seat of any aircraft which an escape at the maximum altitude which
can be operated at altitudes of above 25,000 feet. the aircraft can attain and during descent to
The emergency supply shall operate 10,000 ft.
independently from the primary gas source and be
of sufficient quantity to protect the crewmember (b) aircrew inside the aircraft, in a
following escape at the aircraft’s maximum descent from the maximum altitude which
operational altitude. the aircraft can attain down to 10,000 ft with
the main oxygen system inoperative.

Unless otherwise specified, the oxygen set shall

be of the demand type.
1.4.5.32. The Emergency Oxygen Supply The Emergency Oxygen Supply shall be shown The volume of oxygen required for the defined
shall be of sufficient quantity to protect the to have sufficient stored oxygen capacity to meet worst-case scenario shall be calculated at an
crewmember for the longest anticipated time to the specified worst-case crewmember breathing early stage of the project.
descend to a safe altitude whether the cabin is scenario by calculations during the system The calculations shall take into account aircraft
pressurised or depressurised. design stage. descent rates, ejection requirements, and
Both pressurised and depressurised cabin crewmember pulmonary ventilation requirements
scenarios shall be analysed to determine the for the required workload.
worst case duration requirement in an It should be remembered that at any given aircraft
emergency descent. altitude the volumetric crew breathing demand will
Aircraft system testing in an altitude chamber use stored gas more quickly in a pressurised
should be carried out on the equipment to show cabin than in a depressurised one due to the
that it will meet the oxygen supply requirements higher mass flow associated with denser gas.
and duration.
Man rating testing, as part of the aircraft system,
should be carried out on the equipment to show
that it will meet the crewmembers physiological
and duration requirements.
1.4.5.33. The selection and de-selection of The force required to operate the Emergency Automatic selection may be achieved by either a
the Emergency Oxygen Supply shall be possible Oxygen Supply shall not be greater than 13.6 kg release valve on the equipment or, in the case of
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by both automatic and manual means. (30 lb) or less than 6.8 kg (15 lb). For seat- a live system, by a remote selection valve.
mounted systems, all crew positions shall be The selection valve shall be activated
considered. automatically on seat ejection. Manual selection
may be achieved by activation of the ejection seat
Emergency Oxygen control, or by activation of
another control accessible to the crewmember.
1.4.5.34. Where applicable, the Emergency Design calculations, aircraft system altitude In addition, when the aircrew equipment
Oxygen Supply shall also support the use of testing and man rating testing shall be carried out assemblies (AEA) for the aircraft calls for an
pressure garments and NBC protective on the equipment to show that it will meet the partial pressure garments, the Emergency
equipment. oxygen supply requirements and duration when Oxygen Supply is to be capable of inflating the
used with pressure garments and NBC protective chest / trunk garment bladders and should also
equipment. support aircrew NBC equipment where
necessary.
1.4.5.35. The Emergency Oxygen Supply Design calculations, aircraft system altitude The requirements for Emergency Oxygen may be
may be combined with the Back-up Oxygen testing and man rating testing shall be carried out combined with the requirements for Back-up
Supply, subject to sufficient quantity of breathing on the equipment to show that it will meet the Oxygen Supply, as long as a sufficient volume of
oxygen being provided to accommodate the needs oxygen supply requirements and duration for oxygen is provided and the unit is ejection seat
of both. emergency use. mounted.
1.4.5.36. The Emergency Oxygen Supply Qualification testing shall be carried out on the Leakage shall be minimised to prevent excessive
shall have a minimal leakage rate. equipment to show that it will meet the specified topping up of the Emergency Oxygen Supply.
requirements.
1.4.5.37. An indication shall be provided of Qualification testing shall be carried out on the Pressure gauges, pressure switches and
the contents remaining in the Emergency Oxygen equipment to show that it will meet the specified pressure transducers shall be fitted to the
Supply. requirements. equipment as appropriate to the system
requirements.
1.4.5.38. The Emergency Oxygen Supply Qualification testing shall be carried out on the Storage pressures of 1,800 psig are commonly
shall contain a stored source of high pressure equipment to show that it will meet the specified used for Emergency Oxygen Supply. Other
breathing oxygen. requirements. pressures may be considered if advantageous.
1.4.5.39. The Emergency Oxygen Supply Qualification testing shall be carried out on the The Emergency Oxygen Supply shall be fitted
shall contain the means to allow an outlet of equipment to show that it will meet the specified with an integral pressure-reducing valve to allow
breathing oxygen a suitable reduced pressure and requirements. the unit to supply a suitable regulated pressure
flow. and flow to the aircraft system.
The flow shall be sufficient to allow the
crewmembers physiological requirements to be
met.
1.4.5.40. The Emergency Oxygen Supply Qualification testing shall be carried out on the A bursting disc shall be fitted to protect the high

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shall be fitted with pressure relieving means equipment to show that it will meet the specified pressure oxygen storage.
capable of protecting the equipment and requirements. A pressure relief valve, with sufficient flow
restricting the maximum outlet pressure in the capacity, shall be fitted to the reduced pressure
event of an aircraft fire or failure of the equipment. outlet.
1.4.5.41. The Emergency Oxygen Supply Qualification testing shall be carried out on the A charging valve to STANAG 7106, or alternative
shall be capable of being replenished either whilst equipment to show that it will meet the specified as specified, shall be fitted to allow replenishment
either installed on the ejection seat, or after requirements. of the stored oxygen.
removal to a ground, oxygen charging facility.
1.4.5.42. The Emergency Oxygen Supply Qualification testing shall be carried out on the Def Stan 00-970 Part 1 Section 6.14 for
shall be fitted with a high pressure oxygen storage cylinder to show that it will meet the specified Pressurised Gas Storage Vessels specifications.
cylinder. requirements. Fragmentation testing to the requirements of
LEAFLET 33 shall be carried out if required
according to the energy level defined by
LEAFLET 30.
1.4.6. Specific System Requirements
The majority of military aircraft breathing system applications will fall into one of the categories described below; however it should be noted that fulfilment
of a specific military role may necessitate requirements which are not so easily categorised. In such cases the tenor of this document should be followed,
particularly with respect to Clause 1.4.1.1 (a) to (i) and LEAFLET 3, and appropriate specialist advice taken.
Fast Jet Requirements – Low Differential Cabins
Low differential cabins are defined as those that cannot exceed a maximum of 36.2 ± 1.4 kPa (5.25 ± 0.2 psi) differential pressure between the interior and
exterior of the aircraft.
1.4.6.1. At all times the aircrew shall breath The breathing system shall be such as to An alveolar oxygen partial pressure of 103 mmHg
gas resulting in alveolar oxygen partial pressure maintain a minimum alveolar oxygen partial is produced at sea level by an inspired (dry gas)
that is the same or greater than that which would pressure of 13.33 kPa (103 mmHg) at all cabin oxygen partial pressure of 160 mmHg. This is due
be experienced breathing environmental air at sea altitudes below 34,000 feet. to the change in temperature and pressure
level, unless the aircraft has decompressed to an conditions from cabin (or mask) to body, plus the
altitude of 34,000 feet or above. partial pressures of water vapour and carbon
dioxide present within the lungs. The 103 mmHg
PAO2 level is equivalent to breathing 21 %
oxygen at standard (sea level) atmospheric
pressure.
1.4.6.2. In the event of a decompression of The following three methods shall be used in The two latter techniques result in the "notch"
the cabin to an altitude of 34,000 feet or above, combination to prevent the fall of PAO2 below 30 often portrayed in requirements plots of breathing
the alveolar oxygen partial pressure shall not fall mmHg on decompression: gas oxygen concentration against a base of cabin
below 30 mmHg at any time. altitude. The "notch" is only applicable prior to a
(a) The volume of gas breathed by decompression of the cabin as it protects against

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the crewmember after commencement of the possibility of that event. An example is
decompression prior to receiving 100 % illustrated in Figure 2 of LEAFLET 3.
oxygen in the mask shall not exceed 0.6 L The shape, size and location of the "notch" for a
ATPD. particular aircraft is dependent upon the aircraft
pressurisation schedule, the flight ceiling and the
(b) The breathing system shall PBA schedule. It may be calculated using the
provide breathing gas to the mask with an alveolar gas equation and decompression
oxygen concentration sufficient to maintain equation as described in Section 4 of LEAFLET 3.
an alveolar oxygen partial pressure which PBA is intended as a short-term protective
will ensure that with the cabin pressurised measure during the decompression and initial
the alveolar PO2 is greater than 103 mmHg descent.
and that rapid decompression of the cabin
will not reduce the alveolar PO2 below 30
mmHg, and shall provide 100 % oxygen to
the mask during the decompression and
subsequent descent to an altitude at which
an alternate breathing gas (or air) can
safely be used.

(c) The breathing system shall

increase the breathing mask cavity
pressure above that of the surrounding
cabin pressure after the decompression.
This is Pressure Breathing with Altitude
(PBA).
1.4.6.3. The risk of adverse physiological The breathing system shall limit the maximum Under sustained +Gz acceleration greater than
conditions associated with breathing high oxygen concentration supplied to the breathing about 3 G whilst breathing 100 % oxygen,
concentrations of oxygen, particularly in aircraft mask to 60 % at cabin altitudes up to 15,000 feet particularly when wearing an anti-G suit,
capable of achieving sustained +Gz accelerations rising to 75 % at 20,000 feet, taking into account acceleration atelectasis can result due to the lack
in excess of 3 G, shall be minimised by limiting the the requirements for prevention of hypoxia in the of nitrogen.
maximum oxygen concentration in the breathing event of a cabin decompression. Breathing high concentrations of oxygen can also
gas unless an emergency condition such as a result in delayed otitic barotraumas (absorption of
cabin decompression has occurred. oxygen from the middle ear causes distortion of
the ear drum).
See Section 4 of LEAFLET 3 for a more detailed
discussion of these conditions.
There is a potential conflict between the
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enhanced oxygen concentration required to
protect against hypoxia in the event of cabin
decompression (see 0 above) and the limit to
maximum concentration required to prevent
acceleration atelectasis. Even if there is no direct
conflict, the resultant narrow control band for
oxygen concentration at higher pressurised cabin
altitudes may be difficult to implement.
Such conflicts and difficulties should be resolved
in discussion with the Project Team Leader and
with aeromedical advice, taking into account the
likely frequency of high-G manoeuvres at higher
aircraft altitudes.
1.4.6.4. Physiological protection from cabin 100% oxygen shall be available at the breathing In MSOC systems, minimising the volume of pipe-
decompressions to 25,000 feet and above mask within 2 seconds of the cabin work from the emergency oxygen supply to the
requires the delivery of 100% oxygen within 2 decompressing to 25,000 feet or above. 100% breathing regulator, and minimising the volume
seconds of the commencement of decompression. oxygen shall be maintained whilst the cabin between the regulator and mask, is necessary to
altitude exceeds 25,000 feet. achieve rapid provision of 100% oxygen.
Transport Requirements – High Differential Cabins
High differential cabins are defined as those that can exceed 36.2 ± 1.4 kPa (5.25 ± 0.2 psi) differential pressure between the interior and exterior of the
aircraft. Typically, such aircraft have a maximum differential pressure capability of 55 kPa (8 psi). The requirements for aircrew (including pilots, co- pilots,
communications operatives, loadmaster, etc) are defined herein. It may be stipulated that the requirements of CS 25.1443 apply in place of those below.
The requirements for passive passengers are defined in Clause 1.4.6.7 below.
In this type of aircraft the crew (and passengers) normally breathe cabin air at pressurised cabin altitudes of up to 8,000 feet, the degree of hypoxia
experienced at this altitude having been deemed acceptable by service practice and international standards.
1.4.6.5. The flight crew shall be fully Supplemental oxygen shall be provided by It is typically required by regulations that watch
protected against any significant impairment of means of a demand regulator system providing keeping members of the flight-deck crew be
performance due to hypoxia in the event of a safety pressure at cabin altitudes between wearing oxygen masks when the aircraft is above
cabin decompression at high altitude and any 12,000 and 40,000 feet and pressure breathing 40,000-43,000 feet.
subsequent flight at cabin altitudes above 8,000 at cabin altitudes above 40,000 feet. Where gaseous or liquid oxygen is the source of
feet. breathing gas the regulator should provide air
dilution to economise the rate of consumption of
oxygen and to minimise the risk of delayed otitic
barotraumas.
Consideration should be given to automatic
washing out (purging) of gas containing nitrogen
from the distribution system on decompression in
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order to ensure the delivery of 100% oxygen at all
crew masks within 2 seconds of the System
design should include consideration of mask
stowages and other ergonomic factors to
minimise mask donning time.
1.4.6.6. The crew shall be provided with a The breathing system shall include a mask and In practice this is likely to be the same equipment
means of protection against the effects of smoke regulator for use by each crewmember in the as for hypoxia prevention on cabin
and/or toxic fumes in the cabin. event of smoke and fumes in the cabin. decompression.
Passenger Requirements – High Differential Cabins
High differential cabins are defined as those that can exceed a maximum of 36.2 ± 1.4 kPa (5.25 ± 0.2 psi) differential pressure between the interior and
exterior of the aircraft. Typically, such aircraft have a maximum differential pressure capability of 56 kPa (8psi) and ensure that the cabin equivalent altitude
does not exceed 8,000 feet (typical of commercial airliners).
1.4.6.7. If the cabin altitude exceeds 8,000 The tracheal oxygen concentrations required by The source of supplementary oxygen may be
feet supplementary oxygen shall be made CS 25.1443 shall be met or exceeded. stored gas, chlorate candles, or MSOC.
available to passengers in accordance with the It should be noted that passenger mask
requirements of CS 25.1443 assemblies typically contain a controlling orifice
sized for a specific flow when supplied with 100 %
oxygen at a specific pressure. Orifice sizing
should be re evaluated if an MSOC source is
used to take account of the different oxygen
concentration and pressure conditions.
Consideration should also be given to the use of
an altitude-sensitive control valve governing the
supply pressure to the passenger masks. This will
maximise the endurance of stored gas sources or
minimise the required number or size of MSOC.
1.4.7. Biological and Chemical (BC) Protection
Biological Protection
1.4.7.1. Biological protection is afforded by If filter performance characteristics are not Biological protection shall be co-ordinated through
suitable clothing and the use of suitable grades of available from the manufacturers, then the the appropriate Project Team Leader, who will
filter element media in the breathing gas line. suitability of candidate biological filter protection make available all necessary specialist
Protection is also afforded by ensuring that the media shall be tested (e.g. sodium flame establishments to assist in the design and
breathing mask or respirator does not allow the photometry). validation of the life support system.
user to breathe under negative pressure in a The particulate filter shall be capable of achieving
manner which could allow inwards leakage of an efficiency of 99.997% when challenged with a
biological warfare agents into the breathing sodium chloride aerosol of concentration 10 mg

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supplies (and where applicable, respirator demist m-3 of a mass median diameter of 0.4 micron
supplies). The filter shall possess an absolute filter and as measured using a flame photometer.
rating of 0.4 micron or smaller.
Chemical Protection
1.4.7.2. Chemical protection is afforded by Suitable aircrew equipment to afford chemical Disposable activated and impregnated carbon
suitable clothing and use of stored breathing gas protection shall be defined by the Project Team (charcoal) filters form the basis of suitable
in conjunction with an oronasal mask. Leader. chemical agent filters. These normally combined
Chemical hardening of equipment should be with a particulate high efficiency particulate air
undertaken by the manufacturer, commensurate (HEPA) filter to remove residual radioactive dust,
with the threat identified by the Project Team chemicals in aerosol form and biological agents
Leader. from the breathing gas supply.
1.4.7.3. Use of materials susceptible to The selection of materials for use in breathing
absorb contaminants shall be avoided such that equipment shall be made with regard to
reliability or performance is not degraded. minimising the absorption of contaminants.
1.4.7.4. If cabin air is used in the regulator Cockpit air used to dilute breathing gas shall be The use of air dilution in a chemical warfare
to create physiologically appropriate breathing gas filtered, prior to entering the breathing regulator, operation is not recommended.
oxygen concentrations, then a filter shall be to remove chemical agents.
provided to remove chemical agents from the
dilution (airmix) air.
1.4.7.5. Where breathing gas is sourced Testing the PSA device under representative Up to date toxicological data shall be reviewed to
from the conditioning of engine bleed air by conditions of bleed air composition (of e.g. allow the significance of any dosages of chemical
molecular sieve pressure swing devices, use of pressure, temperature and water content) with agent penetrating into the breathing and demist
conventional filters for protection against chemical
chemical agents identified by the Project Team supplies to be assessed.
warfare agents may be reassessed because such will normally be required unless analogous data Validation testing should be conducted at an
pressure swing adsorption (PSA) devices can is available for closely similar equipment which approved authorised specialist establishment
inherently provide a degree of protection to an operates in the same way, and which is under the control and guidance of the Project
extent dependant on the operating conditions. acceptable to the Project Team. Such testing Team Leader.
should encompass the range of bleed air
compositions known or expected, and should be
carried out for a period of time sufficient to allow
the Project Team to be satisfied that the
equipment will function for an acceptable period
of time, which is also to be identified by the
Project Team.
Influence Of BC Respirators & Man Mounted BC Filters
1.4.7.6. The increase in breathing Test and evaluation of the biological and Guidance on acceptable impedance to breathing

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impedance and the increase in volume between chemical (BC) respirator or man mounted BC whilst wearing BC respirators may be found in
oxygen source and oronasal mask which may filter shall be conducted under conditions AIR STD 61/112/2B.
result from the introduction of BC respirators representative of those which might be BC protection shall be co-ordinated through the
and/or filters shall be minimised, and shall be encountered on the specific aircraft. Project Team Leader, who will make available all
shown to have no significant detrimental effect on necessary specialists to assist in the design and
breathing system performance. validation of the life support system, including
aeromedical advice as to the acceptability of
increases in breathing impedance and system
volume.
1.4.8. System Test Requirements
Engineering Qualification Testing
1.4.8.1. Qualification testing is the process Environmental testing shall be performed to the Qualification may be achieved by test of a
of verifying and declaring conformance with each requirements of BS 3G 100 or equivalent as production standard system or unit, by
requirement of the customer’s equipment directed by the Project Team. analogy/design with similar equipment or by
specification. analysis. An alternative method is for the
equipment to be cleared by in-service experience.
1.4.8.2. A pre-qualification test performance Prior to qualification testing, the equipment shall The pre-qualification test performance results can
record shall be completed. be operated at standard ambient conditions to be used for comparison with data obtained
obtain and record data determining compliance before, during and after the qualification testing.
with Acceptance/Production test procedures.
1.4.8.3. The Qualification Test Programme Consideration shall be given to the environmental Analysis of test data recorded for similar
shall be representative. characteristics for the regions in which the equipment or in service experience may prove
equipment is to be deployed. useful in determining test parameters.
Consideration shall be given to the Whenever practical, specific test levels, ranges,
characteristics of platforms on which the rates and durations shall be derived from
equipment is to be carried or operated. measurements made on actual or appropriately
Consideration should be given to the number of similar equipment.
test units allocated for qualification testing. Tailoring of environmental test conditions to
Consideration shall be given to the orientation of individual equipment applications can be
the test unit during testing A test plan shall be employed to ensure that equipment is tested for
provided to the project director, which identifies resistance to the environmental stresses it will
the proposed tests and test conditions and encounter during its life cycle.
provides a narrative that supports the selected Test items shall normally be installed in the test
(proposed) test conditions and methods. facility in a manner that will simulate service
This should reference the range of ECS usage, with connections made and
conditions found or expected for the aircraft. instrumentation attached as necessary.

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1.4.8.4. The test parameters and Functional testing shall be performed to the Clearance of the system or unit under test to any
environment shall be selected to stress the requirements of AIR STD 61/101/2B one qualification test is achieved by
equipment in the way that reflects worst-case (Development Test and Evaluation of Aircraft demonstration of satisfactory performance of the
conditions. Oxygen Delivery Systems) or equivalent as unit under test for key performance parameters
directed by the Project Team. before and after each qualification test to
The pre and post qualification tests performed to demonstrate that no significant degradation has
verify that no permanent damage results from the occurred.
qualification test conditions provides the When operation of the test item is required during
clearance for that qualification environment. The the test exposure, suitable tests shall be
pre and post tests shall comprise critical performed to determine whether the test exposure
performance parameters that the qualification is producing changes in performance when
tests are likely to stress. compared with pre and post-test data.
Functional endurance test work shall form part of
qualification testing. This may be performed as
part of the system.
1.4.8.5. Where full Formal qualification of This testing shall stress the unit to the worst Typically, Preliminary qualification endurance
the equipment cannot be performed prior to first conditions stated in the respective equipment testing shall be performed as a simple functional
flight, then there shall be, at minimum, sufficient specification. test in real time, preferably in conjunction with
testing performed to provide evidence for safe other components in the system.
development flying. Formal qualification endurance testing is a real
Preliminary qualification testing is a stage between time test as in the preliminary endurance test, but
the minimum safety of flight clearance and the full may incorporate the application of realistic
Formal qualification requirement. vibration levels.
1.4.8.6. Where necessary, and as directed Any NBC compatibility tests are to be at the
by the Project Team Leader, qualification for NBC discretion of the Project Team Leader, who shall
usage shall be carried out using processes and also define the requirements of any necessary
procedures approved by the Project Team. tests and the methodology to be employed.
1.4.8.7. Testing shall be performed to By demonstration or as defined by the Project A recommended schedule of tests for liquid
ensure correct interaction between system Team. oxygen systems is given in LEAFLET 5.
components and demonstrate compliance to the See MAP RA 5208.
aircraft system specification.
Interconnecting aircraft pipe work shall be
simulated for all system test work. Testing shall
also verify the correctness of the installation in
terms of location of components within the
airframe.
The system shall be accessible for the purposes
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of maintenance and inspection, and should be
capable of removal from the aircraft in a time
period defined by the Project Team.
Human Subject Testing (“Man Rating”)
1.4.8.8. Following engineering Qualification Minimum requirements are contained in AIR STD Machine-based testing as is typically used in
Testing of a life support related system, its 61/101/2B. engineering qualification testing cannot easily
performance shall be verified by testing with At least three examples of mask/regulator simulate certain aspects of human physiology
human subjects. combinations covering the minimum, mean and (such as gas diffusion between lung gases and
maximum values of: the blood, or the impedance to flow of the
respiratory tract), nor can it cover the variation in
(a) impedance to breathing, physiology that might be expected in the aircrew
population.
(b) safety pressure, and It is therefore mandatory to demonstrate
satisfactory performance with human subjects
(c) pressure breathing characteristics unless it can be shown, to the satisfaction of the
(if applicable) Project Team Leader and his/her aeromedical
advisors, that the system is sufficiently analogous
should be tested with human subjects. to existing in-service or qualified equipment for
such testing to be unnecessary.
All aspects of testing with human subjects (e.g.
subject selection and training, suitability of test
facility, test protocols, ethical considerations)
should be in accordance with national standards
in the country in which testing will take place.
In the UK such aspects will be subject to prior
approval by a nationally-recognized Research
Ethics Committee. Advice may also be sought
from the RAF Centre of Aviation Medicine.
1.4.8.9. Human subject testing shall verify The test parameters employed shall cover the Changes in equipment operating mode requiring
performance of the system under conditions range of pressure-altitudes, equipment inlet particular attention are, for example, changes in
typically encountered by the aircrew in normal pressures and temperatures, and (if relevant) breathing regulator outlet pressure characteristic /
flight, particularly at the extremes of the equipment +Gz accelerations for which the system is breathing gas mixture at certain cockpit / cabin
design envelope and at points in the envelope designed in normal operation. altitudes, changes in demand flow due to auxiliary
where changes in equipment operating mode Actual aircrew drills should be followed for loads (e.g. demisting), or changes in cyclic control
occur. normal system operation, as far as is practicable of MSOC units.
in the test environment, in order to verify their
suitability.
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1.4.8.10. Human subject testing shall verify The test parameters employed shall cover any Close collaboration with the airframe
performance of the system under abnormal relevant abnormal events including (but not manufacturer is required to ensure that the
aircraft conditions for which the equipment was limited to) cockpit / cabin decompression, low human subject testing is realistic. For example,
designed. temperature environment, operation of backup attention should be paid to the maximum likely
and / or emergency systems. decompression time at various aircraft altitudes,
Testing shall be as representative as is safely the effect on cockpit / cabin altitude of
possible of realistic conditions in the aircraft at aerodynamic suction caused by a structural
the time of the abnormal event. failure at high airspeeds, the triggering altitude for
Actual aircrew drills shall be followed for automatic backup or emergency systems, the
abnormal system operation, as far as is sequencing of changes from primary to backup or
practicable in the test environment, in order to emergency breathing systems.
verify their suitability.
1.4.8.11. Human subject testing shall Tests shall be defined covering (but not limited A cycle ergo meter is typically used to establish
demonstrate satisfactory steady-state and to) the following aspects of the aircraft life different subject work rates in a controlled and
dynamic performance of the aircraft life support support system: breathing impedance, stability, reproducible manner.
system in accordance with the physiological prevention of hypoxia, and (if relevant) Reading aloud is included so that the high
requirements set out in Clause 1.4.2.15 performance enhancement under +Gz inspiratory peak flows and rates of change of flow
acceleration. that are characteristic of human speech may be
A representative range of breathing rates and tested.
flows shall be tested. Safety permitting, these A standardised test may be used to demonstrate
rates and flows shall be tested both with the breathing system stability. This consists of 50
subject at rest and with the subject undergoing a breaths at rest, 50 breaths at rest with speech, 50
work rate representative of the workloads to be breaths at light to moderate work rate (50 - 80W),
encountered by aircrew during aircraft and 50 breaths at this work rate with speech. This
operations. In both cases (resting & working), a test is repeated with 5 different human subjects,
period of testing in which the subject speaks or and is usually conducted at ground level (since
reads aloud shall be included. experience has shown that instabilities are more
likely to occur with the densest gas). However,
additional stability tests may be necessary at
altitude for particular systems. It is important that
the AEA used for stability testing is fully
representative of that to be worn by aircrew.
Instrumentation used should not alter the
characteristics of the system and typically
consists of a single transducer measuring mask
cavity pressure, with a flat response to at least
100 Hz.
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1.4.8.12. Human subject testing shall Each test case shall be satisfactorily completed It is preferable to carry out each test multiple
demonstrate that satisfactory performance of the at least twice, using a different human subject for times using a range of human subjects whose
aircraft life support system is not limited to a each run. relevant physiological characteristics cover the
particular set of physiological characteristics. In all applications, aeromedical advice shall be extremes of the intended aircrew population.
sought as to the minimum number of human However, it is recognized that logistics, cost and
subjects for each test case necessary to verify the availability or otherwise of a sufficient number
satisfactory performance of the system. of suitable subjects may render this impracticable.
B - Rotorcraft
1.4.1 Refer to IR OPS CAT.IDE.H.240

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1.5. ICE PROTECTION
A – Fixed Wing
This Clause applies to all aeroplanes whether or This section is similar to CS 25.1419. It contains
not equipped with airframe ice protection, which the requirements for protecting aeroplanes
are required by the Aeroplane Specification to against ice accretion
enter, or operate in, icing conditions. The atmospheric conditions in which ice accretion
may occur, and their probable extent, are defined
in Section 2 Leaflet 7.
1.5.1. OPERATIONAL REQUIREMENTS
1.5.1.1. The aeroplane shall the capable of In order to satisfy the requirements of this Clause Notes:
meeting the Service operational requirements it shall be shown that, when the aeroplane is
stipulated in the Aeroplane Specification. The operated in the stipulated conditions and for the (1) See Section 2 Leaflet 7 Table 1 for
following conditions with durations are those that required durations, there will be no hazard to the the definition of conditions (a) to (h).
should be defined in the specification: aeroplane or its crew, and no unacceptable Conditions (i) and (j) are to be agreed with
degradation in:- the Project Team Leader (PTL).
(a) The Continuous Maximum icing
conditions. (a) The performance of the aeroplane (2) Conditions (i) is a special case for
or its systems. conditions (a), (b), (c) and (d). It covers
(b) The Intermittent Maximum icing circumstances where the airframe may be
conditions. (b) The handling qualities of the substantially colder than the surrounding
aeroplane. atmosphere with consequent ice accretion
(c) The Ice Crystal cloud or mixed at ambient temperatures higher than the
cloud conditions. (c) The performance of weapon normal freezing limit.
systems carried in or on the aeroplane.
(d) Falling snow, continuous. (3) For condition (j) it is recommended
Note: Any degradation would be unacceptable if that contaminant depths up to not less than
(e) Falling snow, intermittent. it resulted in the aeroplane or its systems being 15 mm Water Equivalent Depth (WED) be
unable to meet the requirements of the considered. (Water Equivalent Depth is the
(f) Blowing or recirculating snow. Aeroplane Specification. product of the mean depth of the
In satisfying the requirements relating to contaminant, and its relative density.
(g) Freezing fog. operation in condition (j), consideration shall be
given to the following:- GUIDE TO ICE PROTECTION SYSTEMS BASIC
(h) Freezing rain/drizzle. SYSTEM TYPES
(a) the effect of the runway
(i) Descent through and/or loiter in contaminant on aeroplane runway The two basic types of ice protection systems are

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moist air after the aeroplane has cold performance (acceleration and braking); as follows:
soaked at altitude.
(b) the effect on aeroplane ground (a) Anti-icing: Systems used to prevent
(j) Operation from or onto runways handling characteristics (e.g. nose wheel formation of ice on part of the surface of the
contaminated with ice, snow (both lying and steering, crosswind limitations, and aeroplane, usually achieved through
hard-picked), slush or standing water. minimum control speed on the ground continuous heating of part of an aeroplane
following a sudden loss of engine thrust); surface, but can be based on use of
freezing point depressant fluids in certain
(c) the impact of water/slush limited applications.
displaced by the undercarriage into engine
or other intakes, against the airframe or (b) De-icing: System used to
externally mounted probes and equipment, periodically remove ice from part of the
and into undercarriage bays or control surface of the aeroplane (before it reaches
gaps, with possible subsequent freezing. a size that could cause unacceptable
degradation of the aeroplane/system
performance or hazard the aeroplane during
shedding).

Anti-icing and de-icing systems can be used in

combination, for example on propeller spinners.
A brief description of the various means of ice
protection is given in Leaflet 8.
SYSTEMS REQUIREMENTS - GENERAL
1.5.1.2. The system shall be designed so The protection system shall be capable of In determining the areas requiring protection the
far, as is reasonable to provide symmetric satisfactory operation throughout the aeroplane designer shall give particular consideration to the
protection in order that the problems associated flight envelope, unless the Project Team Leader following:-
with asymmetric ice accretion and/or shedding (PTL) agrees otherwise.
may be avoided. In designing to achieve this In determining the extent of ice protection (a) Aerofoils
objective the effects of possible system failures coverage and the degree of protection (e.g. heat
shall be considered (see Clause 1.5.1.5) per unit area) required, the designer shall (b) Moving surfaces
consider all critical combinations of airspeed,
altitude, ambient temperature, liquid water (or ice (c) Propellers
crystal) content, water droplet size and droplet
trajectories. (d) Engine and auxiliary intakes
For any but the simplest shapes mathematical
modelling and/or rig testing may be required to (e) Pitot and static heads and masts
provide quantitative results.
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The extent of ice protection coverage shall also (f) Aerials and sensors
take account of the predicted effects of the
maximum likely ice accretion on the unprotected (g) Weapons and weapon carriers
areas.
The manufacturer must show that the protection (h) Transparencies
system meets the Aeroplane Specification by
means of some combination of modelling, rig and (i) Vents and drains
tunnel testing as well as by flight tests in natural
or simulated icing conditions. Where computer (j) Other items as specified
models or simulated icing conditions are used to
prove compliance, they must be supported by Note: For most shapes and ambient conditions
substantiation and/or validation evidence, and there is a pessimism (i.e. worst or least
must be approved for use by the Project Team favourable) speed which results in the worst
Leader (PTL) accretion rate or greatest heat flux required. This
The manufacturer shall show, by analysis or test, speed is not normally either the lowest or the
that the performance of the ice protection highest speed of which the aeroplane is capable.
system(s) will not be degraded to an
unacceptable extent over the required life and
predicted usage.
1.5.1.3. Not withstanding the formation of
ice, the aeroplane services and ancillary
equipment, e.g. undercarriage, flaps, generators,
speed monitored stability aids and flight
instruments, shall continue to function
satisfactorily.
1.5.1.4. Ice protection power requirements
shall be kept to the minimum necessary to achieve
the required level of protection.
1.5.1.5. Back-up systems/power sources The acceptability of any reduction in protection in
must be provided where failure of the protection the event of such a failure will be decided by the
system could lead to unacceptable flight safety Project Team Leader (PTL)
problems or operating restrictions. On multi-
engine aeroplanes where protection is provided by
the engine, for example by bleed air, alternative
sources must be provided to cater for the effect of
engine failure.
1.5.1.6. Ice protection systems shall not
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introduce secondary ice accretion (run-back)
unless it can be shown to be of no consequence.
1.5.1.7. When a de-icing system is used,
the off-time accretion must not produce
unacceptable aeroplane or system performance
degradation.
1.5.1.8. The ice protection systems,
whether functioning or not, shall not:-

(a) Cause corrosion or deterioration of

any part of the structure, or associated and
adjacent systems.

(b) Impair the operational ability of the

aeroplane or its equipment.
1.5.1.9. The ice protection systems shall be
so designed that toxic fluids or vapours cannot
enter the cabin in normal flight, or as the result of
enemy action.
1.5.1.10. Ice protection systems for engine
intakes, pilots' windscreens, air data probes and
any other system requiring protection prior to take-
off shall be designed for operation during start-up,
taxi and take-off as well as in flight.
1.5.1.11. Any ice or slush which may be
shed from unprotected or de-iced areas of the
aeroplane shall not damage the engines, the
airframe, or other systems or components, and
shall not hazard ground personnel.
AEROPLANE ICE PROTECTION
GENERAL
1.5.1.12. The design features of the An airframe ice protection system may be
aeroplane shall be such as to minimise the required to ensure that the requirements of
hazards of flight in ice forming conditions. Surface Clause 1.5.1.1 are met. Since such an
discontinuities and excrescences shall be avoided installation may however impose an
if possible. unacceptable operational penalty, or since, by

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variation in the flight plan, it may be possible to
reduce the amount of protection required, the
airframe ice protection system for each individual
type shall be the subject of negotiation at an
early stage in the design.
1.5.1.13. All aeroplanes shall comply with the
requirements of Clauses 1.5.1.2 to 1.5.1.11
inclusive irrespective of whether an airframe ice
protection system is fitted or not.
PROPELLERS
1.5.1.14 Excessive ice accretion on Propeller ice protection may be continuous, If ice accretion is to be prevented, then protection
propellers shall be prevented, unless it can be cyclic, or a combination of both as specified in may need to be provided for the cuffs and
shown by propeller icing tests that, in the the Aeroplane Specification. Operation of the ice spinners as well as for the blades.
meteorological conditions of Clause 1.5.1.1: control system shall be accomplished either
automatically or manually (as specified) unless a
(a) the propeller operates satisfactorily, continuous system is provided. Irrespective of the
type of system installed, continuous, operation of
(b) loss of thrust, resulting from ice the ice control system in flight shall not damage
accretion, is within acceptable limits, and or affect the life of the propeller or the system.
Indication of the operation of the ice control
(c) there is no hazard to the aeroplane system, and of any failure to operate, shall be
or its systems, its occupants, or to ground provided in accordance with the Aeroplane
personnel. Specification.
1.5.1.15. The fuselage or cabin in the vicinity
of the plane of the propellers shall be protected
against damage by ice thrown from them.
ENGINE AND AUXILIARY AIR INTAKES
1.5.1.16 Engine air intakes, and any If axial flow turbine engines are fitted, the Note: The engine air intake is assumed to begin
components such as sensor probes which may be adhesion of ice to any part of the interior of the at the upstream end of the duct traversed by the
located in them, shall be protected against the intake shall be completely prevented under the air which eventually passes into the engine, and
accumulation of ice in such quantity as to interfere conditions specified in Clause 1.5.1.1 unless to include the whole of such duct.
with the performance or safe running of the tests have shown that the engine can safely
engine, or which, upon separation, could cause accept such ice as might form in the air intake
damage to the engine or the airframe. and subsequently be ingested, and there is no
significant deterioration in engine performance.
Debris guards fitted to engine or auxiliary intakes

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during ground running shall be designed to
minimise the hazard to engines, airframe or
aeroplane systems due to ice shed from the
guards.
Appropriate advice and procedures shall be
included in the aeroplane operating and
maintenance instructions to avoid damage so
caused.
1.5.1.17. The engine intake anti-icing system
shall be designed, as far as is reasonably
possible, to "fail-safe"; that is to say, to remain in
or revert to the anti-icing mode in the event of a
partial system failure. Continuous operation of the
anti-icing system throughout the operational
envelope of the aeroplane shall not damage or
affect the life of the engine or the system.
1.5.1.18. Auxiliary air intakes and their
associated components shall be protected against
ice accretion (and any subsequent shedding)
which could cause damage to, or unacceptable
loss of efficiency of, any aeroplane system.
AIR DATA SENSORS
1.5.1.19. Pitot, Static, Angle of Attack, Stall Note: For the purpose of demonstrating
Warning and other similar air data sensors shall compliance with Clause 1.5.1.1 pitot and pitot
be protected against ice accretion unless it can be static tubes may be tested to the requirements of
shown that such accretion will not cause a hazard, BS 2G 135 for Type B heads.
or result in an unacceptable loss of performance
of the sensor and its associated systems.
1.5.1.20. Where protection is required, it
shall be automatically selected, unless otherwise
stated in the Aeroplane Specification. The pilot
shall be given a cautionary warning (see Part 1
Section 4 Clause 4.19) of the failure of any air
data sensor heating system.
RADOMES
1.5.1.21. When siting radomes, account shall Even though radomes are manufactured from low

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be taken of the possibility of damage being thermal conductivity materials, such structures,
caused by ice shed from other parts of the depending on shape and position, readily accrete
airframe. ice and slush. This accretion may attenuate the
radar signal, and may also degrade aeroplane
performance and hanging.
WEAPONS AND WEAPON CARRIERS
1.5.1.22 Weapons (system, sights, stores,
pods, guns etc.,) whilst being carried by
aeroplanes in icing or snow, must not be
adversely affected by accretions during use,
deployment or jettison, unless the weapon or
aeroplane specification permits otherwise.
Furthermore, ice accretion on such systems must
not hazard the aeroplane during use of the
weapon system.
TRANSPARENCIES
1.5.1.23. Adequate areas of the The ice protection system shall be designed to (See also Part 1 Section 4 Clause 4.17)
transparencies used by the crew for the safe and function by a suitable means approved by the
effective operation of the aeroplane in its Project Team Leader (PTL). Where applicable,
designated role shall be maintained free of ice. the effects of aerodynamic heating will be
included in the thermal analysis. The system
shall be capable of correct operation during
engine warm-up, taxying, take-off and
touchdown, as well as throughout the normal
flight envelope of the aeroplane.
1.5.1.24 Redundant systems shall be The ice protection system shall be tested to
provided if the failure of a single system would demonstrate compliance with Clause 1.5.1.23
result in an inability to comply with Clause without causing overheating or other detrimental
1.5.1.23. If redundant systems are required, failure effects on the transparency (see Part 1 Section 4
of a single system shall not result in the failure of Clause 4.18)
the remaining ice protection provisions to clear an
adequate area of the windscreen.
VENTS AND DRAINS
1.5.1.25. Airframe and systems vents and See also Leaflet 6.
drains shall be designed or protected against
blockage by ice, snow and slush.

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1.5.1.26. Water systems shall be so
designed that the waste water is either collected
and retained in the aeroplane, or is discharged in
such a way as to prevent it eventually falling from
the aeroplane in the form of lumps of ice.
ICE PROTECTION SYSTEMS CONTROLS AND INDICATORS
1.5.1.27. The following indications shall be All ice protection systems shall, unless otherwise Ice detectors shall:
provided to the pilot:- stipulated in the Aeroplane Specification or
agreed with the Project Team Leader (PTL): (a) provide an accurate indication of
(a) Accurate indication of the onset of ice build-up or impending ice formation
icing. (a) operate automatically when ice (depending on whether they are of the
begins to form, accretion or inferential type) over the full
(b) Configuration that the protection range of temperatures and associated
systems are operating. (b) automatically cease to operate as conditions required by the Aeroplane
soon as it is no longer required, Specification;
(c) System failure indication.
(c) be capable of being switched on- (b) be located so as to give an
(d) Indication of overheating if the off manually to override the automatic accurate indication in all phases of flight and
construction of the component is such, (e.g. control (see Part 1 Section 4 Clause 4.19, associated aeroplane configurations;
by bonding process), that overheating would for location of manual controls), and
be a serious hazard. (c) if they project into the air stream,
(d) be capable of operation present the least obstruction to airflow and
(e) Indication of the severity of the throughout the performance range of the be constructed to withstand damage from
icing conditions, unless this is shown to be aeroplane without causing damage, or, the impact of ice or slush shed from other
unnecessary or impractical, and its omission alternatively, be so designed that it cannot parts of the airframe;
is agreed with the Project Team Leader operate automatically in conditions when
(PTL). damage would result. (In such cases, the (d) in the case of accretion type
pilot shall be given warning that, if he detector probes, be positioned with due
operates the system manually, damage will regard to the probable trajectories of ice
result.) and, shed from them in order to obviate damage
to engines, or other systems or
(e) incorporate back-up systems components; and
where flight safety may be impaired by the
failure of primary systems. (e) have a response time such that the
safety of the aeroplane, and the correct
functioning of its engines and systems, are
not jeopardised due to ice accretion.
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Ice detectors, and temperature or thermostat

probes if used, shall be designed for minimum
thermal lag.
DESIGN AND CONSTRUCTION
1.5.1.28. The ice protection system shall The contractor shall show by ground and flight The system shall be designed with due regard to:
conform to the conditions and testing tests that the requirements of this Clause are
requirements defined in Def Stan 00-35 (Multipart) complied with, and in particular that: (a) reliability,

(a) sufficient ice protection is provided (b) availability,

for the icing conditions and durations
required by the specification, (c) maintainability (including inspection
and servicing, see Part 1 Section 4 Clause
(b) the ice detector will detect, and 4.4), and
respond rapidly to, icing conditions,
(d) testability, and shall be shown to
(c) the automatic controls function meet the requirements of the Aeroplane
satisfactorily when flying into and out of Specification in these respects.
icing conditions,
Where flight testing of untried systems or
(d) automatic system operation does applications may involve some risk, consideration
not occur in conditions where this might shall be given to first conducting ground tests on
hazard the aeroplane or system, and representative test specimens using rigs, icing
tunnels or other suitable facilities.
(e) significant failures of the
protection system are indicated to the pilot
and any back-up systems operate
satisfactorily at such times

Testing shall be performed in accordance

with the provisions of Clause 1.5 on an
aeroplane designated by the Project Team
Leader (PTL). With the ice protection
system installed and working, the handling
and performance requirements of Part 1
Section 2 Clause 2.20 must be complied
with.
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1.5.1.29. The protection system and its
installation in the airframe shall meet the relevant
static strength and fatigue damage tolerance
requirements of Part 1 Section 3 Clause 3.1 and
3.2, in addition to any life requirements contained
in the Aeroplane Specification.
1.5.1.30. All electrical installations shall be
designed in accordance with Part 1 Section 6
Clause 6.6
1.5.1.31. Materials used must not be
adversely affected by the de-icing or anti-icing
medium.
1.5.1.32. Temperature limiting devices,
insulation, or other methods shall be used, as
necessary, to prevent any system or component of
the aeroplane from exceeding any of the
following:-

(a) A surface temperature that would

be hazardous to the occupants.

(b) Its design temperature limitations.

(c) The self-ignition temperature of any

flammable material (solid or liquid) to which
it may be exposed either normally or
accidentally.
1.5.1.33. Any insulation which is not naturally
impervious to combustible fluids and which is used
in areas where combustible fluids are present
shall be covered and sealed with abrasion and
fluid resistant covering.
1.5.1.34. If ground checking requires the use
of external supplies or test equipment, then
suitable test connection facilities shall be provided
within the system.

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1.5.1.35. The internal diameter of any filling See BS 2C 13
orifice shall not be less than 45 mm.
TESTING
1.5.1.36. The ice protection systems under The tests described in this Clause are applicable The object of the test of this Clause is to
test shall be fully representative of the production to all aeroplanes fitted with ice protection demonstrate that ice protection systems where
standard aeroplane. system(s). Tests are required to evaluate each provided, protect the aeroplane in all the icing,
new aeroplane/system type and existing snow and mixed conditions in which the
aeroplanes where modification action may have aeroplane is expected to operate.
degraded its capability in icing conditions. In Particular emphasis is placed on establishing the
each case the aeroplane and system or adequacy of anti-icing and de-icing measures in
modification shall be fully representative of the relation to vulnerable areas. Testing does not
Service standard. cover general clearance/handling for flight in icing
conditions, whether or not they are fitted with ice
protection systems.

 Protected areas include:

 aerofoils
 moving surfaces (slats/flaps)
 propellers
 engine and auxiliary intakes
 pitot and static heads and masts
 aerials and sensors
 radomes
 AEW rotating scanners
 weapons and weapon carriers
 transparencies
 vents and drains

and other specified items - as applicable.

INSTRUMENTATION
1.5.1.37. The following requirements shall be Care shall be taken to ensure that the Instrumentation will, to some extent depend upon
met: instrumentation does not hazard the aeroplane in the assessment of risk related to flying in icing
its own right, or modify the ice accretion conditions. This will be based upon calculation,
(a) The aeroplane shall be characteristics of the aeroplane under test. modelling and rig (icing tunnel etc.,) tests and the
instrumented so that a continuous record is extent of icing introduced.
obtained of aeroplane flight, engine and
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protection system parameters.

(b) Meteorological conditions shall be

recorded (outside air temperature and icing
severity)

(c) Visual monitoring of accumulations

of ice in critical areas shall be provided
during the trial with video or cine recording
where warranted.

(d) An ice accretion indicator will be

provided for the pilot, to enable identification
of the onset of airframe icing, ice type and
amount.
GENERAL TEST CONDITIONS AND REQUIREMENTS
1.5.1.38. The ice protection systems under The manufacturer’s protection system limitations The Service operational requirements need to be
test shall be fully representative of the production for the equipment under test and the evidence clearly defined in terms of, severity of snow and
standard aeroplane. from any earlier testing in icing conditions shall icing encounters to be considered, duration of any
Ground and flight conditions shall be fully be sought. and each ice encounter and the number of icing
representative of the Service role of the aeroplane encounters per flight. The level of acceptable
and the range of climatic conditions for which the performance degradation, for aeroplane or system
clearance is required. shall be specified in the requirement.
GROUND TESTS
1.5.1.39. Cold soak ground tests shall be Consideration shall be given to the need for tests
conducted where the windscreen has been to establish the onset of intake icing for engines
allowed to mist or ice over to assess the adequacy and auxiliary power units with and without guards
of the ice protection system (and standby system fitted. To minimise risk during subsequent flight
where applicable) to maintain clear vision testing, consideration shall be given to conducting
throughout the pre-take-off period. a series of tests on the ground using an air stream
icing facility. Advice may be obtained from the
Project Team Leader. Aspects that might usefully
be addressed are:

(a) The ability of the ice detector to

detect icing conditions and the response
time of those conditions.
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(b) Functioning of the pitot/static

heaters under icing conditions.

(c) Operation of stall warning sensor

and heater under icing conditions.

(d) Functioning of the propeller de-

icing system under normal and single failure
conditions.

(e) Functioning of the engine air intake

anti-icing system under icing conditions.

(f) Ice shedding from parts of intake

etc., to assess the possibility of significant
ice quantity being ingested by the engine.

(g) The effect of ice accretion on the

operation of primary and secondary flying
controls.

(h) Assessment of clear vision for the

pilot(s) at circuit flying speeds.

(i) Assessment of the ice formation on

external tanks/weapons and weapon
carriers and the impact this has on jettison
capability.

(j) Initial assessment of the possible

effects of frozen slush spray on the
operation of undercarriage and under-
carriage doors.
1.5.1.40. Taxying tests over snow/slush shall Consideration shall then be given to the possible
take place to determine the extent of airframe effects of the resultant spray freezing and
contamination by spray thrown up by wheels and causing interference with the normal operation of
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engine efflux/propeller slipstream (especially the undercarriage, undercarriage doors, flaps
during reverse thrust braking). and flying controls and auxiliary intakes during or
These tests shall be carried out in the take-off and after take-off. The effect of external stores (where
landing configurations. applicable) or contamination/ice accumulation
shall be determined.
1.5.1.41 Following operation of both anti-
icing and de-icing systems on the ground there
shall be an assessment of the extent of run-back
of melted snow/ice into hinges for control surfaces
where subsequent refreezing could occur.
1.5.1.42. Tests shall be conducted to Tests should take place in:
determine the effect on the aeroplane structure
and other systems resulting from deliberate or (a) in sub-zero temperatures;
inadvertent operation of the ice protection
system(s) on the ground (except where (b) at other temperatures,
automatically inhibited).
(c) in snow conditions.
FLIGHT TESTS
1.5.1.43. Aspects to be addressed during these trials are Icing trials with a hitherto untried aeroplane shall
as follows: be approached with caution since it is a potentially
high-risk area. The planning of individual flight
(a) The ice detector and ice protection tests shall take full account of the hazards,
systems shall be operated in clear air in the particularly with a single engine aeroplane.
required temperature range to confirm Consideration shall be given to the use of an
satisfactory operation of the system, and at airborne tanker* which can dispense the required
higher temperatures where appropriate to concentrations of liquid water content over the
establish any adverse effects. required area of the test aeroplane: these tests
can be carried out in clear air at any time required
(b) Flight shall take place in light icing in the test programme and the test conditions can
conditions to establish initially correct be repeated at will. However, it is unlikely that the
functioning of the ice detector and to icing produced with a tanker will be fully
establish the extent of ice build up prior to representative of natural icing conditions. The
an icing warning. Observation shall be optimum solution is therefore, likely to be a
maintained to establish any other cues that combination of icing tanker and natural icing trials,
the aeroplane has entered icing conditions, bearing in mind the expense and time involved in
for example, ice formation on the outside of testing under natural conditions.
the windscreen, windscreen wipers (when * A&AEE, Boscombe Down no longer have an
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fitted) and/or structure in the immediate icing tanker. A number of tankers of varying sizes
field of view of the pilot. are available in the USA.
Progressively, checks shall be made to
establish:

(1) that normal and standby

windscreen protection systems provide
adequate clear vision for the pilot.

(2) the adequacy of the protection

afforded to air intake and engine.

(3) the effect of icing conditions on

operation of sensors, for example,
pitot/static heads, ADA sensor, stall
warning sensors, aerials etc.

(4) the rate and position of ice

accretion and the effect on the airframe.

(c) As confidence: in the

effectiveness of the ice protection system is
gained, testing can proceed into
increasingly more severe icing conditions
for longer periods and in snow, ice crystals,
freezing rain and mixed conditions as
applicable to the operational requirement.
An additional aspect to monitor at this
stage is the 'run-back' of melted ice and its
effects.

(d) Tests shall be conducted to

establish the effects of partial and total
system failures, followed by a cautious
assessment of the unprotected aeroplane
during escape from icing conditions.

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B - Rotorcraft
1.5.1 For Ice Protection see Part 7,
Leaflet 711-0, Leaflet 711-2, Leaflet 711-3

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1.6 SURVIVABILITY AND RECOVERY
A – Fixed Wing
SONAR LOCATING BEACONS (SLB) AND EMERGENCY LOCATOR TRANSMITTERS (ELT)
See Section 1.3 Clauses 1.3.1 to 1.3.3.7
B - Rotorcraft
1.6.1 For Sonar Locating Beacons and
Emergency Locator Transmitters see Part 7,
Leaflet 100-00
1.6.2. LIFE JACKETS
A – Fixed Wing
1.6.2 Refer to IR OPS CAT.IDE.A.285
B - Rotorcraft
1.6.2 Refer to IR OPS CAT.IDE.H.290
1.6.3. CREW SURVIVAL SUITS
A – Fixed Wing
1.6.3 N/A at this time.
B - Rotorcraft
1.6.3 Refer to IR OPS CAT.IDE.H.295
1.6.4. LIFE RAFTS AND SURVIVAL ELTS FOR EXTENDED OVERWATER FLIGHTS
A – Fixed Wing
1.6.4 Refer to IR OPS CAT.IDE.A.280
B - Rotorcraft
1.6.4 Refer to Part 7, Leaflet 721-00,
Leaflet 721-1
1.6.5. SURVIVAL EQUIPMENT
A – Fixed Wing
1.6.5 Refer to IR OPS CAT.IDE.A.305
B - Rotorcraft
1.6.5 Refer to IR OPS CAT.IDE.H.305
1.6.6. EMERGENCY FLARES
A – Fixed Wing

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SIGNAL PISTOLS
1.6.6.1. A Signal pistol shall be fitted in all (a) The pistol shall be placed so as to fire
RAF aeroplanes except fighters, bombers and upwards and so that:
trainers, and in all multi-seat naval aeroplanes
required for anti-submarine, search and rescue (1) in high altitude aeroplanes, it can
and troop carrying duties. be operated, if necessary by remote
control, from the pressure cabin, and

(2) in maritime aeroplanes, the fuze

trail after ejection does not enter the field of
view of the camera.

(b) The pistol shall be so installed that either:

(1) it can be fired safely in situ, or

(2) if not installed in the firing position,

it is impossible to insert a loaded pistol in
its stowage.
B - Rotorcraft
1.6.6 See Requirement 1.6.5 Part B
Rotorcraft. (Referenced Para a)
1.6.7. FIRST AID EQUIPMENT
A – Fixed Wing
1.6.7 Refer to IR OPS CAT.IDE.A.220
B - Rotorcraft
1.6.7 Refer to IR OPS CAT.IDE.H.220
1.6.8. EMERGENCY MEDICAL KIT
A – Fixed Wing
1.6.8 Refer to IR OPS CAT.IDE.A.225
B - Rotorcraft
1.6.8 N/A to Rotorcraft at this time
1.6.9. SEATING AND NO SMOKING SIGNS
A – Fixed Wing
DESIGN OF SEATS AND HARNESSES
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1.6.9.1. Crew seats, harnesses, and their The strength requirements for seats and
attachments shall meet the crash landing and harnesses of other occupants are given in Section
ditching requirements of Part 1 Section 4 Clause 4 Clause 4.21 and Leaflet 76.
4.22 unless fully factored flight loads are greater. The requirements for ejection seats are given in
Section 4 Clause 4.23
1.6.9.2. Seats with provision for fitting the
equipment detailed in the Aircrew Equipment
Assembly schedule (AEA) for the aeroplane shall
be provided at all crew stations where ejection
seats are not required.
1.6.9.3. Any seat which might not be
occupied during flight shall be provided with a
means of preventing the movement of articles
attached to or associated with the seat and which
normally remain in the aeroplane. The stowage
arrangements shall ensure the security of the
articles under all conditions and manoeuvres
throughout the flight envelope.
1.6.9.4 If smoking is prohibited, there shall
be a notice stating so, and if smoking is to be
allowed:

(a) There shall be an adequate number

of self-contained removable ashtrays.

(b) Where the crew compartment is

separated from the passenger compartment
there shall be at least one sign meeting the
requirements of Part 1 Section 4 notifying all
passengers when smoking is prohibited.
SMOKING
1.6.9.5. See IR OPS CAT.IDE.A.210
B - Rotorcraft
1.6.9 No Smoking – Refer to Part 7
Leaflet 712-00 Para 7.5, Leaflet 714 Para 2.2.2.
Seat Belts – Refer to Part 7 Leaflet 714 Para 2.1.6
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1.6.10. INTERNAL DOORS AND CURTAINS
A – Fixed Wing
1.6.10 Refer to IR OPS CAT.IDE.A.215
B - Rotorcraft
1.6.10 Curtain: Refer to Part 7 Leaflet 105
Para 16 N/A for Internal Doors
1.6.11. FLOODLIGHTING
A – Fixed Wing
1.6.11.1. Floodlighting shall be installed at all The colour of this light shall be white, unless red See ASCC Air Standard 61/113/2H and STANAG
crew stations to provide general and standby is specified by the Project Team Leader. 3224
illumination of the instruments and console panels.
EMERGENCY LIGHTING
1.6.11.2. In those aeroplanes not fully The colour of this light shall be white, unless red See ASCC Air Standard 61/113/2H) and STANAG
safeguarded against electrical failure, an is specified by the Project Team Leader. 3224
independent system of floodlighting shall be
installed to illuminate the essential instruments at
the pilot's station. It shall operate on a single self-
contained circuit and derive its power from an
independent emergency battery.
1.6.11.3. The circuit shall be controlled by a See Part 1 Section 7 Clause 7.4.5 for the marking
master switch readily accessible to the pilot. of the switch.
EVACUATION
1.6.11.4 Emergency lighting, escape
identification and markings, shall be crashworthy
to the requirements of Part 1 Section 4 Clauses
4.22.44 or 4.22.47 as appropriate.
EMERGENCY ESCAPE/EVACUATION ILLUMINATION
1.6.11.5. The following requirements shall See ASCC Air Standard 61/113/09A (Was 10/66)
apply in the design and location of power supplies,
controls, lighting fixtures and associated
equipment used to provide emergency escape
illumination in aeroplanes:

(a) An emergency lighting system,

independent of the main lighting system, must be

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installed. However, the sources of general cabin
illumination may be common to both the
emergency and the main lighting systems if the
power supply to the emergency lighting system is
independent of the power supply to the main
lighting system. The emergency lighting system
must include:

(1) Illuminated emergency exit marking

and locating signs, sources of general cabin
illumination, interior lighting in emergency
exit areas, and floor proximity escape path
marking.

(2) Exterior emergency lighting.

(b) Emergency exit signs - For aeroplanes that See Def Stan 00-970 Pt 1 Section 7 clause 7.4
have a passenger seating configuration, excluding See Def Stan 00-970 Part 13 Section 1.6
pilot seats, of 9 seats or less, each emergency exit
and external door in the passenger compartment
must be internally marked with the word “exit” by a
sign which has white letters 25 mm (1 in) high on
a red background 51 mm (2 in) high, be self-
illuminated or independently, internally-electrically
illuminated, and have a minimum brightness of at
least 0.51 cd/m2 (160 microlamberts). The colour
may be reversed if the passenger compartment
illumination is essentially the same.

(c) General illumination in the passenger

cabin must provide an average illumination of not
less than 0.5 lux (0.05 foot-candle) and an
illumination at any point of not less than 0.1 lux
(0.01 foot-candle) when measured along the
centre line of the main passenger aisle(s) and at
the seat armrest height

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(d) Floor proximity emergency escape path

marking must provide emergency evacuation
guidance for the aeroplane occupants when all
sources of illuminations more than 1.2 m (4 feet)
above the cabin aisle floor are totally obscured.

(e) When certification to the emergency exit

provisions is requested, the following shall apply:

(1) An emergency lighting system,

independent of the main cabin lighting
system, must be installed. However, the
source of general cabin illumination may be
common to both emergency and main
lighting system if the power supply to the
emergency lighting system is independent of
the power supply to the main lighting
system.

(2) The lights must be operable

manually from the flight crew station and
from a point in the passenger compartment
that is readily accessible to a normal cabin
crew member seat.

(3) There must be a flight crew warning

light, which illuminates when power is on in
the aeroplane and the emergency lighting
control device is not armed.

(4) The cockpit control device must

have an ‘on’, ‘off’ and ‘armed’ position so
that when armed in the cockpit or turned on
at either the cockpit or cabin crew member
station the lights will either light or remain

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lighted upon interruption (except an
interruption caused by a transverse vertical
separation of the fuselage during crash
landing) of the aeroplane’s normal electric
power. There must be a means to safeguard
against inadvertent operation of the control
device from the ‘armed’ or ‘on’ positions.

(5) There must be a crew warning light

that illuminates in the cockpit when power is
on in the aeroplane and the emergency
lighting control device is not armed.

(6) The emergency lights must be

operable manually from the flightcrew station
and be provided with automatic activation.
The cockpit control device must have "on,"
"off," and "armed" positions so that, when
armed in the cockpit, the lights will operate
by automatic activation.

(7) There must be a means to

safeguard against inadvertent operation of
the cockpit control device from the "armed"
or "on" position.

(f) The energy supply to each emergency

lighting unit must provide the required level of
illumination for at least 10 minutes at the critical
ambient conditions after activation of the
emergency landing system.

(g) If rechargeable batteries are used as the

energy supply for the emergency lighting system,
they may be recharged from the main electrical
power system of the aeroplane provided the

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charging circuit is designed to preclude
inadvertent battery discharge into the charging
circuit faults. If the emergency lighting system
does not include a charging circuit, battery
condition monitors are required.

(h) Components of the emergency lighting

system, including batteries, wiring relays, lamps,
and switches must be capable of normal operation
after having been subjected to the inertia forces
resulting from the ultimate load factors prescribed
in Part 1 Section 3 Clause 3.1

(i) The emergency lighting system must be

designed so that after any single transverse
vertical separation of the fuselage during crash
landing –

(1) At least 75 percent of all electrically

illuminated emergency lights required by this
paragraph remain operative; and

(2) Each required electrically

illuminated exit sign remains operative
exclusive of those that are directly damaged
by the separation.

(j) Emergency escape illumination shall be

designed so that no beam of light is directed into
occupants' eyes in such a way as to compromise
their ability to escape.
1.6.11.6. Emergency escape illumination (see Part 1 Section 4 Clause 4.23)
shall be provided independent of normal power
sources.
B - Rotorcraft
1.6.11 Refer to Part 7, Leaflet 105-00 Para

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15.5, Leaflet 307-77 Para 4.15.2, Leaflet 706-00
Para 2.7.3, Leaflet 714 Para 2.1.5
1.6.12. HAND FIRE EXTINGUISHERS
A – Fixed Wing
1.6.12.1. In all multi-seat aeroplanes and in (a) One hand extinguisher shall be easily
single seat high speed aeroplanes carrying accessible from outside as well as inside the
photographic reconnaissance or specialised aeroplane and its location shall be marked in
electronic equipment, hand operated fire accordance with the DAP119A-0601-Series.
extinguishers complying with AFNOR NF EN 3
Parts 1-6, BS 7863 and BS 7867 shall be provided (b) It is desirable that this hand extinguisher
to a total scale of one to each crew member, be stowed together with, or adjacent to, the
positioned at each normal crew station. fireman's axe and heat resisting gloves called for
in Clause 1.6.13.2.
1.6.12.2. There shall be at least the following (a) The number and location of hand held
number of hand held fire extinguishers Fire extinguishers shall be such as to provide
conveniently located in passenger compartments: adequate availability for use, account being taken
Passenger Capacity Minimum Number of the number and size of the inhabited
of Hand Held Fire compartments and the location of toilets and
Extinguishers galleys etc. These considerations may result in
the number being greater than the minimum
7-30 1 prescribed. Hand held fire extinguishers shall
31-60 2 comply with Clause 1.6.12.4
61 or more 3
(b) Where a built-in fire extinguishing system
is used it shall comply with Clause 1.6.12.5
1.6.12.3. There shall be at least the hand
held fire extinguisher suitable for both flammable
fluid and electrical equipment fires conveniently
located on the flight decks of aeroplanes normally
occupied by 2 or more persons.
1.6.12.4. For hand held fire extinguishers the
following apply:

(a) Each hand held fire extinguisher

including bracket must be approved to
AFNOR NF EN 3 Parts 1-6, BS 7863 and
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BS 7867).

(b) The types and quantities of

extinguishing agent used shall be
appropriate to the kinds of fires likely to
occur and the materials involved. The use of
Methyl Bromide is prohibited.

(c) Hand held fire extinguishers shall

be easily accessible and mounted to
facilitate quick removal, and their locations
marked in accordance with the DAP119A-
0601-Series.

(d) Extinguishers used in personnel

compartments shall be designed to minimise
the hazard of toxic gas concentrations.

(e) The design of the extinguisher

bracket shall be such to minimise the
possibility of inadvertent opening of the
release by snagging of clothing etc.
1.6.12.5. If a built-in fire extinguishing system
is specified:

(a) The capacity of each system, in

relation to the volume of the compartment
where used and the ventilation rate, shall be
adequate for any fire likely to occur in that
compartment.

(b) Each system shall be installed so

that no extinguishing agent likely to enter
personnel compartments will be hazardous
to the occupants.

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(c) Discharge of the extinguisher shall
not cause structural damage.

(d) The use of Methyl Bromide is

prohibited.
1.6.12.6 Refer to IR OPS CAT.IDE.A.250
Para’s c, d and f.
B - Rotorcraft
1.6.12 Refer to Part 7, Section 1, Leaflet
105-00 Para 20 and Section 7, Leaflet 714 Para
2.1.7
1.6.13. CRASH AXE OR CROWBAR AND HEAT RESISTING GLOVES
A – Fixed Wing
1.6.13.1. In multi-seat aeroplanes, fireman’s The number of stowages will be decided not later
axes and stowages shall be provided according to than at the Mock-up Conference.
the number and position of crew stations.
1.6.13.2. Stowage shall be provided on all (a) The stowage for one fireman's axe and
multi-seat aeroplanes for a pair of heat resistant the heat resisting gloves shall be accessible and
gloves. readily identifiable from both inside and outside
the aeroplane and its location shall be marked in
accordance with the DAP119A-0601-Series.

(b) It is preferable that these items shall be

together in one stowage (Clause 1.6.12.1.)
B - Rotorcraft
1.6.13 Refer to Part 7, Section 1, Leaflet
105-00, Para’s 20 and 21
1.6.14. MARKING OF BREAK IN POINTS
A – Fixed Wing
1.6.14.1. See Clause 1.6.15.3
B - Rotorcraft
1.6.14 Refer to IR OPS CAT.IDE.H.260
1.6.15. MEANS FOR EMERGENCY EVACUATION
A – Fixed Wing
1.6.15.1. The approach to the exits and the (a) To facilitate passage to the exits, ladders
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exits themselves shall be free from obstructions and foot and/or hand holds of rigid construction
and projections on which clothing, parachute shall be provided wherever necessary.
harnesses or other items of personal equipment
might be caught. (b) Ladders, when in position, should slope
upwards from aft to forward. If not permanently
located they shall be capable of being rapidly
secured in position.
NUMBER OF EXITS
1.6.15.2. The number of exits provided and It is acceptable, and often desirable, that doors,
their situations shall be such that not more than hoods, hatches, etc., primarily provided for other
three persons need to use any one exit in any reasons (e.g. the normal entrance door) should
given emergency, except that, on transport serve also as emergency exits. Any exit may
aeroplanes, the number of exits shall be related to serve for more than one emergency condition
the number of passengers (e.g. 1: 30) provided that the requirements stated for each
such condition are met.
MARKING OF EXITS
1.6.15.3. All emergency exits shall be Details of the required markings are given in Part STANAG 3230 - Emergency Markings on Aircraft.
adequately marked so that their intended use and 1 Section 7 Clause 7.4
their means of operation are quite obvious to the
occupants of the aeroplane and also, where
appropriate, to rescue personnel approaching the
aeroplane from outside.
SIZE OF EXITS
1.6.15.4. Each exit shall be of the largest
practicable size and shall in every case be such
as to give a freedom of passage not less than that
provided by a rectangular opening 609.6 mm x
609.6 mm. To permit the evacuation of injured
personnel from multi-seat aeroplanes, of the exits
provided at least one for every nine occupants
shall give a freedom of passage not less than that
provided by a rectangular opening 762 mm x 762
mm.
CONTROLS
1.6.15.5. Each emergency exit shall be (a) When the exit is jettisonable directly Def Stan 00-250 and Part 1 Section 4 Leaflet 63
openable and jettisonable, when applicable, by outwards the control shall be such that there is cover anthropometric data.
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one hand by a single positive movement of a no risk of the operator's hand being pulled
single control, operated by a pull of between 111 outward by the cover.
N and 178 N.
(b) The control for jettisoning the pilot's hood
shall be in accordance with Part 1 Section 4
Clause 4.19 and shall be operable under all
conditions specified in Part 1 Section 4 Clause
4.23.1

(c) In addition to the requirements of Clause

1.6.15.6. consideration should be given to the
operation of external controls by members of
crash/rescue crews who may be wearing bulky
protective clothing. Crash rescue crew members
should be able to operate the external controls
whilst wearing the full protective clothing outfit.
For this purpose the Design Organisation should
use Anthropometric data for Metacarpal Breadth
relative to 95th percentile man, making due
allowance for the wearing of protective
gloves/gauntlets. The Project Team Leader shall
be approached to confirm details of the protective
clothing outfit to be catered for unless this
information is provided in the Aircraft
Specification.
1.6.15.6. On aeroplanes with ejection seats, All emergency controls shall be operable with the
the following controls, which shall be so shaped man strapped into the seat with the harness
and positioned to avoid any possible chance of retracted and locked; this should also apply when
confusion in their operation, shall be provided: upper limb restraint is installed on the seat.

(a) a control for jettisoning or

fragmenting the hood or hatch conforming to
Clause 1.6.15.5 Where Miniature Detonating
Cord (MDC) is used to fragment the hood
the separate firing control shall permit the
firing of the MDC independently of the seat

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for both internal and external operation and
provision shall be made to reduce the
likelihood of injury to an outside rescuer from
hood debris, e.g. by pulling a long lanyard,

(b) a single control, operated by a pull

of between 111 N and 289 N to initiate the
entire escape sequence, so arranged that
the trajectory of the occupant, his seat and
all personal equipment is automatically
cleared of dangerous obstructions,

(c) a single control, to enable the

occupant to separate himself from the seat
after ejection with parachute and personal
survival pack intact should automatic
separation fail to occur, and to permit
manual escape should this be necessary,

(d) a "go-forward" lever in accordance

with Part 1 Section 4 Clause 4.19,

(e) a control to permit the occupant to

raise or lower the seat in accordance with
Part 1, Section 4, Clause 4.19,

(f) a control for manual operation of

the emergency oxygen set in accordance
with Part 1, Section 4, Clause 4.19 (see also
Clause 1.4.4.33), and

(g) if a seat mounted oxygen system is

fitted, all oxygen controls, located adjacent
to the package. These controls shall include:

(1) an air mix/100% oxygen control

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(this control may however be fitted on a
console),

(2) a bypass regulator control, and

(3) the necessary test buttons,

(h) combined services release on the

port side of the seat.
ESCAPE PATH CLEARANCE
DESIGN REQUIREMENTS
1.6.15.7. The following human factors criteria See ASCC Air Standard 61/102/04A.
shall be applied during the design and evaluation Escape Path and Escape Path Clearance are
of escape path clearance mechanisms: defined in Part 0.

(a) The escape path shall permit the

safe egress of the most critical combination
of aircrew and equipment specified for use
with that escape system.

(b) The escape path clearance

mechanisms should minimize the risk to
aircrew and their equipment.

(c) The various potential environmental

hazards to which the aircrew might be
exposed on the escape path or due to the
clearance mechanisms shall be controlled to
be compatible with established human
exposure limits. Depending upon the
method used, these potential hazards may
include overpressure, acoustic noise, flame,
fragmentation and others.

(d) Failure of the escape path

clearance system shall not prevent escape

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nor expose the crew to undue risk of
unacceptable injury.

(e) The method of escape path

clearance should produce minimal
interference with the crew tasks.
ESCAPE FROM AEROPLANE ON THE GROUND
1.6.15.8. For Cat B and C Aircraft it shall be
demonstrated in accordance with CS 25-803 and
CS-23-803 that a full complement of crew and
passengers can be evacuated from the aeroplane
to the ground within 90 seconds.
ESCAPE FROM THE AEROPLANE ON WATER
1.6.15.9 For Cat B and C Aircraft it must be
shown that, under reasonably probable water
conditions, the flotation time and trim of the
aeroplane will allow the occupants to leave the
aeroplane and enter the provisioned life rafts. If
compliance with this provision is shown by
buoyancy and trim computations, appropriate
allowances must be made for probable structural
damage and leakage. If the aeroplane has fuel
tanks (with fuel jettisoning provisions) that can
reasonably be expected to withstand a ditching
without leakage, the jettisonable volume of fuel
may be considered as buoyancy volume.
B - Rotorcraft
1.6.15. Refer to Part 7, Leaflet 102-00,
Leaflet 102-1, and Leaflet 102-2 (Emergency
Escape).
1.6.16. MEGAPHONES
A – Fixed Wing
1.6.16 Refer to IR OPS CAT.IDE.A.270
B - Rotorcraft
1.6.16 Refer to IR OPS CAT.IDE.H.270

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1.6.17. PERSONAL SURVIVAL PACKS
A – Fixed Wing
1.6.17.1. Stowage space shall be provided (a) The packs shall be capable of being
for the carriage of Personal Survival Packs, easily installed and shall be readily removable
together with the fittings necessary for their after a wheels-up landing.
attachment, on a basis of one for each member of
the crew. (b) Personal Survival Packs are intended for
use after a forced landing and their installation
must not prejudice the fitting of life rafts or life
saving equipment intended for use after a
ditching.
B - Rotorcraft
1.6.17.1 Refer to Part 7, Leaflet 105-00 Para
18

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1.7 SAFETY RELATED PGRAMMABLE ELEMENTS
1.7.1 For the requirements, design, To meet the stated requirement, Compliance is Guidance for this requirement is provided in four
verification and validation of safety related provided in four sections: system level safety sections mirroring those for compliance.
Programmable Elements (PE) refer to Def Stan considerations; airworthiness related cyber
00-55 Issue 3. security; Safety Related Software (SRS); and
Complex Electronic Hardware (CEH).

(a) For the assurance of system level (a) Guidance for system level safety
safety considerations: considerations:

At the system level, the Safety Assessment Civil system developers apply ARPs to ensure
process should define the top level safety that the system design is failure tolerant and that
requirements and design objectives of the PE as a catastrophic failure condition (e.g. loss of
detailed in the guidance contained within aircraft) should not result from the failure of a
Aerospace Recommended Practices (ARPs) critical function implemented in a PE component.
4761 and 4754A. The associated Safety Assessment process
should define the top level safety requirements
and design objectives of the PE as detailed in the
guidance contained within Aerospace
Recommended Practices (ARPs) 4761 and
4754A.

All aspects of the PE should be supported by a As required by Def Stan 00-56 Issue 5, the Safety
Safety Assessment Report as described within Assessment Report should provide a complete,
Def Stan 00-56 Issue 5. evidence-based, robust, compelling, documented
and auditable argument for all aspects of the
safety related PE including providing evidence
that the criticality of any previously developed PE
remains valid when used within the context of the
Military Air Environment (MAE).

Both the Safety Assessment process and

resulting Safety Assessment Report activities
should also be cognizant of security assessment
requirement detailed below.

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REQUIREMENT COMPLIANCE GUIDANCE

(b) For airworthiness related cyber (b) Guidance for airworthiness related
security assurance: cyber security:

RTCA DO-326A/EUROCAE ED-202A and It is necessary to ensure that platform cyber

associated RTCA DO-356/EUROCAE ED-203 security vulnerabilities do not purposefully or
combined with arguments made against JSP 440 accidentally threaten airworthiness.
should be used as an acceptable means of In keeping with threats to the continued safe
compliance with the cyber security requirements operation of SRS and CEH, Def Stan 00-55 Issue
of Def Stan 00-55. 3 places a requirement to demonstrate that
potential cyber security threats to safe operation
are mitigated. Def Stan 00-55 highlights that JSP
440 provides guidance on security policy but the
latter does not specifically provide AMC for design
assurance of the security aspects of
airworthiness, therefore a combined approach is
required.
It is recognised that DO-326A/ED-202A has been
developed for use on large civil aircraft. As such,
some tailoring of the guidance provided therein
may be required for military aircraft and the
military environment.
As is the case for conventional software
assurance, the level of airworthiness-related
security assurance should be commensurate with
the risk associated with failure. Usefully, some of
the activities associated with safety assurance
and airworthiness-related security overlap, it is
therefore recommended that an integrated and
coherent approach is taken to reduce
unnecessary overheads.
Due to the evolving nature of cyber security
threats, where airworthiness-related security risks
are identified, it would also be anticipated that a
continuing airworthiness-related security strategy
(for example, as described in RTCA DO-
355/EUROCAE ED-204, Information Security
Guidance for Continuing Airworthiness) would be
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REQUIREMENT COMPLIANCE GUIDANCE

implemented.

(c) For Safety Related Software (SRS) (c) Guidance for Safety Related Software
assurance: (SRS):

RTCA DO-178C and its appropriate supplements The guidance in Def Stan 00-55 Issue 3 on the
(DO-248C; DO-330; DO-331; DO-332; and DO- adoption of the DO-178 family identifies additional
333), can be considered to be an acceptable considerations relating to governance and
means of compliance to provide design shortfalls against the Def Stan 00-55
assurance of airborne SRS when supported by a requirements; these should be addressed along
robust, documented and auditable Safety with any ‘military delta’ particular to the
Assessment as described within Def Stan 00-56 application.
Issue 5.

For legacy software which is intended to be used For legacy software which is intended to be used
in a new application, or as a significant in a new application, or as a significant
development of an existing system, the following development of an existing system the
principles apply: acceptability of remaining with the legacy means
of compliance is based on the principle that
(i) For systems developed under Def Stan switching development activities to a different
00- 55 Issue 2, it may continue to be applied standard may inherently increase the risk of
as an acceptable means of compliance introducing errors into the software due to
provided the requirements of that standard applicants applying unfamiliar processes,
continue to be met; and methods or techniques. Should this not be an
issue for the applicant, it is acceptable to switch to
(ii) For software developed using RTCA the current acceptable means of compliance (i.e.
DO- 178B, it may continue to be used as an DO-178C) provided that a complete and coherent
alternative means of compliance under the assurance argument can be maintained for all of
following circumstances: the SRS.
When considering the use of software previously
a. The new application does not developed for civilian applications using civil
require a higher level of software aviation standards, including RTCA DO-178C, the
assurance; applicant should note that some SRS components
applied in a MAE would require additional
b. The development cycle is not mitigation e.g. additional functional, design or
updated to include technologies that have physical independence. Where the appropriate
specific supplements in DO-178C; functional, design or physical independence
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REQUIREMENT COMPLIANCE GUIDANCE

cannot be obtained, an alternate military SRS
c. No new software criteria 1 or 2 (as system design should be sought with either a
defined in DO-178C) tool qualification is higher level of assurance chosen or the civil
required. If this is the only differentiator standard applied but with additional assurance
then DO-178B can continue to be applied methods in order to gain the necessary level of
with the tool qualification objectives confidence to meet the requirements of Def Stan
provided by DO-330 being used for the 00-55 Issue 3.
new tools; The re-use of previously developed Def Stan 00-
55 Issue 2 or DO-178B (and DO-178A) SRS
d. No new Parameter Data Item files within a new or existing military airborne system
(as defined in DO-178C) are introduced. can only be considered to be acceptable to the
Where this is the case, DO-178C should authority on a case by case basis and should be
be applied to all affected areas of the supported by documented evidence and a full
software and an argument developed in audit trail of the development history of the SRS.
the supporting safety case to show that
the change has been contained. Where
this is not feasible DO-178C should be
applied; and

e. All of the lifecycle processes and

artefacts from prior certification have
been maintained.

(d) For safety related Complex Electronic (d) Guidance for safety related Complex
Hardware (CEH) assurance: Electronic Hardware (CEH):

RTCA DO-254/EUROCAE ED-80 can be This element of the requirement focuses on safety
considered to be an Acceptable Means of related Complex Electronic Hardware (CEH), also
Compliance to provide design assurance of known as complex custom micro-coded
airborne safety related CEH when supported by a components. These include: Application Specific
robust, documented and auditable Safety Integrated Circuits (ASIC); Programmable Logic
Assessment as described within Def Stan 00-56 Devices (PLD); Field Programmable Gate Arrays
Issue 5. (FPGA); and other similar electronic components
or devices. In keeping with DO-254, this clause
assumes that function allocations made during
system level considerations are to either software
or hardware. This part of the clause refers to
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REQUIREMENT COMPLIANCE GUIDANCE

those functions specifically allocated to hardware.
A hardware item is considered ‘complex’ if a
comprehensive combination of deterministic tests
and analyses cannot ensure correct functional
performance under all foreseeable operating
conditions with no anomalous behaviour. Meaning
that, if the item is so complex that it is impossible
or impractical to completely test and analyze it,
one must rely on design assurance to give
confidence in its correct operation.
Def Stan 00-55 Issue 3 provides guidance for the
development of requirements, design, verification
and validation of Safety Related Complex
Electronic Hardware (CEH). Additional
considerations relating to governance and
shortfalls against the Def Stan 00-55 Issue 3
requirements should be addressed along with any
‘military delta’ particular to the application.
Any contractor using Safety Related CEH that has
been previously developed and does not use DO-
254/ED-80 as its means of compliance is required
to justify the alternative means to the authority.
Justification for the use of the alternative means
of compliance should show that those means
meet the safety objectives of the regulations and
be supported by documented evidence, including
a full audit trail of the development history of the
Safety Related CEH.
Access to this documentation should be made
available to the authority to establish sufficient
confidence in the evidence. The System Safety
Assessment process and Safety Assessment
Report (or Air System Safety Case as
appropriate) will allow the authority to judge the
acceptability of previously developed CEH.

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SECTION 2

MILITARY COMMON FIT EQUIPMENT

SECTION 2 - LEAFLETS

0 References 2
1 View and Clear Vision - Standards of Rain 5
2 View and Clear Vision - Methods of Rain Clearance from Windscreens 6
3 Oxygen Systems - Physiological Requirements for Oxygen Systems 8
4 Oxygen Systems - Pressure Losses in Oxygen Delivery Systems 28
5 Oxygen Systems - Tests on Liquid Oxygen Systems 31
6 Ice Protection - Precautions to Prevent Waste Water Leaving Aeroplanes
as Ice 33
7 Ice Protection - Icing Conditions 34
8 Ice Protection - Ice Protection Systems 40

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SECTION 2

LEAFLET 0

REFERENCES

Each set of references is divided according to the reference number within Section 1.

1.5 ICE PROTECTION

MOD Specifications

Def Stan 00-970 (Part 1, Section 5, Clause 5.1.35 to 5.1.36) General specification for
aircraft gas turbine engines

DTD/RDI 3961 Windscreen de-icing pumps

RAE Reports

SME 3380 Protection of aircraft against ice

Mech Eng 2 Kinetic temperature of propeller blades in conditions of icing

Mech Eng 6 Design of heat exchangers

TR 77090 Calculation of surface temperatures and ice accretion rate in a mixed

water droplet/ice crystal cloud

Memo MAT/ST 1004 An investigation into the anti-icing of a heated cylinder in mixed
conditions

TR 82128 Ice accretion on aerofoils in 2-dimensional compressible flow - a

mathematical model

TR 84060 Calculation of water droplet trajectories about an aerofoil in steady 2-

dimensional compressible flow

TR 87013 Measurement of drag increase due to ice accretion on aerofoils of

NACA 0012 and RAE 9645 section

TR 88052 HOVACC - An aerofoil ice accretion prediction program for steady,

two-dimensional, compressible flow conditions

TR 90054 TRAJICE2 - A combined droplet trajectory and ice accretion

prediction program for aerofoils. 1990

RAE Technical Notes

Eng 124 Ice guards for aero-engine air intakes

Mech Eng 12 Icing tests on engine air induction systems

Mech Eng 19 Provision of heat on aircraft for protection against ice and for cabin
heating

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 DEF STAN 00-970 PART 13/11
SECTION 2

Mech Eng 57 The maintenance of clear vision through aircraft transparencies

Mech Eng 58 Tests of water spraying for simulating icing conditions ahead of a
turbine engine in flight

Mech Eng 62 The problem of icing as it affects modern military aircraft

Mech Eng 104 The maintenance of clear vision through fighter aircraft
transparencies

Mech Eng 208 Impingement of water droplets on aerofoils

Mech End 283 The analysis of measurements of free ice and ice-water
concentrations in the atmosphere in the equatorial zone (Tables 1 -
6)

NGTE Memorandum

M106 The anti-icing of compressor blades by surface heating - Part 1 -

Stator blades

ARC Reports

R & M 2805 Evaporation of drops of liquid (formerly RAE Report Mech Eng 1)

RAeS Journal

August, 1959 Water and ice in the atmosphere

AGARD

AGARD 16 Icing problems and recommended solutions

AGARD AR127 Aircraft icing

AGARD CPP236 Icing tests for aircraft engines

Defence Standard

01-5 Fuels lubricants and associated products

Note The following US reports contain useful bibliographies on the subject of aircraft icing
and ice protection:-

(a) Advisory Circular AC 20-73, Aircraft Ice Protection, (Appendices 1 and 2).
FAA, 1971

(b) NASA TM 81651, Selected Bibliography of NACA - NASA aircraft icing

publications

(c) Society of Automotive Engineers (SAE) Aerospace Information Report (AIR)

4015, Icing technology bibliography. (To be published)

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 DEF STAN 00-970 PART 13/11
SECTION 2

1.4 OXYGEN INSTALLATIONS

Standard Instruction Sheets

2611 Installation and testing of oxygen systems in aircraft
RAE Technical Notes
Mech Eng 292 Oxygen regulator system review
Defence Standards
16-1(Superseded by Def Stan 68-284) Breathing oxygen characteristics, supply pressures
and hoses for aircraft systems

81-24 Identification marking of cylinders, compressed gas

Flying Personnel Research Committee Reports
FPRC 1018(a) Oxygen requirements for high flying civil aircraft

FPRC 1142 Choice of gas mixtures for breathing in high performance

aircraft

FPRC 1163 Effects of breathing high concentration of oxygen upon the

diffusing capacity of the lung in man

FPRC 1165 The relation between the capacity of the regulator-mask hose
and the incidence of anoxia following rapid decompression
Institute of Aviation Medicine Reports
IAM 102 High altitude oxygen equipment
British Standards
C5 (Superseded by BS 2C 5) Mating dimensions for liquid oxygen replenishment couplings
for aircraft

C20 Aircraft gaseous oxygen replenishment connection (inch

dimensions)

N100 (Superseded by BS 4N 100) General requirements for aircraft oxygen systems and
equipment
Aerospace Information Report (SAE)

Air 822 Oxygen systems for General Aviation

Air 825A Oxygen equipment for Aircraft

US Military Specifications/Standards

Mil-STD 810(Superseded by 810F Chg Notice 3) Environmental Test Methods

4
 DEF STAN 00-970 PART 13/11
SECTION 2

LEAFLET 1

VIEW AND CLEAR VISION

STANDARDS OF RAIN
1 INTRODUCTION

1.1 This Leaflet defines the standards of rain to be used in designing and evaluating
windscreen rain removal systems.

2 RAIN STANDARDS SCALE

2.1 The following Table gives the droplet size, water content per cubic metre of air,
etc., of the rain referred to in Clauses 1.1.2.01 to 1.1.2.03, 1.5 and Part 1, Section 4 clause
4.17

3 NOTES

3.1 Very heavy rain as defined above is not uncommon during heavy showers and
thunderstorms in temperate latitudes, the horizontal extent being probably 1 to 3 miles.
Over short distances intensities of 100 mm per hour may be exceeded. Even greater
intensities may be experienced in the tropics.

3.2 Rain removal systems should be designed if possible to provide adequate vision at
the stated speed in very heavy rain. Satisfactory performance in heavy rain is mandatory.

3.3 These standards of rain have been reproduced in a rig at A. & A.E.E., and a good
correlation shown to exist between rig tests and the performance of the system in flight.

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 DEF STAN 00-970 PART 13/11
SECTION 2

LEAFLET 2

VIEW AND CLEAR VISION

METHODS OF RAIN CLEARANCE FROM WINDSCREENS

1 INTRODUCTION

1.1 This Leaflet, which is based on Ref 1, lists the current methods of rain clearance for
windscreens and the considerations which lead to the choice of the best system in a
particular case.

2 WINDSCREEN WIPERS

2.1 These are most useful on large screens and at low aeroplane speeds. They can
also be used in conjunction with an approved washing fluid to clean the screen. When used
in conjunction with an approved rain repellent they are effective even in cloud burst
conditions.

2.2 The positions of the wiper axes are important and, on a divided screen, if the pivots
are on the bottom inboard corners, the wipers give maximum clearance for minimum
power.

2.3 With development to suit the particular aeroplane, wipers can be effective at high
subsonic speeds.

2.4 Wipers should not be used on screens made of those plastic materials which can be
easily scratched.

3 HOT AIR JETS

3.1 A nozzle near the base of the screen discharges high velocity hot air over the
screen. Correct orientation of the nozzle(s) is essential in order to avoid local overheating of
the screen. The temperature needs to be over 100°C to evaporate the water but must not
be so hot as to crack the glass, damage an interlayer or edge sealant, or exceed the
temperature limits for plastic screens (especially those of stretched acrylic where surface
shrink-back can occur at temperatures as low as 105°C).

3.2 The advantages of the system are that:

(a) it can be used on glass or plastic,

(b) it is effective at high subsonic speeds,

(c) it can be used for de-icing and de-misting, and

(d) it has low drag when not in use.

3.3 The disadvantages are that:

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 DEF STAN 00-970 PART 13/11
SECTION 2

(a) use of the system at take-off can significantly reduce the power available for
flight,

(b) during landing the air supply with engines throttled may be inadequate, and

(c) it is sometimes difficult to find space in the right place for the hot air ducting
and mixers. Furthermore the provision for the supply of air at correct temperature
and pressure is also very difficult.

4 RAIN REPELLENTS

4.1 There are two varieties:

(a) A type (1) which needs to be rubbed on to a clean dry screen before flight at
intervals depending on weather conditions and the role of the aeroplane.

(b) A type (2) which must be sprayed on to a wet screen in rain as required in
flight.

4.2 The advantage of both types is that:

(a) they use no power,

(b) type (1) requires no space in the aeroplane, type (2) takes up only very little
space,

(c) salt spray and insect debris appear to adhere less to the treated surfaces
and the build up of ice deposits is retarded.

4.3 A disadvantage of both types is that with steeply raked screens a haze effect can be
produced in light rain when flying towards the light.

4.4 A disadvantage of the rub on type is that it may be difficult in operational conditions
to get the screen clean and dry before application.

5 SUMMARY

5.1 It may be necessary to use rain repellents in conjunction with windscreen wipers in
order to provide maximum vision in rain over an extended range of speed and rain intensity.
At low forward speed (e.g. on the ground) the distribution of rain repellent due to airflow can
be inadequate and reliance must be placed on wipers.

5.2 The surfaces of plastics materials are less readily wetted than those of glasses so
that less help from external sources is required to provide clear vision in rain through plastic
screens.

REFERENCES

No Author Title
1 Booker J D Aircraft windscreen rain clearance – A review.
RAE Technical Report No. 71022.

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 DEF STAN 00-970 PART 13/11
SECTION 2

LEAFLET 3

OXYGEN SYSTEMS

PHYSIOLOGICAL REQUIREMENTS FOR OXYGEN SYSTEMS

1 Introduction

1.1 This leaflet describes the basis of the physiological requirements for breathing systems
which are fitted to aircraft primarily to provide individual protection against hypoxia for aircrew and
passengers. Breathing equipment is also employed to provide protection against the inhalation of
toxic gases and fumes which may arise in the cabin environment and of airborne nuclear,
biological and chemical (NBC) warfare agents. Recently the delivery of positive pressure breathing
has been adopted as a means of enhancing aircrew tolerance of high sustained +Gz accelerations
(pressure breathing with G, PBG). Breathing equipment may, by the composition of the gases
delivered to the respiratory tract and/or the associated impedance to breathing induce undesirable
or indeed unacceptable physiological and/or performance disturbances in the individual.

1.2. The physiological requirements for the performance of aircraft breathing systems represent
practical compromises between the physiological ideal that the equipment should produce no
disturbance whatsoever to the user, the performance of available designs and the operational and
logistic requirements of simplicity, reliability, low maintenance and low financial cost. They are
based upon laboratory and airborne research and operational experience. As the requirements are
compromises they may vary with the application. Thus the degree of hypoxia which is acceptable
in seated passengers following decompression of the cabin of an aircraft at high altitude differs
markedly from that which is acceptable in the pilot of high performance combat aircraft during flight
with the cabin pressurised.

1.3 The design and operation of an aircraft breathing system must be very closely related to the
breathing requirements of the wearer. The three major aspects of these are the respiratory flow
demands, the pressure at the entrance to the respiratory tract (nose and mouth) and the
composition of the gas which is delivered to the respiratory tract. The physiologically acceptable
values of these breathing requirements are addressed in this leaflet together with the other
physiological factors which affect the performance of aircraft breathing systems.

2 Respiration in Flight

2.1 The ranges of instantaneous and average flow rates which can be demanded by fit adults
are extremely large. Thus the peak inspiratory flow rate can vary from 0.4 - 0.5 L (BTPS) s-1 at rest
to 10 L (BTPS) s-1 in maximum exercise, and the mean inspired pulmonary ventilation from 6 L
(BTPS) min-1 at rest to 150 L (BTPS) min-1 in maximal exercise. Whilst it is self-evident that aircrew
will not perform maximal exercise in flight a pilot who climbs into the cockpit of his aircraft after
running may well have a very high respiratory demand. Knowledge of the pulmonary ventilation
(average inspiratory or expiratory flow) and the instantaneous respiratory flow rates demanded by
aircrew (and passengers) in flight and on the ground is essential for the specification of the
performance required of an aircraft breathing system. Thus the pulmonary ventilation which occurs
under various conditions of flight will determine the size of the main, back-up and emergency
stores of oxygen required in an aircraft. Knowledge of the ranges of pulmonary ventilation which
may be demanded by aircrew over relatively short periods of flight [30 seconds or so] is required
for the specification of the performance of molecular sieve oxygen concentrator systems. Finally
the impedance to respiration imposed by any breathing system is a function of the instantaneous
inspiratory and expiratory flow rates created by the wearer.

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 DEF STAN 00-970 PART 13/11
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Effect of Altitude

2.2 The gases in the lungs are always saturated with water vapour at body temperature (37°C)
so that the partial pressure of water (PH2O) in the alveolar gas is always 47 mmHg. These
conditions of a temperature of 37°C and a PH2O of 47 mmHg are termed body temperature,
pressure and saturated with water vapour (BTPS). Thus as a dry gas at ambient pressure and
temperature (ATPD) enters the lungs, it expands, not only due to the rise in temperature, but also
due to the addition of water vapour. Whilst the increase in volume due to a change of temperature
from the standard temperature of NTPD conditions (15°C) to 37°C of 7.6% is independent of
changes in ambient pressure, the increase in the volume of the gas due to the PH2O rising from 0
to 47 mmHg varies with altitude from 6.6% at ground level, to 14.1% at 18,000 feet and 50% at
40,000 feet. Whilst the dependence of the relationship between gas volumes at ATPD and BTPS
conditions upon ambient pressure (altitude) is of great importance in calculating the size of oxygen
stores this effect is normally neglected when considering instantaneous respiratory flow rates over
the normal range of cabin altitudes. Indeed specifications of instantaneous flow rates are by
convention considered to be unaffected by altitude and are stated as flow rates of dry gas at 15°C
and at the absolute pressure within the respiratory tract (mask cavity) (ATPD). These conventions
are followed in the ASIC Standards and NATO STANAGS and in Section 1 Clause 1.4

2.3 The level of pulmonary ventilation of an individual in the absence of hypoxia and emotional
disturbances is very closely related to the rate of production of carbon dioxide by the body, which
in turn is very closely related to the physical activity of the individual. Generally, pulmonary
ventilation is adjusted in relation to the rate of production of carbon dioxide to maintain a constant
partial pressure of carbon dioxide (PCO2) in the alveolar gas and the arterial blood Thus at a
constant level of activity (rate of production of carbon dioxide) the pulmonary ventilation expressed
as volume under BTPS conditions is unaffected by ascent to altitude, provided that the
concentration of oxygen in the inspired gas is raised in order to prevent any hypoxia (see below).

Average Pulmonary Ventilation

2.4 The maximum average pulmonary ventilation is the essential component of any calculation
of the quantity of gas required to supply aircrew using a demand type of flow regulated breathing
system. The value used for the latter must take into account the effects of various stages of a
sortie and of types of flight on pulmonary ventilation, and the variation in the individual responses
to each condition. Extensive flight trials in combat aircraft the UK in the early 1960s led to the
adoption of the maximum [to include 97% of occurrences] pulmonary ventilations, averaged over
the whole sortie of the crews of combat aircraft presented in Table 1. Equivalent data for the front
and rear crews of large aircraft are not yet available. Increasing the number of crew will reduce the
mean pulmonary ventilation per individual as it is very unlikely that the pulmonary ventilations of
several crew members will be at the maximum value measured for a single crew member. Flight
trials conducted in fighter and bomber aircraft in the early 1960s formed the basis of the values
presented for multiple crews in Table 1. These values together with the air dilution characteristics
of the MK17, 20 and 21 series of pressure demand regulators form the basis of the oxygen
requirements for aircrew in Section 1.4. Measurements of pulmonary ventilation in mock air-to-air
combat in a Hunter T7 aircraft yielded a mean value for 18 pilots of 18.8 L (BTPS)min-1 with a
maximum average value (to include 97% of all observations) of 24L(BTPS)min-1. Measurements of
pulmonary ventilation in 12 rotary wing rear crew in a ground based simulation of flight tasks
yielded a mean value of 13.9 L (BTPS) min-1 at rest, and pulmonary ventilation measured in 12
rotary wing pilots using a flight simulator revealed a mean value of 12.9 L (BTPS) min-1 during
cruise.

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 DEF STAN 00-970 PART 13/11
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Pulmonary Ventilation of Aircrew operating combat aircraft averaged over a sortie [values to
include 97% of occurrences].
Number of Seats in aircraft Average Pulmonary Ventilation (L (BTPS) min-1)
1 21.8
2 or 3 18.6
4 or more 16.2

Table 1 Pulmonary ventilation of aircrew operating combat aircraft.

2.5 It should be emphasised that the values for maximum average pulmonary ventilation over
complete sorties presented in Table 1 should be used with caution in circumstances outside those
in which the information on which they are based was collected. Although estimates of the quantity
of oxygen required in aircraft storage systems in UK military aircraft have been based upon these
values for pulmonary ventilation since the early 1960s the adoption of standard sizes of LOX
converters in the mid-1960s led in general to the capacity of the oxygen store being greater than
that required by the values presented in Table 1. This greater margin of oxygen supply would tend
to mask in service any underestimate of the quantity of oxygen required based on the values in
Table 1.

Minimum and Maximum Pulmonary Ventilation

2.6 The pulmonary ventilation of an aircrew member may vary markedly with time during a
flight. Thus pulmonary ventilation is typically raised during the stress of take-off and landing. It is
raised by flight at low level and when performing the anti-G straining manoeuvre and by air
combat. The minimum and maximum levels of pulmonary ventilation which may occur in flight are
of importance to certain aspects of the design of breathing systems, such as the performance of
the injector form of air dilution mechanism employed in many demand regulators and the
concentration of oxygen in the product gas provided by a molecular sieve oxygen concentrator. In
this context the relevant value of the pulmonary ventilation is typically that averaged over a period
of 30 seconds or longer. The minimum pulmonary ventilation which will be demanded by a pilot
during undisturbed straight and level flight or by a seated passenger is very similar to the minimum
seen in subjects seated at rest on the ground i.e. 6.0 L (BTPS)min-1. Thus the value for the
minimum pulmonary ventilation to be met by aircraft breathing systems used by either aircrew or
passengers adopted in current national and international standards is 5.0 L (ATPD) min-1.

2.7 The few studies which have been made of maximum pulmonary ventilation in flight in pilots
performing simulated aerial combat and other manoeuvres have yielded values between 51 and 60
L(BTPS) min-1 for the maximum pulmonary ventilation maintained for 30 seconds or longer. Values
higher than 40 L (BTPS) min-1 were recorded on 1-2% of occasions. Although there are no data for
large aircraft crews available, pulmonary ventilation measured in 12 rotary wing rear crew in a
ground based simulation of flight tasks yielded a mean maximum value of 66.7 L (BTPS) min-1
during brief heavy exercise before landing. In 12 rotary wing pilots, pulmonary ventilation
measured in a flight simulator showed a mean value of 19.1 L (BTPS) min-1 during complex flight.
The standard adopted by the ASIC and NATO nations for the maximum pulmonary ventilation
which can be sustained in flight for 30s or more is 50 L (ATPD) min-1. It is unlikely that the flight
deck crew of transport aircraft would exhibit pulmonary ventilations as high as 50 L(ATPD) min-1,
whilst seated in flight. A realistic value for the maximum pulmonary ventilation to be sustained by
these aircrew for 30 seconds or longer is 40 L (ATPD)min-1. The maximum pulmonary ventilation
demanded by flight deck crew when moving out of their seats may well, however, exceed 50 L
(ATPD) min-1. A sustained pulmonary ventilation of 50 L (ATPD) min-1 would be demanded by an
aircrew member fighting an in flight fire. The pulmonary ventilations of seated passengers will
under normal flight conditions vary between 5 and 20 L (ATPD) min-1, depending upon their level of
activity. The emotional disturbances, such as fear, which may be engendered by decompression of
the cabin can raise the pulmonary ventilation of passengers above this range. European Aviation
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Safety Authority and US Federal Aviation Agency requirements for passenger breathing equipment
specify a maximum sustained pulmonary ventilation of 30 L (BTPS) min-1.

Instantaneous Respiratory Flow Rates

2.8 The instantaneous rates of flow of gas into and out of the respiratory system is one of the
principal factors which determine the magnitude of the changes in mask pressure imposed by a
breathing gas delivery system, the other being the pressure-flow characteristics of the breathing
equipment itself. There is considerable variation in respiratory flow patterns between individuals,
especially when breathing at rest. Typically, the breathing frequency is 15 breaths a minute with
each inspiration occupying 1.6 seconds and each expiration lasting 2.4 s. During inspiration the
instantaneous flow typically rises rapidly to a maximum after 0.5 seconds and then falls at a slower
rate to zero. Expiration usually follows the end of inspiration without a break. The instantaneous
flow in expiration rises to reach a maximum which is somewhat less than the maximum attained in
inspiration. The flow then declines slowly to reach zero at the end of this phase of the cycle. There
is frequently a pause between the end of expiration and the commencement of the next inspiration.
Of particular significance to the design of aircraft breathing systems are the maximum (peak)
inspiratory and expiratory flows which occur during the respiratory cycle. Although for some
purposes the instantaneous respiratory flow can be simulated by a sine wave (when the peak flow
equals 3.14 times the pulmonary ventilation) the peak inspiratory flow at rest is typically 3.2-3.8
times the pulmonary ventilation, and the peak expiratory flow is about 2.7 - 3.0 times the
pulmonary ventilation. Respiratory flow patterns tend to become more regular as the pulmonary
ventilation is increased by physical exercise. Inspiratory and expiratory times become more equal
as do peak inspiratory and expiratory flow. Moderate increases of pulmonary ventilation produced
by physical exercise typically occur by an increase in the size of individual breaths (i.e. increase of
tidal volume) rather than an increase in the frequency of breathing. When the respiratory frequency
increases it principally occurs by a shortening of the duration of expiration; the duration of
inspiration only decreases slightly with increasing frequency of breathing.

2.9 Respiratory flow patterns are modified by numerous factors ranging from exercise, speech
and swallowing to the imposition of external resistance to breathing (see paragraph 3.1) and
pressure breathing. Of particular relevance to the requirements for aircraft breathing systems are
the changes produced by speech and the anti-G straining manoeuvre (AGSM). The duration of
inspiration is markedly reduced by speech and since there are usually only minor changes in the
tidal volume the peak inspiratory flow is increased to 5 to 10 times the pulmonary ventilation. The
expiratory flow is modulated during speech and the maximum flow is less than that during
breathing at the same level without speech. The voluntary breathing manoeuvres involved in the
AGSM greatly reduce the duration of inspiration and expiration. The breathing cycle is typically
completed in 1.0-1.5 seconds and the peak inspiratory and expiratory flows are increased to
between 7 and 15 times the pulmonary ventilation. Pressure breathing without counter-pressure to
the chest also produces marked changes in respiratory flow patterns with an increase in peak
inspiratory flow and expiration becoming prolonged with a relatively constant expiratory flow. The
application of full counter-pressure to the chest tends to restore the breathing flow patterns to
those seen in the absence of pressure breathing. The flow patterns during pressure breathing with
+Gz acceleration with counter-pressure applied to the chest and abdomen are similar to those
which occur in light to moderate exercise provided that the individual does not perform any
respiratory straining manoeuvre.

2.10 The specifications of the maximum peak flows to be met by aircraft breathing systems are
based principally upon the breathing patterns of aircrew recorded in flight. These have shown that
the peak inspiratory and expiratory flow of pilots operating high performance combat aircraft can be
as high as 5L (ATPD) s-1. Analysis of the frequency distributions of peak flow recorded in several in
flight studies has shown that the occurrence of peak flows in excess of 3.3L (ATPD) s-1 is 2.5%.
ASIC and NATO specifications require that breathing systems must be capable to meeting peak
inspiratory and expiratory flows of at least 3.3 L (ATPD) s-1. Ideally aircraft breathing systems
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should be capable of meeting peak flows of up to 4.2L (ATPD) s-1. Since speech at moderate
levels of pulmonary ventilation (20L (ATPD) min-1) will produce peak inspiratory flows of 2.5-3.3 L
(ATPD) s-1 breathing systems designed for use in multi-crew aircraft, including transport aircraft,
must also be capable of meeting peak inspiratory flow rates of 3.3L (ATPD) s-1.

2.11 In pressure demand breathing systems, the impedance to respiration imposed by the
system is a function not only of the instantaneous respiratory flow, but also the rate of change of
flow. The rates of change of flow which occur during breathing are related to the nature of the
respiratory manoeuvre e.g. quiet breathing, speech, AGSM, pressure breathing and also to a
limited extent the peak respiratory flow. Speech at rest increases the median rate of onset and
offset of inspiratory flow during quiet breathing from 1.6 L s-2 to 18 L s-2. In practice, the highest
rates of change of flow occur in speech and whilst performing the AGSM. The minimum rates of
change of inspiratory and expiratory flow specified by ASIC and NATO requirements for aircrew
breathing systems are 10L (ATPD) s-2 at a peak flow of 1.5 L (ATPD) s-1 increasing to 20 L (ATPD)
s-2 at a peak flow of 3.3 L (ATPD) s-1. These rates of change of flow define the rate of change
between 0 flow and 90% of the relevant peak flow.

2.12 As already discussed (paragraph 2.4) it is unlikely that the breathing patterns of the
members of two crew or multi-crew aircraft will coincide exactly in time. Monte Carlo simulation of
the inspiratory demands of two crew members suggests that 95% of all instantaneous peak
demand flow can be met by a breathing system which will provide 70% of the flow demanded when
the two crew members are breathing exactly in phase. An in-flight study in which the inspiratory
flow patterns of the two crew of a two seat combat were recorded during level flight, high G
aerobatics and simulated combat manoeuvring, showed that the beginning of inspiration occurred
simultaneously in the two pilots in less than 1% of 5,000 breaths. The UK standard requires that a
breathing system for two crew members provides 85% of the peak inspiratory flow which could be
demanded by both crew members breathing exactly in phase i.e. 5.6L (ATPD) s-1.

3 Resistance to Respiration

Effects of external resistance

3.1 Excessive external resistance to breathing can give rise to breathing discomfort, fatigue of
the respiratory muscles and to changes in pulmonary ventilation which generally causes
hypoventilation but on occasions can cause hyperventilation.. Excessive resistance also impairs
speech and the ability to perform the ASGM. Finally, changes in the mean pressure in the lungs
induced by external resistances can disturb the cardiovascular system and the distribution of body
fluids. Many laboratory based studies of the effects of adding external impedances to breathing
have employed resistances which had a linear relationship between pressure drop and flow rate.
Such studies showed that subjective discomfort occurred when resistances with a pressure drop
greater than 0.5 kPa at a flow of 1.4 L (ATPD) s-1 were imposed in inspiration and expiration
together. Other studies in which the additional work of breathing produced by a variety of levels of
external resistance was measured suggests that the limit of breathing comfort is reached when the
external work exceeds [0.5 + 0.02 x (pulmonary ventilation)] Joule per litre of pulmonary
ventilation.

3.2 In practice the resistance to respiration imposed by an aircraft breathing system is defined
in terms of the relationships between the pressure in the cavity of the mask and the corresponding
respiratory demands. It is generally most appropriate to relate the minimum and maximum mask
pressures during the respiratory cycle to the corresponding peak inspiratory and expiratory flows
demanded by the wearer. It is normal practice to describe the resistance imposed by aircrew
breathing equipment in terms of the total change of pressure in the mask cavity [the pressure
swing] and the minimum and maximum mask pressures which are produced by equal inspiratory
and expiratory flows. The pressure in the mask cavity averaged over the whole of the respiratory

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cycle is also a valuable expression of the performance of a breathing gas system, as this quantity
determines in part the stresses imposed on the heart and circulation by the equipment.

Total change of mask pressure

3.3 The total change (swing) of pressure in the mask cavity during the respiratory cycle (i.e. the
difference between the minimum and maximum mask cavity pressures) should be as low as
possible. The greater the swing, the greater is the sensation of resistance to breathing and the
greater is the likelihood of incidents of hyperventilation, particularly in situations of high mental
workload. The current standard for the maximum permissible change of mask cavity pressure
during the respiratory cycle (with equal peak inspiratory and expiratory flow rates) is presented in
Table 2. This standard ensures breathing comfort at pulmonary ventilations between 5 and 50 L
(ATPD) min-1. Although internal airway resistance is reduced at altitude, the effect on the total work
of breathing is relatively small and it is present practice to require the resistance to breathing
imposed by an aircrew breathing system to be within the same maximum limits at all altitudes from
ground level to 38,000 feet, above which altitude pressure breathing is operative.

Peak Inspiratory and Expiratory Flow Rates Maximum Acceptable Change of Mask
(L (ATPD) s-1) Cavity Pressure during the Respiratory
Cycle (kPa)
0.5 0.5
1.5 0.85
2.5 1.75
3.3 3.0

Table 2 The maximum acceptable change of pressure in the mask cavity during the respiratory cycle at
altitudes between ground level and a pressure altitude of 38,000 feet.

Safety Pressure

3.4 The design of an oro-nasal mask and its suspension system should be such that a good
seal between the edge of the mask and the face is maintained under all conditions of flight. The
standard of this seal should be such that the inboard leakage of ambient air into the mask cavity
does not exceed 5% of the pulmonary ventilation when the mean pressure in the mask cavity is
between 0 and 1 kPa less than that of the environment. There are, however, situations in which
this level of sealing of the mask to the face may not be achieved. Indeed a serious disadvantage of
suction demand breathing systems in practice is that hypoxia can occur at altitude due to the
inspiration of air through a leak between the mask and the face. Safety pressure which is the
maintenance of the pressure in the mask cavity during inspiration at a value greater than that of the
environment is widely employed in aircrew breathing systems to prevent the flow of environmental
gas into the mask when there is a failure of the seal of the mask to the face. The ingress of air,
toxic fumes or NBC warfare agents through a leak between the mask and the face could have
serious consequences. In large aircraft, the requirement to provide effective denitrogenation by
pre-breathing 100% oxygen to limit the risk of decompression sickness is also highly dependent on
safety pressure. As long as the pressure in the mask cavity remains greater than that of the
environment, then a failure of the seal of the mask to the face will result in a flow of breathing gas
from the mask to the environment thus preventing the contamination of the breathing gas in the
mask by air or toxic materials in the air. Although it is desirable that safety pressure is maintained
in the mask cavity even at high inspiratory flow rates and in the presence of large leaks, the
pressure-flow characteristics of most breathing gas delivery systems, in which the mask pressure
falls with increasing flow, and the compensation of the expiratory valve, make this goal difficult, if
not impossible, to meet. In such systems, a high safety pressure will be associated with a high
resistance to expiration. A mean pressure in the mask cavity of +0.5 kPa will, however, minimise
the total work of breathing and increase breathing comfort.
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3.5 The fraction of the inspired gas which enters a mask through a typical mask leak in a
suction demand system is greatest at low inspiratory flows. The ratio of flow through the leak to
total inspiratory flow falls rapidly as the latter rises. The presence of safety pressure is therefore
most important during quiet breathing. It is thus possible to strike a compromise between the
maximum inspiratory flow rate at which safety pressure is required to be present and the rise in
mask expiratory pressure produced by the safety pressure. The UK standard requires safety
pressure to be present in aircrew breathing systems at inspiratory flows of up to at least 1.2L
(ATPD) s-1 and limits the maximum mask pressures during expiration to the values presented in
Table 3 (Section B - safety pressure present). The minimum mask pressures allowed when safety
pressure is present are also presented in Table 3. These limits to the peak mask pressures when
safety pressure is present ensure that the effects of the associated increase of mean lung pressure
of +0.25 to +0.5 kPa upon the circulation and distribution of body fluids are minimal and acceptable
for many hours.

Peak Inspiratory and Expiratory Acceptable Mask Cavity Pressures (kPa)

Flows (L (ATPD) s-1)
Minimum Maximum
A. Safety Pressure Absent
0.5 -0.38 +0.38
1.5 -0.55 +0.65
2.5 -1.12 +1.00
3.3 -1.90 +1.50

B. Safety Pressure Present

0.5 +0.02 +0.75
1.5 -0.20 +0.95
2.5 -0.90 +1.25
3.3 -1.75 +1.65

Table 3 The minimum and maximum acceptable mask cavity pressures during the respiratory cycle at
altitudes between Ground Level and a pressure altitude of 38,000 feet.

3.6 In some aircraft breathing systems safety pressure is only operative at altitudes above
either 10,000-15,000 feet or 30,000 feet. Below these altitudes, gas only flows from the regulator
when the pressure in the mask is reduced below that of the environment. The reduction of mask
pressure which occurs during inspiration in these circumstances should not give rise to the
sensation of excessive inspiratory resistance. The suction in the mask cavity is not to exceed the
values specified in Table 3 for the absence of safety pressure (section A). The maximum mask
pressures which occur when safety pressure is not operative should be such that there is no
sensation of excessive expiratory resistance. The maximum acceptable values are specified in
Table 3 (Section A - safety pressure absent).

Further Increases of Mask Pressure

3.7 In use, certain routine and emergency conditions tend to raise the pressure in the mask
cavity above the values seen during breathing in the steady state. Thus, in a typical pressure
demand system in which the outlet valve of the mask is compensated to the pressure in the inlet
hose of the mask, head movement increases the pressure in the mask hose and hence the
resistance to expiration and similarly a rise of mask hose pressure produced by a rapid ascent also
increases expiratory resistance. In order to maintain breathing comfort, the rise of mask cavity
pressure induced by realistic head movements or by the maximum rate of ascent of cabin altitude
(with the cabin pressurised) is not to exceed 0.25 kPa. A continuous flow failure of the demand
valve in a conventional compensated mask outlet valve system will result in a continuous rise of
mask pressure. If the flow through the demand valve is relatively low, the wearer will experience
expiratory difficulty. A high continuous flow will produce a rapid rise of mask and lung pressures,
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provided that the seal of the mask to the face is maintained. Inflation of the lungs to an
intrapulmonary pressure of 10.7-13.3 kPag will, if the expiratory muscles are relaxed, result in
overdistension of the lung tissue, rupture of the walls of air sacs and the passage of gas into the
lung tissue, into the tissues within the chest and neck, into the pleural space (producing lung
collapse) and most seriously into the ruptured pulmonary capillaries, allowing bubbles of gas into
the heart and arterial vessels with a high probability of blocking arteries supplying parts of the brain
which may cause unconsciousness and death. The rise of mask pressure produced by a high
continuous flow failure of a demand valve must not exceed 5.5 kPa.

Venting of lungs on rapid decompression

3.8 Rapid decompression of the pressure cabin of an aircraft produces an almost equally rapid
expansion of the gases in the lungs and airways and can produce over-inflation of the lungs with
damage to the lung tissue with the consequences discussed in paragraph 3.7. The incidence and
severity of the damage to the lungs produced by rapid decompression are determined primarily by
the ratio of cabin pressure before the decompression to that after the decompression, the speed of
the decompression (the reciprocal of the time constant of the decompression), the degree of
opening the glottis (the orifice between the vocal chords) and the resistance to the flow of gas from
the respiratory tract imposed by the breathing equipment. The breathing equipment worn by
aircrew should allow free venting of the expanding gases from the lungs in these circumstances.
The peak pressure difference between the gas in the lungs and the environmental pressure
produced by a rapid decompression should not exceed the 10.6 - 13.3 kPag required to produce
pulmonary damage by over-inflation of the relaxed chest. Present standards for aircrew breathing
equipment require that the mask pressure on a rapid decompression to a final altitude of 38,000
feet (above this altitude pressure breathing is operative) in 0.1 seconds shall not exceed 5.5 kPag.
This limit is somewhat arbitrary. It is one half of the intrapulmonary pressure required to damage
the lungs by over-distension of the relaxed chest. There is some experimental evidence that short
duration (<50 ms) peak mask pressures of up to 13.3 kPag on rapid decompression over a 35 kPa
pressure change in 0.2 seconds will not cause lung damage. The probability of lung damage on
rapid decompression is reduced if over-distension of the lungs is prevented by the application of
counter pressure to the chest wall and abdomen during the decompression.

Oscillatory activity

3.9 Aircrew breathing systems can exhibit oscillatory activity which produces oscillations of
pressure in the mask, usually during inspiration. Such oscillations of mask pressure, particularly if
they are of sufficient amplitude, are subjectively disturbing, may induce hyperventilation and can
interfere with communication. The incidence, amplitude and frequency of these oscillations are
determined by the oscillatory mechanics of the breathing equipment, by the impedance of the
respiratory tract [when present, oscillatory activity is frequently much greater when the wearer
breathes through the nose as compared with breathing through the mouth] and the respiratory flow
pattern. Ideally any oscillatory activity which occurs should not be detectable subjectively; it must
not be disturbing. Thus the double amplitude of any oscillation of pressure in the mask cavity which
persists for longer than 0.25 seconds should not exceed 0.06 kPa.

4 Composition of Inspired Gas

4.1 Several physiological factors influence the requirements for the composition of the gas
delivered to the respiratory tract. It is convenient to consider these requirements in terms of the
limits to the concentration of oxygen in relation to cabin altitude. In conventional oxygen systems
the diluting gas is virtually entirely nitrogen since the oxygen from the aircraft store is diluted with
cabin air. The performance of molecular sieve oxygen concentrators is such that the product gas
contains argon as well as oxygen and nitrogen. The maximum concentration of argon in the
product gas is 5-6%. In this context argon has no specific physiological effects and can be
regarded solely as an inert diluents gas. The concentration of oxygen in the gas delivered by a

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breathing system to the nose and mouth not infrequently fluctuates during a single breath
(especially in continuous flow oxygen systems) and from one breath to another (as occurs in some
molecular sieve oxygen concentrator systems). The physiological requirements with respect to the
concentration of oxygen in the inspired gas discussed in the following paragraphs assume that the
inspired gas is thoroughly mixed before it enters the respiratory tract. In mechanical testing the
mean volume weighed concentration of oxygen in the gas delivered to the mask cavity should be
determined by passing the gas from the expiratory port of the mask through a mixing box fitted with
baffle plates and measuring the concentration of oxygen in the mixed gas flowing from the box.
The final definitive measure of the “effective” concentration of oxygen delivered by a breathing
system is the measurement of the alveolar PO2 in human subjects breathing from the system
during man rating. The composition of the alveolar gas may, with certain precautions, be
determined in normal healthy subjects by measuring the PO2 and PCO2 of the gas flowing from the
nose and mouth towards the end of expiration [the end-tidal PO2 and PCO2].

4.2 The composition of the gas which enters the respiratory tract during inspiration when
wearing breathing equipment depends not only on the composition of the gas delivered to the oro-
nasal mask (or pressure helmet) through the inlet hose, but also on the proportion of the tidal
volume which is gas which had been breathed out in the previous expiration. The re-breathing of
previously expired gas adds external dead space to the respiratory tract and lowers the
concentration of oxygen delivered to the alveolar gas. It also impairs the elimination of carbon
dioxide from the body which raises the alveolar PCO2 which in turn increases the pulmonary
ventilation. The volume of the external dead space added by the oro-nasal mask (or pressure
helmet) must therefore be minimised. Depending upon the shape of the cavity of the mask and the
positioning of the inlet and outlet valves the effective respiratory dead space of a mask may be less
than the volume of the mask cavity when the mask is sealed to the face of the wearer. The
respiratory dead space added by most modern aircrew masks is of the order of 0.10-0.15L (ATPD).
The maximum acceptable effective respiratory dead space of an oro-nasal mask or pressure
helmet is 0.2L (ATPD).

Minimum Concentration of Oxygen in the Steady State

4.3 The principal consideration is that the concentration of oxygen in the inspired gas shall be
adequate to prevent significant hypoxia. The partial pressure of oxygen (PO2) in the alveolar gas
when breathing air at ground level (barometric pressure - 760 mmHg) is normally 103 +3 mmHg.
The ability of a subject to respond rapidly to a novel situation is marginally impaired when the
alveolar PO2 is reduced to 75 mmHg by breathing air at a pressure altitude of 5,000 feet and
significantly reduced when the alveolar PO2 is reduced to below 60 mmHg by breathing air at
pressure altitudes greater than 8,000 feet. When breathing equipment is worn throughout flight as
by the aircrew of high performance combat aircraft, the concentration of oxygen in the inspired gas
is to be such that the alveolar PO2 is maintained at or above the value produced by breathing air at
ground level, i.e. 103 mmHg. The alveolar PO2 should never be allowed to fall below 75 mmHg
[the alveolar PO2 produced by breathing air at a pressure altitude of 5,000 feet] during normal flight
with the cabin pressurised. The devices employed in molecular sieve oxygen concentrator systems
to provide warning when the PO2 of the product gas falls below an acceptable value have a
significant tolerance band within which they may or may not provide a warning of a low PO2. In
order to ensure that adequate warning of impending hypoxia is given without spurious warnings,
the minimum PO2 of the product gas when the system is operating correctly should not be less
than that required to maintain an alveolar PO2 of 103 mmHg. The warning system shall always
provide a warning when the PO2 of the product gas falls below that required to maintain an alveolar
PO2 of 75 mmHg.

4.4 The concentration of oxygen required at a given altitude to produce a given alveolar PO2 is
calculated using the Alveolar Gas Equation with assumptions with respect to the partial pressure of
carbon dioxide (PCO2) in the alveolar gas and the respiratory exchange ratio, R. The Alveolar Gas
Equation states that:
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PA O2 = (PI O2 - PA CO2 [FI O2 + (1-FI O2)])/R

Where
PA O2 = P O2 in alveolar gas
PI O2 = P O2 in inspired gas saturated with water vapour at 37°C
PA CO2 = P CO2 in alveolar gas
FI O2 = Fractional concentration of oxygen in the mixed dry inspired gas
R = Respiratory Exchange Ratio

The normal resting value of the alveolar PCO2 is 40 mmHg and of R is 0.85. The concentration of
oxygen required in the inspired gas and altitude to produce an alveolar PO2 of 103 mmHg is
presented in Figure 1. The concentration of oxygen required in the inspired gas to maintain an
alveolar PO2 of 103 mmHg rises to 100% at an altitude of 33,700 feet (barometer pressure = 190
mmHg). Above this altitude the alveolar PO2 will fall below 103 mmHg even when 100% oxygen is
breathed.

4.5 Breathing systems for aircrew whether in aircraft in which an oxygen mask is worn
throughout flight or in an aircraft in which the aircrew don oxygen masks only when the cabin
altitude exceeds 8,000 feet should provide the minimum concentration of oxygen in the inspired
gas in relation to cabin altitude which will maintain an alveolar PO2 of 103 mmHg [at cabin altitudes
up to 33,700 feet]. Some degree of hypoxia is, however, acceptable in passengers in the
emergency of loss of cabin pressure at altitude. European Aviation Safety Authority and US
Federal Aviation Agency specifications for passenger oxygen systems allow the mean
concentration of oxygen in the inspired gas to fall to a level which will produce an alveolar PO2 of
around 55 mmHg at pressure altitudes between 10,000 and 18,500 feet and an alveolar PO2 of
around 45 mmHg at pressure altitudes above 18,500 feet.

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Figure 1 The relationships between the concentration of oxygen in the inspired gas and cabin altitude
required (i) to maintain an alveolar PO2 of 103 mmHg - GL equivalent; (ii) to produce an alveolar PO2 of 30
mmHg on rapid decompression to various final altitudes and intrapulmonary pressures - broken lines and (iii)
to ensure rapid decompression of a 35 kPag pressure cabin will produce a minimum alveolar PO2 of 30
mmHg when using two common pressure breathing schedules at pressure altitudes above 40,000 feet - solid
lines.

Minimum Concentration of Oxygen to prevent Hypoxia on Rapid Decompression

4.6 A second factor which influences the relationship between the concentration of oxygen in
the inspired gas and cabin altitude is the need to prevent impairment of performance due to
hypoxia following a failure of the pressure cabin at high altitude. When the inspired gas breathed
before the decompression contains a significant concentration of nitrogen, the fall of the total
pressure of the alveolar gas produced by rapid decompression produces a concomitant reduction
of the alveolar PO2 which may be to such a level that it produces impairment of performance or
even unconsciousness. If the decompression is to a pressure altitude greater than 30,000 feet then
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100% oxygen must be delivered to the respiratory tract immediately the decompression occurs if
there is not to be a significant impairment of consciousness. There will be a significant impairment
of performance if the alveolar PO2 is reduced during the decompression to below 30 mmHg even
for only a few seconds. If the magnitude of the area enclosed between an alveolar PO2 of 30
mmHg above and the time course of alveolar PO2 below exceeds 140 mm Hg.s then the individual
will become unconscious. The decrement of performance at a choice reaction task is proportional
to the magnitude of the area bordered above by a PO2 of 30 mmHg and the time course of the
alveolar PO2 below. The breathing gas delivery system shall therefore prevent the alveolar PO2
falling below 30 mmHg during and subsequent to a rapid decompression.

4.7 The major factors determining the minimum value of the alveolar PO2 immediately after a
rapid decompression are the initial and final absolute pressures of the alveolar gases, and the
composition of the gases breathed before and after the decompression. Assuming that 100%
oxygen is delivered to the respiratory tract immediately the decompression occurs, the alveolar
PO2 can be prevented from falling below 30 mmHg by ensuring that the gas breathed before the
decompression contains an adequate concentration of oxygen and that the total intrapulmonary
pressure does not fall below 115-120 mmHg (15.3-16 kPa) absolute. Assuming that the duration of
the decompression is so short that there is no significant exchange of oxygen between the alveolar
gas and the blood flowing through the lungs, then the alveolar PO2 immediately after a
decompression in which the absolute pressure of the lungs falls from PL(i) to PL(f) is related to the
alveolar PO2 immediately before the decompression by the equation:

Final alveolar PO2 = (Initial alveolar PO2 x (PL(f)-47)) / (PL(i)-47)

[all pressures expressed as mmHg]

4.8 This simple relationship may be employed to calculate the value of the alveolar PO2 before
the decompression which will produce an alveolar PO2 of 30 mmHg immediately after the
decompression from the initial to the final absolute pressures of the lung gas. The Alveolar Gas
Equation (paragraph 4.4) can then be used to calculate the concentration of oxygen required in the
inspired gas to ensure that the specified decompression will produce an alveolar PO2 of 30 mmHg
(but no lower) immediately after decompression. The concentrations of oxygen required in the
inspired gas to produce an alveolar PO2 of 30 mmHg immediately after a rapid decompression
from a given initial cabin altitude to a given final cabin altitude [total absolute alveolar gas pressure
at final cabin altitudes above 40,000 feet] are indicated by the interrupted curves of Figure 1. The
relationship between initial cabin altitude and the final cabin altitude is determined by the
pressurisation schedule of the cabin of the aircraft. The final alveolar gas pressure is also
determined by the safety pressure/pressure breathing characteristics of the breathing gas delivery
system. Thus the curve relating the minimum concentration of oxygen in the inspired gas to cabin
altitude before a decompression required to prevent the alveolar PO2 falling below 30 mmHg
immediately after the decompression will depend upon the cabin pressurisation schedule of the
aircraft and the safety pressure/pressure breathing characteristics of the breathing gas delivery
system. In large aircraft a watch-keeping pilot is usually required to don and wear an oxygen mask
on the flight deck when aircraft altitude is 40,000 feet or greater, to ensure appropriate oxygenation
before decompression and timely provision of 100% oxygen.

4.9 The minimum inspired oxygen concentration-cabin altitude curves for two commonly used
pressure breathing systems employed in aircraft with a cabin pressure differential of 35 kPa at
aircraft altitudes above 23,000 feet are presented in Figure 1. Both of these pressure breathing
systems commence pressure breathing at a cabin altitude of 40,000 feet and deliver oxygen at an
absolute pressure which falls linearly with the reduction of environmental pressure at altitudes
above 40,000 feet. One system employs a breathing pressure of 30 mmHg at 50,000 feet which
provides an intrapulmonary pressure of 117.5 mmHg absolute at 50,000 feet. The other system
employs a breathing pressure of 70 mmHg at 60,000 feet which provides an intrapulmonary
pressure of 124 mmHg absolute at 60,000 feet. It may be seen from Figure 1 that the minimum
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concentration of oxygen required in the inspired gas to prevent significant hypoxia being induced
by the rapid decompression is greater than that required to maintain an alveolar PO2 of 103 mmHg
in the steady state at cabin altitudes above 16,000 feet. The concentration of oxygen required in
the inspired gas at cabin altitudes above 16,000 feet is greater with the pressure breathing system
which employs a breathing pressure of 30 mmHg at 50,000 feet than the system which employs a
breathing pressure of 70 mmHg at 60,000 feet. The minimum concentration of oxygen required in
relation to cabin altitude to prevent hypoxia in the steady state and in the event of a rapid
decompression in an aircraft with a 35 kPa differential pressure cabin and using a breathing
pressure of 30 mmHg at 50,000 feet is summarised in Figure 2.

Maximum Concentration of Oxygen

4.10 Breathing high concentrations of oxygen during flight in high performance, combat aircraft
has two important disadvantages. It results in acceleration atelectasis and delayed otitic
barotrauma.

4.11 Exposure to sustained positive acceleration whilst breathing high concentrations of oxygen
produces marked collapse of the lower part of the lungs due to the absorption of alveolar gas whilst
the small sized airways are collapsed by the increased weight of the lungs. The symptoms of the
condition are attacks of coughing accompanied often by a sense of difficulty of breathing or, less
frequently, by discomfort in the chest. The coughing is usually provoked by an attempt to take a
deep breath either in flight or, more frequently, on standing up in the cockpit after flight. The cough
and difficulty in breathing may last a few moments or repeated attacks may occur over a period of
10 to 15 min. Field studies have shown that 80-85% of pilots develop the condition with symptoms
in flights in which 100% oxygen is breathed and manoeuvres above +3 to 4Gz are performed. The
lung collapse which often reduces the vital capacity by 40% is associated with a large right to left
shunt (20-25% of the cardiac output) of venous blood flowing through the collapsed lung, which
reduces the concentration of oxygen in the arterial blood. The collapse remains after the return to
+1 Gz until the individual takes a deep breath and/or coughs.

4.12 The causative factors of acceleration atelectasis are exposure to accelerations greater than
+3 to 4 Gz and breathing 100% oxygen. The degree of lung collapse and the intensity of the
symptoms are greatly increased by inflation of the G trousers. The mechanism is absorption of gas
from non-ventilated alveoli in the lower parts of the lungs. The ventilation of these alveoli ceases
on exposure to +Gz acceleration as the increased weight of the lung above compresses the lower
parts of the lung, closing the small and intermediate sized airways. Inflation of the abdominal
bladder of the G trousers accentuates this process. A high concentration of nitrogen in the non-
ventilated alveoli will maintain the patency of the latter whilst the increased accelerative force is
operative and ventilation of the alveoli will recommence on return to +1 Gz. If, however, the gas
breathed before the exposure to +Gz acceleration is 100% oxygen so that the concentration of
nitrogen in the alveoli is very low, the blood flowing through the non-ventilated alveoli rapidly
absorbs all the gas trapped in the alveoli and surface forces maintain the alveoli in the collapsed
state after the return to +1 Gz until they are reopened by a deep inspiration and coughing. The rate
of absorption of gas from nonventilated alveoli is increased sixty times when 100% oxygen is
breathed instead of air before the cessation of ventilation of the lungs. The presence of a
significant concentration of nitrogen which has a much lower solubility in blood than oxygen and
carbon dioxide acts as a brake on the absorption of gas from the non-ventilated alveoli.

4.13 Although no long term deleterious effects have been found in aircrew who have had the
condition repeatedly in flight, many air forces consider that the chest discomfort which is produced
and the potential hazard to safety of coughing in flight make acceleration atelectasis unacceptable.
Acceleration atelectasis is less likely to occur if the concentration of nitrogen in the gas breathed
before and during the exposure to the sustained acceleration does not fall below 40%. In this
context, the argon which is present in breathing gas produced by molecular sieve oxygen
concentrators behaves as nitrogen as it is also relatively insoluble in blood. Laboratory studies
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suggest that the concentration of nitrogen required to prevent significant acceleration atelectasis at
altitudes up to 25,000 feet is also 40%. Flight experience at cabin altitudes up to 20,000 feet
confirms this finding.

4.14 Breathing 100% oxygen, especially if it is associated with ascent to and descent from even
moderate altitudes, is followed in the vast majority of individuals by the development of ear
discomfort and deafness (delayed otitic barotrauma). A typical picture is that, on waking from a
night’s sleep, following flights in which 100% oxygen has been breathed, the individual has
discomfort in the ears and is moderately deaf. Breathing 100% oxygen results in the nitrogen
normally present in the middle ear cavity being washed out and replaced by oxygen through the
pharyngo-tympanic tube. In the absence of nitrogen or the presence of a high concentration of
oxygen in the middle ear cavity the blood flowing through the wall of the cavity rapidly absorbs gas
from the cavity. The absorption of gas reduces the pressure in the middle ear which draws the
eardrum into the cavity causing discomfort and deafness. The reduction in pressure also draws
fluid into the cavity. The process of absorption of gas from the middle ear can be slowed and
arrested after flight by “clearing the ears” whilst breathing air. The re-introduction of nitrogen into
the middle ear must be repeated several times over the 12-18 hours following a flight in which
100% oxygen is breathed if delayed otitic barotrauma is to be avoided. The use of 100% oxygen to
reduce risk of decompression sickness during deliberate depressurised operations means that for
certain mission profiles the occurrence of otitic barotrauma will be inevitable.

4.15 The incidence of delayed otitic barotrauma is reduced by the presence of a minimum
concentration of nitrogen in the gas breathed during flight. The concentration of nitrogen required
in the inspired gas to reduce the incidence and severity of this condition to negligible levels is
between 40% and 50%. Laboratory evidence suggests that the incidence of delayed otitic
barotrauma will be very low when the nitrogen concentration is between 30% and 40%.

4.16 The requirements to avoid acceleration atelectasis and delayed otitic barotrauma in flight
set the limit to the maximum concentration of oxygen which should be present in the gas delivered
to the respiratory tract by the breathing system of a high performance combat aircraft. This
requirement can be met by ensuring that the maximum oxygen concentration does not exceed
60%. There are obvious limits to the maximum altitude up to which this requirement can be
applied. Three factors play a part in deciding the range of cabin altitudes over which it should be
applied. The first factor is cabin pressurisation schedule. Aircrew operating combat aircraft will only
be exposed to cabin altitudes greater than 20,000-22,000 feet in the rare event of decompression
of the cabin at high altitude when 100% oxygen must be breathed in order to prevent hypoxia. The
second factor is the effect of high altitude upon the ability of the aircraft to sustain significant levels
of acceleration. Some high performance combat aircraft are able to sustain high G at aircraft
altitudes where the cabin altitude may be up to 22,000 feet. The third factor which is relevant is that
the design of the breathing system becomes more technically difficult and costs rise if the
difference between the minimum and maximum allowable oxygen concentrations is very small.
Such would be the case if the specification of performance required that the concentration of
oxygen should not exceed 60% at cabin altitudes much above 15,000 feet. Taking all these factors
into consideration, the compromise requirement in ASIC and NATO agreements is that the
concentration of oxygen in the inspired gas delivered by the breathing system of an aircraft in
which the aircrew will be exposed to sustained +Gz accelerations above +3 Gz should not exceed
60% at cabin altitudes up to 15,000 feet and 75% at a cabin altitude of 20,000 feet (Figure 2).

5 Pressure Breathing at Altitude

5.1 The principal physiological hazards associated with loss of cabin pressure at pressure
altitudes above 40,000 feet are hypoxia, decompression sickness and cold injury. A full pressure
suit assembly is necessary if protection against all three hazards is required over a prolonged
period. However, if the aircraft can descent promptly and rapidly (within 3-4 minutes) to an altitude
of less than 40,000 feet, protection against hypoxia only is required. A full pressure suit assembly
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will provide the ideal physiological protection, but it is bulky, cumbersome, impairs operational
efficiency during routine flying with an intact cabin, and imposes major ground procedural
problems. Pressure breathing combined with partial pressure garments at altitudes in excess of
50,000 feet is used therefore to provide short term or “get-me-down” protection against hypoxia.
Partial pressure garments are required to combat the undesirable physiological disturbances
produced by pressure breathing, but in order to exploit the advantages of the partial pressure
approach (less restriction when un-inflated and inflated, greater routine comfort and lower thermal
load), it is desirable that the area of the surface of the body to which counter pressure is applied
should be the minimum required to provide the specified protection. Thus the design of the counter
pressure garments represents a compromise between ideal physiological requirements and
functional convenience. In addition, since the protection against hypoxia using a partial pressure
assembly is required for only a short period of time during emergency descent, some compromise
in the level of alveolar PO2 which is required is also acceptable. It is the interaction of the
deleterious effects of hypoxia upon mental performance and the cardiovascular system, with the
undesirable consequences of positive pressure breathing, which determine the acceptable
minimum alveolar PO2. Virtually all pressure breathing systems and partial pressure assemblies
employ 100% oxygen in order to minimise the magnitude of the breathing pressure required at
altitudes above 40,000 feet to maintain the required alveolar PO2. The use of product gas from a
molecular sieve oxygen concentrator comprising 5-6% argon and 94- 95% oxygen during pressure
breathing at an altitude of 50,000 feet, requires the breathing pressure to be increased in order to
maintain the alveolar PO2 at the appropriate level. Pressure breathing with a pressure sealing
mask and no counter pressure to the body is widely used to provide short duration protection
against hypoxia on exposure to altitudes up to 48,000-50,000 feet. The mean mask cavity pressure
required at 50,000 feet is a compromise between too high a pressure which will produce a faint,
and too low a pressure which will not prevent a serious deterioration of performance due to
hypoxia. The acceptable compromise is a mean mask pressure between 4.0 and 4.5 kPag (30.0-
33.8 mmHg) at 50,000 feet. Between 38,000 feet and 50,000 feet the mean mask pressure should
increase linearly with fall of environmental pressure, the limits of mean mask pressure at 40,000
feet being +0.1 to 1.0 kPag (0.75-7.5 mmHg). During pressure breathing with a mask alone the
total change of mask cavity pressure during the respiratory cycle should not exceed 0.5 kPa at
peak inspiratory and expiratory flows of 0.5L (ATPD) s-1 and 1.0 kPa at peak inspiratory and
expiratory flows of 1.83L (ATPD) s-1 [the absolute pressure in this context is the absolute pressure
in the mask and respiratory tract].

5.2 The magnitude of the breathing pressure required to prevent unacceptable hypoxia at
pressure altitudes above 50,000 feet requires the application of counter pressure to the chest and
abdomen to support breathing and at higher altitudes counter pressure to at least a portion of the
limbs to counteract the effects of the raised intrapulmonary pressure upon the cardiovascular
system, and maintain an adequate arterial blood pressure and blood flow to the brain. Thus all
partial pressure assemblies apply counter pressure to the external surface of the chest, most
commonly by means of a bladder covering part or the entire chest and restrained within an outer
inextensible fabric layer. The bladder is connected into the hose between the breathing gas
demand regulator and the oro-nasal mask/pressure helmet so that it is inflated with breathing gas
to the breathing pressure provided by the regulator. The bladder of the pressure jerkin not only
applies counter pressure to the chest, but also to the whole of the abdomen which ensures the
minimum of respiratory disturbances during pressure breathing. In some partial pressure
assemblies counter pressure is applied to the abdomen and lower limbs by means of the G
trousers which the crew member is primarily wearing to enhance tolerance of +Gz acceleration.
The pressure in the G trousers during pressure breathing at altitude is raised to 1.5 to 3.2 times the
breathing pressure. The optimum ratio of G trouser to breathing pressures varies with the degree
of coverage provided by the G trousers and is about 2.0 when using the UK full coverage anti-G
trousers.

5.3 A well-sealing oro-nasal mask can be used to deliver breathing pressures of up to 9.3 kPag
(70 mmHg) for several minutes. A limited proportion of subjects can even tolerate pressure
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breathing with a mask at pressures up to 10.7 kPag (80 mmHg). The practical limit to the use of an
oro-nasal mask without external support to the upper neck is a breathing pressure of 9.3 - 10.0
kPag (70-75 mmHg). The standard of sealing of a mask employed to deliver high pressures to the
respiratory tract should be such that the outboard leakage from the mask when sealed to the face
does not exceed 0.10 L (ATPD) s-1 at a mask pressure of 4 kPa and 0.25 L (ATPD)s-1 at a mask
pressure of 9.3 kPa. If any leakage does occur during pressure breathing the fit of the mask should
be adjusted so that the leaks do not occur into the eyes.

5.4 Partial pressure assemblies which employ a partial pressure helmet to deliver 100%
oxygen to the respiratory tract maintain the absolute pressure in the lungs at 18.7-20.0 kPa (141-
150 mmHg) at all altitudes above 40,000 feet which, in the absence of hyperventilation, gives an
alveolar PO2 of 50-60 mmHg. The use of a breathing pressure of only 30 mmHg at 50,000 feet
results in an intrapulmonary pressure of 117 mmHg (15.6 kPa) absolute and an alveolar PO2 of 40
mmHg with a moderate degree of hyperventilation (alveolar PCO2 = 30 mmHg). When pressure
breathing is performed with an oro-nasal mask and counter pressure to the trunk and lower limbs
at breathing pressures up to 70 mmHg (9.3 kPag) an intrapulmonary pressure of 130 mmHg (17.3
kPa) absolute produces mild to moderate impairment. Several current partial pressure assemblies
comprising an oro-nasal mask with counter pressure to the trunk and lower limbs employ a
breathing pressure of 70 mmHg (9.3 kPag) at an altitude of 60,000 feet, which provides an
intrapulmonary pressure of 124 mmHg (16.5 kPa) absolute and an alveolar PO2 of 45-50 mmHg.
The relationship of breathing pressure (mask pressure) to altitude between 40,000 and 60,000 feet
can take several forms (Figure 3). The mask pressure can be held at 141 mmHg (18.8 kPa)
absolute with ascent above 40,000 feet until the breathing pressure reaches the maximum of 70
mmHg (9.3 kPag) [Figure 3 - solid line]. This relationship minimises the hypoxia at the intermediate
altitudes. An alternative relationship is one in which the absolute pressure in the mask falls linearly
with environmental pressure from 40,000 to 60,000 feet [Figure 3 - broken line]. This form of the
relationship minimises the cardiovascular stress at the intermediate altitudes. The mask pressure
averaged over the respiratory cycle during pressure breathing at altitudes over 40,000 feet is to be
within 0.27 kPa of the nominal mask pressure. Partial pressure breathing systems employing an
oro-nasal mask, trunk counter pressure (pressure waistcoat and G trousers or pressure jerkin) and
G trousers can provide acceptable to aircrew at altitudes up to 60,000 feet. There is at present an
insufficient body of evidence to support their use to provide protection at altitudes above 60,000
feet.

5.5 The resistance to breathing during pressure breathing with respiratory counter pressure is
determined by the relationships of the pressures in the mask cavity, the pressure applied to the
chest by the respiratory counter pressure garment [which is generally assumed to be the pressure
in the bladder, if a bladder system is used] and the pressure applied to the abdomen by the G
trousers. The swings of pressure in the mask cavity and the chest counter pressure garment during
pressure breathing with counter pressure should not exceed the limits specified in Table 2. The
difference between the pressure in the mask cavity and the chest counter pressure garment shall
at no time exceed 0.5 kPa.

5.6 The required intrapulmonary pressure must be established rapidly on a sudden

decompression to high altitude if hypoxia is to be avoided. Thus on a rapid decompression (in 0.1
s) to an altitude above 45,000 feet the pressures in the mask cavity and in the respiratory counter
pressure garment shall not fall below 120 mmHg (16 kPa absolute) for longer than 2 s. This
standard determines the requirement for the rate of inflation of the respiratory counter pressure
garment. In practice, however, where the garment will usually be inflated to safety pressure prior to
a decompression, there is a need to vent excess gas from the garment to avoid over-
pressurisation, although the latter can provide some protection against lung damage on a very
rapid decompression.

6 Pressure Breathing for +Gz Protection

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6.1 Pressure breathing for G protection (PBG) used in combination with anti-G trouser inflation
is now a well established technique for raising the tolerance of +Gz acceleration. Pressure
breathing together with extended cover G trousers can reduce the need for sustained G straining
manoeuvres during exposures to +8 to +9Gz. Use of a chest counter pressure garment has not
been demonstrated to provide any meaningful improvement in lung protection over the counter-
pressure inherent in the acceleration forces acting on the chest wall, and is not a mandatory
requirement for PBG (unlike pressure breathing for altitude). The pressure demand regulator
provides pressure breathing in response to the rise in the pressure at the outlet of the anti-G valve.
The latter typically controls the flow of cooled engine bleed air into and out of the G trousers. The
anti-G valve inflates the G trousers rapidly (within 1-2 s) to the desired pressure in relation to the
total applied +Gz. The relationship between pressure in the G trousers and applied G is trouser
pressure rising linearly with acceleration from 0 at +2 Gz to 70 kPag at +9 Gz.

6.2 The optimum breathing pressure at +9 Gz is 60-65 mmHg (8.0-8.7 kPag). The preferred
relationship is to commence pressure breathing at +4 Gz and for the breathing pressure to rise
linearly to 60-65 mmHg (8.0-8.7 kPag) at +9 Gz.

6.3 The resistance to breathing during pressure breathing with G should be minimal. The total
swing of mask pressure should not exceed the limits specified in Table 2.. Pressure breathing must
not be operative unless the G trousers are pressurised as pressure breathing on exposure to +Gz
acceleration without pressurisation of the G trousers will cause rapid loss of consciousness. The
rise of pressure in the mask on the sudden application of +Gz must not lag more than 0.5s behind
the rise of pressure in the G trousers, and the fall of pressure in the mask should not lag more than
0.5s behind the fall of pressure in the G trousers.

7 Pressure Breathing - Press-to-Test

7.1 A facility whereby pressure breathing may be obtained by the operation of a manual control
is required to enable the user to test the standard of seal of the breathing gas delivery system up to
and including the mask. The performance of this facility is to be such that the user can perform
several respiratory cycles with the mask pressure raised.

7.2 The test pressure to be employed varies with the pressure breathing assembly in use. The
mean mask pressure produced on press-to-test when a mask is worn alone should be within the
limits +3.5 to +4.5 kPag. When chest counter pressure and G trousers are worn for protection on
decompression at high altitude the facility should provide a mask pressure of 6.7 to 8 kPag and
inflation of the G trousers to 1-2 times breathing pressure. This mask pressure is also to be
provided in a press-to-test facility for a pressure breathing with G assembly. It should be noted that
it is not acceptable to provide this facility simply by inflating the G trousers to the appropriate
pressure, 62- 72 kPag, as the application of these G trouser pressures at +1 Gz gives rise to pain.
The total change of mask cavity pressure during the operation of the press-to-test facility should
not exceed 0.75 kPa at peak respiratory flows of 0.5L (ATPD) s-1 and 1.0 kPa at peak flows of 1.0L
(ATPD) s- 1.

8 Protection against Hypoxia after Ejection

8.1 The delivery of breathing gas to the respiratory tract following ejection from an aircraft at
altitude shall be such that significant hypoxia does not occur during the subsequent descent of the
crew member to below 10,000 feet. Typical descent times from various altitudes to 10,000 feet are
presented in Table 4.

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Starting Altitude (feet) Time to descent to Time to descent to 10,000 feet (second)
10,000 feet (second) Aircrew in ejection seat*
Aircrew Alone

20,000 40-60 70

30,000 70-110 130

40,000 95-160 170

50,000 110-190 215

60,000 130-220 245

Table 4 - Time to descend to 10,000 feet following ejection. * Ejection seat with 1.62 m diameter drogue.

8.2 The time taken to descend from altitudes up to 20,000-25,000 feet is such that breathing air
throughout the whole of the descent will not cause significant impairment of performance. Thus it is
not essential to provide supplemental oxygen for escape at altitudes up to 25,000 feet. Breathing
gas with a PO2 greater than 130-150 mmHg is required to prevent hypoxia on escape at altitudes
above 25,000 feet. Pressure breathing is required at altitudes above 40,000 feet. Inward relief
whereby the ejectee/parachutist can breathe ambient air in the event of either cessation of the
breathing gas supply or separation from the ejection seat, is required. The headgear including the
mask and its supply system must remain intact and remain in place during ejection and perform
satisfactorily thereafter. The breathing equipment must perform satisfactorily at low temperature (-
60°C) in the presence of representative air movement [at least 20 knots (37 km.h-1)].

9 Provision of Inward Relief

9.1 The ability to breathe air is required in the event that the flow of breathing gas provided by
the breathing equipment is inadequate to meet the inspiratory demand. This facility is necessary in
order to avoid a sudden failure of the supply of breathing gas imposing a very high resistance to
inspiration, a situation which could threaten flight safety. The inward relief facility must not allow the
ambient air to dilute the breathing gas delivered by the breathing system during normal operation
of the equipment and thereby cause hypoxia or allow toxic material in the cabin air to enter the
breathing system. The crew member should be aware immediately that air is entering the breathing
system. In many conventional breathing systems inward relief is obtained either by loosening the
mask so that air can be inspired around it or by disconnecting the inlet hose of the mask from the
supply system. Neither of these methods is satisfactory. Some systems employ a spring-loaded
inward relief valve (anti-suffocation valve) in the wall of the mask or mask hose connector. The
minimum suction required to open such an inward relief valve should be 1.25 - 1.75 kPag in order
to ensure that the opening of the valve is noticed immediately by the wearer but that the inspiratory
resistance is acceptable for breathing up to at least 30 minutes and it will not depress the
respiration of an unconscious crewmember.

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Figure 2 A specification embodying the physiological requirements for the relationship of the
concentration of oxygen in the inspired gas and cabin altitude in the intact pressure cabin of a typical agile
combat aircraft with a 35 kPag pressure cabin and a ceiling of 50,000 feet.

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Figure 3 Two acceptable forms of the relationship between mean mask pressure and altitude
for a partial pressure assembly comprising a mask, pressure waistcoat or jerkin and G trousers.
Note that both of these lines are nominal relationships - in practice, engineering tolerances would
require that upper and lower limits be defined around either line (or any other acceptable line)
based on aeromedical and engineering discussion.

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LEAFLET 4

OXYGEN SYSTEMS

PRESSURE LOSSES IN OXYGEN DELIVERY SYSTEMS

1 INTRODUCTION

1.1 This Leaflet gives some guiding principles on the design of oxygen systems to minimise
pressure loss, and an acceptable technique for measuring the pressure loss in the oxygen delivery
system, i.e., between the regulator outlet and the mask plug socket.

2 GENERAL PRINCIPLES

2.1 The main causes of pressure losses in low pressure oxygen systems and methods of
minimising these are as follows:

(a) The viscous and turbulent resistance to flow ("drag") occurring in pipes and other
components. Generally, these can be considerably reduced by a small increase in bore,
particularly in long pipe runs (the approximate pressure loss varies inversely as (inside
diameter) 4.5). However, care should be taken if local increases in bore are employed, such
as in a long run of piping, otherwise the gains will be offset by the additional losses incurred
due to change in areas.

(b) Sudden variations in gas velocity caused by change in cross-sectional areas,

particularly increases in area causing a sudden reduction in velocity. The "constant area" rule
should, therefore, be applied as far as possible. When this cannot be done, transition from
one cross-sectional area to another should be as gradual as possible, particularly for any
increase in area; the aim should be to provide an included angle of transition not exceeding
15°.

(c) Changes in the direction of flow caused by bends, etc. These should be eliminated
and made as gradual as possible, both for bends in the piping and changes in direction in
components.

3 MEASUREMENT OF PRESSURE DROP

3.1 An acceptable type of rig for measuring the pressure drop is illustrated in Fig. 1
Measurements should be made for each regulator separately.

3.2 The pressure drop in the system when the flow of oxygen is 120 litres per minute corrected
to NTP conditions is determined as follows:

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(a) Measure the following quantities during the test:

Pd --- in. wg total pressure drop recorded between piezos A and B at flow FM
FM --- litres/min volume flow as recorded on flow meter (this should be
approximately 120 litres/min)
δPA -- in. wg - pressure drop in downstream limb of piezometer A (obtained in
separate test by butting back to back two similar piezos together and
halving the pressure drop measured between them at 120 litres/min)
δPB -- in. wg - pressure drop in upstream limb of piezometer B (obtained in
separate test as in the case of δPA above)
pB -- psia back pressure at piezometer tapping B
Pa -- psia atmospheric pressure
Ta – oK atmospheric (and gas) temperature

(b) Calculate the volume flow corrected for discharge conditions at B (i.e. FB) by
correcting FM for any flow meter scale error and for density differences between the flow
meter and point B.

(c) The pressure drop in the system corrected to 120 litres per minute flow at NTP
conditions, assuming flow is turbulent (Reynolds Number greater than 2,500) is then given by
the formula:

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LEAFLET 5

OXYGEN SYSTEMS

TESTS ON LIQUID OXYGEN SYSTEMS

1 INTRODUCTION

1.1 This leaflet gives a recommended test schedule for the testing of liquid oxygen systems for
aircrew breathing.

2 RIG TESTS

2.1 GENERAL

2.1.1 Test should first be made on working rig called for in Part 1, Section 1, Clause 1.1.2 and
MAP RA 5211.

2.2 PRESSURE/FLOW TESTS

2.2.1 Before connecting the converter and regulator to the system, pressure/flow tests should be
carried out using gaseous oxygen.

2.2.2 The pressure at the inlet to each regulator should be measured when the supply, pressure
at the converter outlets is equal to the normal design working pressure of the converter and the
maximum flow required by Section 1, Clause 1.4 is supplied.

2.3 ALTITUDE CHAMBER TESTS

2.3.1 Where pressure jerkins are provided, it should be established with the rig (from the
converter up to but not including the regulator) in an altitude chamber that the maximum demand
from the simultaneous inflation of all jerkins can be obtained 15 minutes from the commencement
of recharging at all relevant altitudes and under all flight conditions, including, where applicable,
inverted flight.

2.3.2 The test procedure should be obtained from the Aeroplane Equipment Installation
Information (AEII) for the liquid oxygen system and the inflation flows from the AEII for the
regulator.
Note: It may be necessary with certain installations to reduce the overall size of the working rig by
coiling pipes when carrying out altitude chamber tests.

3 INSTALLATION TESTS

3.1 Before making the installation tests, the converter(s) should be submitted to an evaporative
loss test on the bench, the converter(s) should be vented to atmosphere and loss recorded by
weighing.

3.2 The following tests should be made in temperature conditions on the prototype installation
and on any modified installation:

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(a) prove the functioning of the relief valves at the correct pressure, by pressurising the
system with gaseous oxygen,

(b) if the filler connection/disconnection is satisfactory, fill the system with liquid oxygen.
(With a system incorporating rapid stabilisation, fill to the agreed mark on the contents gauge
and after 15 minutes check that the gauge reads "FULL"; with un-stabilised liquid, fill until the
contents gauge reads "FULL" and ensure that this can be reached before overflow occurs),

(c) empty the container and ensure that all contents gauges return to zero,

(d) recharge the system with liquid oxygen and record the elapsed time from
commencement of filling until the working pressure in the system is reached,

(e) check that all contents gauges indicate "FULL", and

(f) after 30 minutes under pressure, examine the system for leaks, which are usually
indicated by local frosting, and examine the vent outlet for indication of leaking valves.

4 FLIGHT TESTS

4.1 The following flight tests should be made on the prototype installation and on any modified
installation, in a suitably instrumented aeroplane:

(a) empty and recharge the converter, check the time from commencement of
recharging until the working pressure is reached,

(b) take-off as soon as the working pressure is reached,

(c) record the pressure and temperature at the regulator inlet under all flight conditions
while carrying out a flight plan which has been agreed by the appropriate Project Team
Leader, Note: As soon as possible after take-off, an inverted flight test or a test under
negative g conditions as appropriate to the aeroplane type should be made and the pressure
should not fall by more than 68 kPa,

(d) descend and land, and examine the system for frost decomposition and traces of
melted frost which may have collected in the structure.

5 TESTS OF INSTALLATIONS IN PRODUCTION AND OTHER DEVELOPMENT

AEROPLANES

5.1 A pressure/flow test should be made on each and every system to prove that under the
most adverse conditions the pressure and flow to each regulator remains adequate.

5.2 All the ground tests in Para. 3 should be made on each and every oxygen installation.

5.3 No special flight tests are needed, but the behaviour of the systems should be checked
during other flight tests.

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LEAFLET 6

ICE PROTECTION

PRECAUTIONS TO PREVENT WASTE WATER LEAVING AEROPLANES AS ICE

1 INTRODUCTION

1.1 Investigation into dangerous falls of ice from aeroplanes in flight has led to the conclusion
that, in many instances, the ice was formed by the discharge of waste water on to the cold airframe
structure; the ice then became detached in lumps when the aeroplane entered warmer air.

1.2 It is, therefore, required on some military transport aeroplanes that waste water shall not be
so discharged as to leave the aeroplane in the form of lumps of ice and this Leaflet gives
recommendations for compliance with this requirement.

2 COLLECTION OF WASTE WATER

2.1 The surest way of meeting the requirement is to collect all waste water (by use of a soil
tank, if necessary) and retain it in the aeroplane.

3 DISCHARGE OF WASTE WATER IN FLIGHT

3.1 Where the method of Para. 2.1 is not adopted, the waste water should be prevented from
freezing in the outlet pipe and should be discharged clear of the aeroplane.

3.2 The chances of ice forming on the airframe structure will be further reduced if the waste
water is discharged in large amounts infrequently rather than in small amounts frequently. Hence it
is desirable to store, at a suitable temperature, as much waste water as feasible and then to
discharge it quickly.

4 RECOMMENDATIONS FOR DESIGN OF WASTE WATER OUTLETS

4.1 The water drain mast should either be heated to provide de-icing, or manufactured of low
thermal conductivity materials if not de-iced.

4.2 The outlet pipes should cause the least possible interference with the air flow and should
be long enough to clear the boundary layer (see also Para 3.1).

4.3 A design of outlet which has proved satisfactory on large transport aeroplanes is given in
Fig. 1. The pipe is a polythene moulding having a trailing edge thickness of 1.6 mm.

5 FLIGHT TESTS

5.1 Where it is doubtful whether waste water discharged in flight will be carried clear of the
aeroplane, flight tests should be made with whitewash on the aeroplane surfaces near to the outlet
and colour dye in the water.

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LEAFLET 7

ICE PROTECTION

ICING CONDITIONS

1 INTRODUCTION

1.1 The design atmospheric icing conditions are defined in this leaflet.

1.2 These conditions represent standards which an aeroplane and its equipment may be
required to meet in order to ensure the ability to fulfil the Service operational requirements in
inclement weather.

1.3 The extent to which these conditions are applicable to any particular aeroplane or
operational role fit will be stipulated in the Aeroplane Specification or agreed with the Project Team
Leader.

1.4 These design conditions have been based on statistical analyses of thousands of
observations obtained over several decades in many geographic regions.

1.5 In natural icing the conditions experienced are unlikely to correspond precisely to any one
of these design conditions; indeed, natural icing conditions may well be mixed, and may change
rapidly in a short distance (or time).

2 METEOROLOGICAL FACTORS INFLUENCING ICING

2.1 Atmospheric icing is a complex phenomenon influenced by many interdependent factors.

The more significant of these are as follows.

2.2 Ambient Temperature: Icing can occur at ambient temperatures between -40°C (-80 in ice
crystal cloud) and +5°C or above. At these higher temperatures icing can occur in the following
situations:

(a) in engine air intakes, carburettor venturis, and other places where the air
experiences adiabatic cooling due to expansion of the airflow, and

(b) on external parts of the aeroplane which have been subjected to cold soak at
altitude and which are then exposed to moist air during the descent.

2.3 Liquid Water Content: The design icing conditions specify liquid water content (LWC) from
0.15 to 5 gm/m3. The higher concentrations are associated with the higher temperatures within the
icing range. The liquid water content in layer (stratiform) cloud seldom exceeds 1 gm/m3, whereas
much higher concentrations can occur in convective (cumuliform) cloud. However, convective
cloud is normally much more limited in horizontal extent than layer cloud.

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2.4 Ice Crystal Content: Clouds containing ice crystals can occur at temperatures from 0°C
down to -80°C, and at altitudes up to 18000 m (60000 feet). At temperatures down to, and possibly
below, -20°C (but no lower than -40°C) the ice crystals usually occur in combination with super
cooled water. In ice crystal cloud, accumulation of slush can occur even on heated surfaces such
as turbine engine air intakes, and on pitot heads and other sensor probes.

2.5 Droplet Size: The median volume diameter droplet size is a function of the icing condition.
Reference to Table 1 shows a range of 10 microns (in freezing fog) to 5000 microns (in freezing
rain). Within each condition the droplets are assumed to be distributed in size about the median
diameter. The distribution is defined in Table 2. Water droplet and ice crystal size has a marked
influence on both the extent and the severity of ice accretion on a body (see paragraph 3.5). In
determining the impingement areas for Icing Conditions I and II of Table 1 it is acceptable to
assume a droplet size of 50 microns. In calculating accretion characteristics it is recommended
that, in addition to using the distribution defined in Table 2, the analysis should also be performed
assuming a constant droplet diameter (equal to the median volume diameter appropriate to the
Condition).

2.6 Pressure Altitude: The altitude ranges in which icing can occur depend considerably on the
condition. For example, freezing fog often extends no higher than 15 m (50 feet) above ground
level, and seldom above 100m (330 feet). However the severe icing associated with the
Intermittent Maximum Condition (normally associated with cumuliform cloud) is most likely to be
experienced between say, 1200 and 12000 m (4000 and 40000 feet).

3 THE INFLUENCE OF ICING CONDITIONS ON ICE ACCRETION

CHARACTERISTICS

3.1 The distribution, and the type, of ice accretion is strongly dependent on air temperature, but
is also influenced by the following factors:-

(a) the shape, size, and attitude relative to the airflow of the accreting body,

(b) the surface temperature and thermal conductivity of the body,

(c) the liquid water or ice crystal content,

(d) the size of the water droplets, and the airspeed,

(e) the atmospheric pressure at the altitude the aeroplane is flying, since this
determines the vapour pressure, and hence the rate of evaporative heat losses.

3.2 At the lowest temperatures in the icing range the super cooled water freezes on impact on
a cold surface, normally in a narrow band centred on the stagnation point, to form rime ice. This is
usually opaque, and sharply pointed.

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3.3 At the highest temperatures in the icing range, close to 0 degrees Celsius, the super cooled
water does not freeze immediately on impact but runs back, losing heat by evaporation, conduction
and adiabatic cooling until ice forms. This is glaze ice which may be smooth and fairly transparent.
Since the runback of the impinging water occurs on both sides of the stagnation point, the ice
formation grows as two horns which may be separated by a relatively ice-free area.

3.4 At intermediate temperatures in the icing range the extent and type of ice accretion lie
between those described in 3.2 and 3.3. The ice may take on an arrowhead or a mushroom shape,
and the ice texture may range from rime through glime (or cloudy ice) to glaze, depending on the
temperature.

3.5 The combined influences of body shape and size, liquid water content, water droplet
diameter, and airspeed determine the rate at which super cooled water impacts with the surface,
and the extent of the impact area. The larger the water droplets, the smaller the body, and the
higher the airspeed - the greater the 'rate of catch' per unit frontal area, and thus the greater the
rate of ice accretion if the air and surface temperatures are sufficiently low.

Note: Great caution should be exercised in extrapolating ice accretion rate data, whether
obtained from calculation or by test. The very process of accreting ice modifies the profile of the
accreting body. This alters the rate of catch, as well as changing such factors as the rate of
adiabatic cooling due to flow acceleration around rapid changes of profile. These factors modify the
subsequent rate of ice accretion, usually resulting in an accelerating rate of ice accumulation.

4 RELATIONSHIP BETWEEN ICING SEVERITY STANDARDS

4.1 There is no simple or precise relationship between the design icing conditions defined in
this leaflet and the Meteorological Office forecasters' terminology, viz, "light", "moderate" and
"severe".

4.2 Reference to Table 1 shows that the Liquid Water Contents (LWCs) appropriate to the
Continuous Maximum and Intermittent Maximum design icing conditions are functions of OAT and
altitude. For the Continuous Maximum conditions LWC ranges from 0.9 gm/m3 at +5°C to 0.2 at -
30°C, whereas the Intermittent Maximum LWCs range from 2.7 at +5°C to 0.2 at -40°C.

4.3 As a rough guide the Meteorological Office forecasters' "light icing" covers LWCs up to
approximately 0.5 gm/m3, "moderate icing" covers LWCs from 0.5 to 1.0, and severe icing" from
1.0 to 4.0 gm/m3.

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TABLE 1
DEFINITION OF DESIGN ICING CONDITIONS

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TABLE 1(Cont.)
DEFINITION OF DESIGN ICING CONDITIONS

Notes

1. At altitudes below 1200 m (4000 feet), water content is assumed to decrease linearly with
decreasing altitude to zero at sea level, except that below 300 m (1000 feet) the content for 300 m
(1000 feet) applies.
2. In determining the limits of the impingement area, droplet sizes up to 40 microns shall be
considered.
3. At altitudes below 4500 m (15000 feet), water content is assumed to decrease linearly with
decrease of altitude such that, if extrapolated to zero altitude the content would be zero.
4. In determining the limits of the impingement area, droplet sizes up to 50 microns shall be
considered.
5. In the temperature range 0 to -20°C the ice crystals are likely to be mixed with water
droplets (with a maximum diameter of 2 mm) up to a content of 1 gm/m3 or half the total content,
whichever is the lesser, the total content remaining numerically the same. Below - 20°C all the
water present may be assumed to be in the form of ice crystals.
6. When the horizontal extent is shown as 'continuous' it is acceptable to show that the
aeroplane functions satisfactorily during 30 minutes continuous exposure to the condition.
7. See Figure 1 - Empirical Relationship between Snow Concentration and Observed Visibility
- for a guide to the severity of snow conditions.

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TABLE 2
RANGE OF DROPLET SIZES

Note:

1. The droplet sizes quoted in Table 1 are the volume median diameters (dV) for the
distribution, shown in Table 2; dF is the particular drop diameter under consideration.

Fig. 1 - Empirical Relationship between Snow Concentration and Observed Visibility

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LEAFLET 8

ICE PROTECTION

ICE PROTECTION SYSTEMS

1 GENERAL

1.1 The design or selection of ice protection systems requires consideration of the following:

(a) the Design Icing Conditions in which the aeroplane is required to be capable of
meeting its Service operational requirements,

(b) the areas over which ice accretion can occur, and

(c) the rate at which ice accretion can occur.

1.2 The Design Icing Conditions are defined in Leaflet 7. The Aeroplane Specification will
stipulate which of the conditions shall be met, and for what durations.

2 AREA OF ICE ACCRETION

2.1 The area over which ice can accrete on an object is determined by:

(a) its shape, size and disposition relative to the airflow,

(b) its surface temperature when exposed to the icing condition, and

(c) the water droplet or ice crystal size and air velocity.

2.2 The shape of the ice accretion is largely a function of icing surface temperature and
freezing fraction. These are complex functions of the parameters (a) and (c) above, and of
liquid/solid water concentration, ambient pressure, temperature and relative humidity. A concise
relationship between accretion shape and surface temperature and freezing fraction has not been
established. In general a freezing fraction of unity and low icing surface temperature gives a sharp
pointed rime ice growth. As temperature rises to zero the ice type changes to glaze ice and the
accretion shape changes with decrease in freezing fraction through arrowhead, blunt arrowhead
and mushroom to double or single horned at low freezing fraction.

2.3 For aeroplanes the ice accretion area and chord wise limits can be calculated using
suitable computer programs, (see for example Ref. 11 for aerofoils). Further relevant information is
contained in Refs. 1 to 10 inclusive.

3 RATE OF ICE ACCRETION

3.1 The rate at which ice may form on a surface is determined by:

(a) the shape, size and disposition of the surface relative to the airflow;

(b) the water droplet or ice crystal size;

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(c) the airspeed, ambient air temperature and pressure;

(d) the amount of water blown off;

(e) the liquid water or ice crystal content;

(f) the surface freezing fraction (which depends on the balance of the heat transfer
from the surface); and

(g) the density of the resulting ice.

Items (a) and (d) control the surface water catch efficiency, which with item (e) determines the total
rate of catch of water/ice. Item (f) determines how much of the collected water/ice freezes.

3.2 The rate of catch of water drops on aerofoil sections can be calculated using computer
programs. In the absence of a suitable program the methods of Ref. 1 may be used. An allowance
should be made for the water which blows off and therefore does not require evaporation, and for
kinetic heating. Fig. 1 shows the percentage of the calculated catch to be evaporated for the
continuous maximum icing condition. The kinetic temperature rise should be taken as half the
value assumed for dry air conditions. The rate of catch of droplets on a radome may be assumed
to be equal to that of a sphere with the same frontal area.

4 NOTES ON ICE PROTECTION SYSTEMS

4.1 GENERAL

4.1.1 Ice protection systems can be either of two types:

(a) Anti-icing Systems, where the surface is maintained free from ice accretion at all
times, or

(b) De-icing Systems, where accretion is allowed to occur and is periodically remove
before its effects, and those of the shed ice, are hazardous.

4.1.2 Since there may be an unacceptable operational penalty associated with the provision of
ice protection systems, and also with flight in icing conditions even when protection systems are
fitted, it is recommended that systems be discussed and agreed with the Project Team Leader at
an early stage in the design.

4.2 ANTI-ICING SYSTEMS

4.2.1 Anti-icing can be achieved by continuous heating, employing either electrical or hot air
systems, or by the use of freezing point depressant fluids.

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4.2.2 If a continuous heating method is employed, either an adequate surface should be

maintained above 0°C, so that freezing of any run-back water which might cause a hazard is
prevented, or sufficient heat should be applied to the wetted area to evaporate all impinging water
not blown off. Computer programs have been developed to perform these calculations (Ref. 11.) In
the absence of a suitable program the heat required may be calculated by the methods given in
References 2 and 6. The amount of heat should be evaluated for the droplet sizes given in Leaflet
7, Table 2 with due allowance for the latent heat of fusion of ice for the mixed and snow conditions.

4.2.3 When methods employing freezing point depressants only are used as anti-icing systems,
the quantity of fluid required to depress the freezing point below the local temperature of the
surface should be calculated. The amount of fluid applied should be 1.25 times the calculated
quantity to allow for local variations both in the mixing process and in distribution (see Ref 7). The
pumps should be duplicated, each capable of providing the full requirements. Allowance should
also be made for the mixed conditions given in Leaflet 7, Table 1 and for the effects of the fluid
when flying in snow.

4.3 DE-ICING SYSTEMS

4.3.1 De-icing can be achieved by the application of heat through electrical or hot air systems to
weaken the bond between the ice and the surface, by the use of freezing point depressant fluids,
by mechanical means or by the use of low adhesion (ice-phobic) coatings or pastes.

4.3.2 If a heating method is employed, care should be exercised in the selection of heating power
and duration to ensure that catch on the surface, after shedding of the ice, does not result in an
unacceptable amount of run-back icing on unprotected parts of the surface. Methods of calculating
heating requirements are given in Ref 7.

4.3.3 Where pneumatically inflatable de-icing boots are employed, the correct sequencing and
duration of the inflation cycles needs to be established, appropriate to the range of icing conditions
in which the aeroplane may be required to operate.

4.3.4 If electro-impulse methods are used to provide de-ice protection, the airworthiness of the
resulting structure must be fully substantiated.

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FIG. 1 - ALLOWANCE FOR 'BLOW-OFF' OF WATER CATCH BASED ON CONTINUOUS MAXIMUM

ICING CONDITIONS

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REFERENCES

To assist the design of ice protection systems, the following references should be consulted:

Ref No Author Title, etc

1 Bigg F G Impingement of water droplets on aerofoils.
Baughen J E RAE Technical Note ME 208, 1955.
2 Hardy J K Protection of aircraft against ice. RAE Report SME
3380, 1946.
3 Brun E A Icing problems and recommended solutions.
Agardograph 16, November 1957.
4 Dorsch R G Impingement of water droplets on an ellipsoid with
Brun R L fineness ratio 5.
Gregg J L NACA TN 3099, 1954.
5 Coles W C Icing limit and wet-surface temperature variation
for two airfoil shapes under simulated high speed
flight conditions.
NACA TN 3396, 1955.
6 Bowden D T Engineering summary of airframe icing technical data
et al FAA Tech Report ADS-4 1964.
7 Messinger B L Equilibrium temperature of an unheated surface as a
function of airspeed.
J Ae Sc January 1953 Vol 20 No 1
8 Cansdale J T Calculation of surface temperature and ice accretion
rate in a mixed water droplet - ice crystal cloud.
RAE Technical Report TR 77090, 1977.
9 Cansdale J T Ice accretion on aerofoils in 2-dimensional
Gent R W compressible flow - a mathematical model.
RAE Technical Report TR 82128, 1982.
10 Gent R W Calculation of water droplet trajectories about an
aerofoil in steady, 2-dimensional, compressible flow.
RAE Technical Report TR 84060, 1984.
11 Gent R W TRAJICE2 - a combined droplet trajectory and ice
accretion prediction program for aerofoils.
RAE Technical Report TR90054, 1990.

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MILITARY SPECIFIC SYSTEMS

3.0.1 Requirements are provided to cover the following aspects:

3.1 - Armament Installations – General 2

3.2 - Armament Control Systems 4
3.3 - Gun Installations 15
3.4 - Installation of Explosive devices. 19
3.5 - Air to Air Refuelling 20
3.6 - Arresting Hooks. 43
3.7 - Installations for Emergency Recovery from Stall and Spin. 52
3.8 - Target Towing Installations 61
3.9 - Reduction of Vulnerability to Battle Damage 76
3.10 - Protection of Aircrew against Conventional weapons 79
3.11 - Protection from the Effects of Nuclear Explosions, Laser weapons, Chemical and Biological Warfare Agents 82
3.12 - Aircrew Equipment. 91
3.13 - Brake Parachute Installations. 92
3.14 – Integration of Stores 96
3.15 – Not Issued 101
3.16 - Defensive Aids Systems (DAS) 102

Tables 115

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3.1 ARMAMENT INSTALLATIONS - GENERAL

3.1.1 Clauses 3.1 to 3.3 state the airworthiness requirements for the armament installations of all aircraft. Armament installations, for the purpose
of these requirements, comprise the weapon installations, associated armament equipment, electrical installation and associated software
concerned with carriage, monitoring, control, release or firing, or jettison of all stores. (See note). Where stores are carried internally, all
mechanisms required for the operation of weapon bay doors or weapon release system shall be considered part of the armament installation.

Note: "Stores" includes all explosive, non-explosive and pyrotechnic items that can be carried on or in, and released or jettisoned from, an
aircraft.

3.1.2 The requirements of clause 3.3 apply to all aircraft which are fitted with fixed or free (movable) guns. A fixed gun is controlled by the aircraft
armament installation and may be mounted internally or externally. A free gun is an airframe-mounted gun that is manually operated. A gun
installation includes the guns, ammunition containers and ammunition, feed and ejection mechanisms, and spent links or empty cases collection or
disposal systems.

REQUIREMENT COMPLIANCE GUIDANCE

GENERAL
3.1.3 The armament system shall be such The armament installation shall be such that a These requirements aim to ensure that the
that no single fault or failure shall adversely affect single fault or failure shall neither: equipment has the degree of safety and reliability
the safety or operation of the system. adequate for the carriage of stores on or in
(a) Prevent release or jettison of the aircraft. Notwithstanding this, the cumulative
store(s) when required. probability for the inadvertent or uncommanded
release of a store due to a technical fault, or the
(b) Result in inadvertent or cumulative probability of a technical failure
uncommanded release or jettison of the preventing the intended release of a store, should
store(s). be no worse than 1 in a million per flying hour
(probability of occurrence of 1 x10-6 per flying
(c) Prevent the weapon being hour).
released live and in the correct condition,
when required.

(d) Result in arming of the weapon

before release.

(e) Prevent the weapon being made

safe after having been selected live.

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REQUIREMENT COMPLIANCE GUIDANCE

The Aircraft Design Organisation shall

demonstrate that the armament system shall
function in accordance with specified
requirements under all required operational
conditions.
3.1.4 The Aircraft Design Organisation Evidence from modelling, simulation, working
shall demonstrate that: rigs, wind tunnel tests, ground and air trials,
would be acceptable, subject to the agreement
(a) Stores can be carried and armed, of the Project Team Leader (PTL). Flight trials
released, fired and jettisoned within the representing operational scenarios are the
parameters specified. preferred source of evidence for the
performance of the armament installation.
(b) Stores can be loaded and unloaded
easily when the aircraft is at rest on the
ground.
3.1.5 The aircraft Design Organisation Universal Armament Interface /NATO Armament
shall prepare an Interface Control Document for Interface (Future) (UAI/NAI (Future))
each store carried on the aircraft that meets the requirements are to be addressed in accordance
requirements of the store DO. with weapon specification.
3.1.6 All store stations shall be fitted with Currently, MIL STD 1760D is the required issue
an Aircraft Station Interface (ASI) that conforms for integration. However, 1760E is being used to
with the latest issue of MIL STD 1760 unless implement UAI/NAI and has both hardware and
specified to the contrary. software implications for integration.

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REQUIREMENT COMPLIANCE GUIDANCE

3.2 ARMAMENT CONTROL SYSTEM
ARMAMENT ELECTRICAL POWER SUPPLIES
3.2.1. The Armament System shall be (a) A failure mode analysis shall be carried (a) Armament power supplies should where
immune to single point failure of the electrical out to demonstrate the immunity of the system possible be duplicated and separate from the
power supply. to single point failure. main aircraft bus bars.

(a) Armament electrical safety shall not (b) It shall be possible to operate the (b) The armament wiring should not normally
to be degraded by any form of load shedding emergency jettison circuits correctly despite a be formed into cable assemblies with wires that
or any other legitimate action. complete failure of the normal generating are not associated with armament circuits.
system. Services other than armament services should
(b) Electrical power shall always be not be routed through armament junction boxes,
adequate for every safety purpose. (c) A load analysis shall be carried out to terminal blocks or connectors.
demonstrate the immunity of the Armament
(c) The status of armament electrical Electrical Power Supplies to transients. (c) Each side of a duplicated circuit should
power shall be clearly indicated to the crew. have its own independent frame connections and
both should be independent of the frame
connections of other systems. See also Part 1,
Section 6, leaflet 14.
ARMAMENT CONTROL SYSTEM SAFETY FEATURES
3.2.2 The Armament Control System shall (a) In addition to the operation of the trigger, (a) The second break is sometimes called the
incorporate protection against human factors and a minimum of 2 crew controlled layers of safety “Late Arm switch” and should be positioned so
technical failure causing inadvertent firing or shall be required (e.g. MASS and Late Arm). that it can be closed during the final stages of an
release of weapons or countermeasures. Control These controls shall be guarded against attack. It is meant to minimise the consequences
of this protection shall be provided to such crew accidental operation and at least one shall be of a short circuit fault across the release switch.
members as may be specified in the Aircraft implemented solely in hardware and shall not The second switch may be software controlled
Specification and shall involve the progressive merely interrupt a discrete input to a processor. and the software must be developed in
removal of safety layers from the point at which the accordance with Def Stan 00-56. Hands on
aircraft is on the ground until the final stages of an (b) The preferred solution for the hardware Throttle and Stick (HOTAS) controls do not
attack and trigger press. break is a rotary switch guarded against normally provide guards for switches.
unintentional operation by a 'gate' system such In the auto mode the release switch/button may
that it shall not be possible to select LIVE act as an enabling device or commit button
without a positive first action to allow passage thereby allowing the weapon aiming computer to
through the 'gate'. Detent indexing should be initiate release pulses which may be derived from
provided at each selectable position. the main computer; if at any time this switch is

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REQUIREMENT COMPLIANCE GUIDANCE

opened, store release is interrupted.
(c) Where more than one position of the
hardware break control is required to energise
armament system circuits, the fusing and
arming, firing, release and jettison circuits shall
not be made live until the control position
furthest from SAFE is reached.

(d) For highly autonomous remotely piloted

air vehicles where operator interaction may be
limited, the level of protection provided against
inadvertent/accidental operation needs to be the
same as that provided through the use of
guarded hardware switches on manned aircraft.

(e) Controls and switches shall be

positioned to minimise the possibility of
inadvertent operation.
3.2.3 Arming, release or jettison shall not
be possible with the aircraft on the ground unless
specifically required by the aircraft specification.
ARMAMENT ELECTRICAL CIRCUITS

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REQUIREMENT COMPLIANCE GUIDANCE

3.2.4 Identification and protection:

(a) All cables shall be uniquely With the exception of the High Bandwidth, the (a) To provide unique identification of the
identified as part of the armament system, Low Bandwidth and the Multiplex Data Bus armament control system wiring all the cables
and shall be run in protective encasements cables that form part of the MIL-STD-1760D should be coloured red. This identification
(e.g. ducting, conduit or sleeving). installation, all armament control system cables ensures that any damaged cable can be
shall be coloured red. immediately identified as safety critical and
appropriate action taken to ensure that repairs are
promptly expedited and the system re-certified.

(b) The cables and encasements should be

(b) All wiring runs external to the See also leaflet 2. coloured red, or encasements sufficiently
aircraft shall, without exception, be protected transparent for the colour of the enclosed cables
by encasement. Where camouflage is to be easily identified.
necessary, marking may be by 2 red lines of
2 mm width, 180° apart running along each
cable.
3.2.5 Cable terminations in connector A circuit analysis shall be carried out to ensure (a) It is preferable to use an entirely separate
systems that control release lines or any form of that the safe operation of the armament connector for the sensitive circuits, and to run the
electro-explosive device (EED.) shall not occupy installation is not compromised. associated wiring in separate looms where this is
pin or socket connections adjacent to pins or possible. The close proximity of standing and
sockets carrying standing voltages. switched voltages for armament and non-
armament services, in connector systems, makes
it necessary to ensure that the armament circuits
are properly segregated from all other services,
particularly at the terminal arrangements on
switches and other electrical components.

(b) The term standing voltages includes not

only those which are present throughout the flight,
but also those which are switched on sometime
before release or firing is intended, for example,
fusing or station selection supplies.
3.2.6 When the use of 2 or more The prevention of cross connection solely by
connectors of the same size in the same location is cable or loom clipping is not permitted.

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REQUIREMENT COMPLIANCE GUIDANCE

unavoidable, shell oriented connectors to different
orientation, or a different insert arrangement, shall
be used.
3.2.7 Cable splices shall not be used.
3.2.8 Cables associated with the circuits See also Leaflet 2.
for firing, fuzing, release and jettison are to be
identical in all aircraft of the same type, mark, role
and mod state.
JETTISON SYSTEMS
3.2.9 Provision shall be made for:

(a) Emergency Jettison, and (a) The initiation of jettison shall be as See Definitions in Part 0 for Emergency Jettison
simple as possible and involve the minimum and Selective Jettison.
(b) Selective Jettison. number of control actions on the part of the
aircrew. Where a safe order of jettison is See also leaflet 3.
required, this must be predetermined and
automatic. Where relevant, operation of weapon Selective Jettison only for stores in weapon bays.
bay doors shall be automatic.
Selective Jettison only for forward firing jettison
(b) The jettison control(s) shall be systems.
independent of the normal weapon release
control(s). Emergency jettison shall always be under the
control of the pilot or operator except for
(c) Emergency Jettison shall be highly/fully autonomous vehicles where control
implemented by high integrity electronic shall reside with the vehicle management system.
hardware that meets the requirements of Def
Stan 00-56.

(d) Emergency Jettison shall always be

under the control of the pilot.
RELEASE SYSTEMS
3.2.10 Arming initiation:

(a) Arming initiation of the store shall

not be possible until separation from the

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REQUIREMENT COMPLIANCE GUIDANCE

release unit or launcher is irrevocable,
although the store may still be connected to
the aircraft by any of the devices listed in
Clause 3.2.24

(b) No single failure shall allow arming

initiation until separation is irrevocable.
3.2.11 Automatic release The pilot should not have to intervene until after
the attack manoeuvre has been completed and
(a) Automatic release of stores shall be the aircraft has recovered to a safe height and
carried out in a predetermined order to speed.
ensure that no dangerous asymmetric or out-
of-trim conditions occur.

(b) The system shall also automatically

take into account the possibility of a failure to
release a weapon in a "stick".
3.2.12 Release, arming or jettison of Opening and closing should be considered as
internally carried stores shall not be part of the weapon release sequence; i.e.
possible until the weapon bay doors (WBD) automatic upon Trigger Press (TP).
are fully open.
3.2.13 An indication shall be provided to This requirement also applies to stores fired or
the aircrew if any of the stores selected for release released from dispensers or launchers.
or firing cannot be released or fired, or suffers Manual operation of the WBD should be available
hangfire or misfire. in the event of hang-up/misfire.
ELECTRO-MAGNETIC INTERFERENCE
3.2.14 EMI shall not initiate the release or (a) Prove that no EED can receive sufficient (a) EMI includes EMC, EMP, lightning,
stimulate generation of a valid signal of any kind in signal from EMI to reach the "No Fire Threshold" electrostatic and radio frequency hazards from
the armament installation, or prevent normal as defined in OB ProcP101 with the appropriate both on-board and external sources.
operation of any armament function. safety factor from either the onboard sources or
the defined external environment. (b) Where digital data transmission systems
form part of the armament control installation, it
(b) Prove that signals induced by EMI may be necessary to run the signal lines
cannot themselves be of such strength that they separately from the main armament cable runs, to
can cause any operation of any kind in the minimise electromagnetic interference.

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REQUIREMENT COMPLIANCE GUIDANCE

armament installation, or prevent a desired
operation, in accordance with OB ProcP101.
3.2.15 The armament installation shall not Prove that the strength of emissions from the
cause EMI to other onboard systems. armament installation is less than the limit to
which other on-board systems are qualified, and
in accordance with OB ProcP101.
IN-SERVICE TESTING
3.2.16 Provision shall be made for testing The need for end-to-end armament installation,
in service to a strategy agreed by the PTL. and “no-volts” testing should be considered.
STORE STATIONS
POSITION OF STORE STATIONS
3.2.17 The position of each station shall be This requirement shall be met for all the release The effects of discarding weapon containers,
such that when a store is fired or released, it falls, modes defined in the Aircraft Specification. empty cases or links, frangible covers and
is ejected or propelled clear of the aircraft, The requirement shall meet the required static protective devices at launch, firing or jettison,
adjacent stores and equipment. clearances declared by the aircraft Design including the random impact of discarded pieces
Organisation. on the aircraft structure, shall be considered.
3.2.18 The positions of the store stations Where applicable the aircraft structure should be
shall be chosen to avoid high temperature efflux designed to tolerate the effect of blast and debris
from weapon propulsion system, rockets or guided from the weapons, and the effects of high velocity
missiles impinging upon other stores and release gas streams.
systems.
3.2.19 The positions of the store stations If necessary protection shall be provided.
shall ensure that contamination of the stores and
the release systems by propulsion system exhaust,
fuels, oil or any substance which could adversely
affect the armament system is avoided.
3.2.20 The position of the stations for any
of the specified combination of stores shall be
chosen to minimise undesirable aircraft trim
changes following store release.
3.2.21 Readily accessible strong points Avoid flailing lines using retractors or lose with
shall be provided at store stations for the weapon.
attachment of static lines.
WEAPON BAYS

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3.2.22 A positive means shall be provided (a) Inadvertent or uncommanded operation (a) The operating system is usually under the
for operating the weapon bay doors in flight. The of the weapons bay doors shall not be possible. control of the pilot.
system shall be such that a single failure will not This may require the use of switch guards and
prevent the weapons bay doors from operating. system interlocks etc. (b) Care should be taken in the design stage
to ensure that correct operation of the weapon
(b) This is normally achieved by the use of bay doors will not be prejudiced by:
an emergency operating system which is
independent of the main system and operable (1) Deformation of the aircraft
regardless of any previous setting or failure of structure by external aerodynamic load.
the main system.
(2) Jamming caused by deformation of
(c) When two systems are employed power the doors or surrounding structure.
for both systems shall be from independent
sources. (3) Incorrect alignment of the hinges
under load.
(d) Spring operated doors are not
acceptable (4) Environmental conditions.

(c) For related information see also the

design requirements for undercarriage doors and
locks. See also Part 1, Section 4, leaflet 55.
3.2.23 Automatic operation of WBD should WBD may be operated in flight without weapon
be inhibited until the selection of ‘Late arm’. release.
ARMING CONTROL SAFETY
3.2.24 Installation of static lines, umbilicals,
leads electrical fuse arming (LEFAs), cable
assemblies fuse arming (CAFAs), shear wire
assemblies and lanyards between aircraft and
store shall be laid out such that:

(a) It shall not be possible to make

incorrect mechanical or electrical
connections.

(b) Failure to make any connection

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properly shall not create a situation where
damage to the aircraft can occur during
carriage or after release of the store.

(c) They are not crossed and cannot

become entangled with the lines to other
stores during carriage or after release of the
store.

(d) They can not become crossed or Retract lines or release with weapon.
tangled upon release of the store and cannot
become entangled with, or cause damage to,
the store or the aircraft after release of the
store.
RELEASE UNITS
3.2.25 The release unit shall:

(a) Satisfy the requirements of

STANAG 3441 and 3558 (and STANAG
3605 where applicable) with respect to the
interface between the release unit and the
store,

(b) Be equipped with mechanical

ground safety lock that shall not be possible
to activate unless the release unit is in the
locked condition.

(c) It shall not be possible to remove

the lock following:

(1) Inadvertent firing of the cartridges of

an ERU, until the gas pressure has fallen
to a safe level.

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REQUIREMENT COMPLIANCE GUIDANCE

(2) Inadvertent release of the hook
mechanism actuators of an EMRU, until
the actuators have been reset to the
cocked position.
3.2.26 The operating mechanisms of
release units shall be so designed that a deliberate
action is required to lock or unlock them.
3.2.27 It shall be possible to manually lock This should be achieved by a single tradesman
and unlock a release unit. using the minimum of hand tools (preferably one
that is already in the service inventory) at the
store station.
3.2.28 Activation of the release unit
mechanism shall:

(a) Break the electrical release circuits,

(b) Activate the arming circuits.

(c) Provide “store on station”

information as required.

(d) When an ERU, be provided with a

means for controlling the ejection and
reaction loads.
3.2.29 Indication shall be provided that the
release unit linkage is correctly locked.
MULTI-STORE CARRIERS, SPECIALISED CARRIERS, ADAPTORS AND LAUNCHERS
3.2.30 These carriers shall be provided It should be possible to attach these carriers to
with strong points for store loading, pull-off the release units normally fitted at the store
lanyards, and arming wires to STANAG 3605. station.
3.2.31 The design of Multi-store carriers See general guidance in leaflet 3.
shall ensure that the deflections due to ERU
reaction loads do not introduce release
disturbance or weapon aiming problems.
3.2.32 An electrical bond which meets the

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requirements of Part 1, Section 4, Clause 4.27
shall be established between the hooks of release
units and the airframe.
FIRING, RELEASE & JETTISON
SAFETY CRITICAL SOFTWARE
3.2.33 The safety criticality of the firing, Software assessed as safety critical shall be Software may be used to implement the
arming and jettison system shall be determined in developed in accordance with Def Stan 00-56. sequence and timing of the Selective Jettison
accordance with Def Stan 00-56. See guidance in Section 1.7 function. The criticality of such software should be
determined in accordance with Def Stan 00-56.
3.2.34 The Emergency Jettison function Software may be used to ensure safe symmetrical
shall ensure symmetrical release. release to avoid out of balance conditions.
SAFETY CRITICAL HARDWARE
3.2.35 Electronic hardware assessed as
safety critical shall be developed in accordance
with Def Stan 00-56.
GUN, ROCKET AND GUIDED WEAPON FIRING CIRCUITS
3.2.36 The firing of forward firing weapons (a) Should be read in conjunction with Part 1,
and guns shall not affect the operational function Section 6, clause 6.6 and Part 1, Section 4,
or safety of the aircraft or its propulsion system, or clause 4.26. See also Leaflet 4 and Leaflet 7
of other stores.
(b) The design should mitigate against
propulsion system surge or flameout resulting
from ingestion of weapon efflux, nor should
weapon efflux contaminate other stores.

(c) When considering the provision of a delay

in gun firing circuits, designers should be aware
that a delay between trigger press and gun firing
in excess of 0.1 seconds is operationally
unacceptable.

(d) In the event that propulsion system control

modulation (e.g. fuel dip, bleed, nozzle area
change) is deemed desirable to minimise the

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adverse consequences of weapon wake ingestion
should it occur, adequate provision in the weapon
firing circuit for delaying weapon release or gun
firing to allow implementation of the propulsion
system controlled modulation shall be provided.
3.2.37 Full gun gas purging flow shall be Where the gun bay purging system employs an See also Leaflet 6.
established before firing commences. electrically operated inlet scoop, the scoop
actuator is to be initiated by a switch on the last
safety break and not the trigger itself to ensure
that full gas purging flow is established before
firing commences.
SAFETY INTERLOCK
3.2.38 Means shall be provided to inhibit
the firing or release of stores while any part of the
aircraft (e.g. undercarriage, flaps), its equipment or
other stores obstructs the line of fire or is likely to
be damaged by blast or debris.
3.2.39. In cases where a store is moved Consideration should be given to conditioning the
from a flight carriage position to a firing position, weapon before deployment to firing position.
firing shall be inhibited until the weapon is correctly
positioned.

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3.3 GUN INSTALLATIONS
3.3.1 General

(a) Free or fixed gun installations and (a) Compliance shall be (a) The design should be such that the lateral
gun firing shall not prejudice aircraft demonstrated on a working rig of the and vertical separations between the sight and
structure, other aircraft systems, or aircraft installation or by ground firing trials, and by the gun barrel bore axis are kept to a minimum.
operation. air firing trials. The strength and stiffness of the gun installation
must take account of gun kinematics, and avoid
(b) Demonstration of compliance resonance and vibrations induced by gun
may be supported by structural analysis. operation.

(b) The effects of the installation on the

aircraft aerodynamics and the ease of servicing
and replenishment of the gun installation should
also be carefully considered.

(c) See also leaflet 5.

3.3.2 Location:

(a) The location of the guns shall be See also Leaflet 7.

chosen with particular regard to the effects of
gun firing directly on the propulsion systems,
structures and other systems and indirectly,
by changes to aircraft or equipment
permanent magnetism, on compass detector
units.

(b) Free guns shall be located so that,

throughout the full range of gun movement,
there shall be no obstruction to the line of
sight or restriction of the operator equipped
as provided for in Part 1, Section 4, clause
4.15
HARMONIZATION

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3.3.3 Provision shall be made to It shall be possible to change guns without
harmonise the gun installation, with and without affecting harmonisation. A fixed gun mounting
feed and ejection systems connected. system shall permit angular movement of the
gun to allow for harmonisation adjustments of a
minimum of ½° around the nominal gun line, in
addition to the angular displacement arising from
manufacturing tolerances. The adjustment shall
be continuous throughout the range required
(i.e., not by shims) and its setting shall not vary
through normal service use of the aircraft.
Provision shall be made for locking the system
to maintain the required setting.
3.3.4 A free gun mounting system shall be See also Leaflet 5.
designed to prevent the gun coming into contact
with the airframe.
GUN FIRING/SAFETY
3.3.5 Where a gun-firing trigger is
provided, it shall be protected from inadvertent
operation by a safety device.
3.3.6 For each fixed gun an accessible This may be satisfied by electrical disconnection (a) This is in addition to requirements of the
means shall be provided to inhibit firing on the or a mechanical interlock. hardware safety break at Clause 3.2.2
ground and to enable personnel to ensure that the
system is inoperative. (b) Provision may be required for the ground
firing of guns.
GUN INSTALLATION ENVIRONMENT
3.3.7 The flammable gas from a gun This requirement may be satisfied by use of a
installation shall not present a hazard to the purging system that dilutes the gas as close to the
aircraft. source as possible to prevent the concentration in
all other areas of the installation exceeding 80%
of the Lower Limit of Aircraft Hazard (LLAH)
Leaflet 6.
3.3.8 The design of the gun installation
shall ensure that the maximum and minimum
temperatures and pressures of the guns and

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REQUIREMENT COMPLIANCE GUIDANCE

ammunition are not exceeded throughout the
aircraft-operating envelope.
AMMUNITION CONTAINERS
3.3.9

(a) Access shall be provided in the (a) The design and position of the containers,
chutes for loading and unloading, and in relation to the gun, should minimise friction and
making and breaking of belts where inertia of moving ammunition.
applicable.

(b) The ammunition storage and feed Stoppage rates shall be confirmed by flight trials. (b) Ideally, ammunition should be fed to the
system and the system for disposing of gun from above. In addition, the feed path should
expended cases and links shall not cause be as short as possible, small radii bends should
stoppages. be avoided, and changes of direction should be
minimal. In a fully loaded container there should
be sufficient space to allow the ammunition to be
withdrawn freely.
3.3.10 The routing of ammunition or the
position of the ammunition shall be shown on
ammunition containers.
3.3.11 Provision shall be made for the
drainage of fluids from the containers.
3.3.12 Ammunition container installations
shall incorporate a blast relief mechanism which
will operate, without endangering aircraft or crew,
in the event of an explosion in the container.
EXPENDED CARTRIDGE CASES AND LINKS
3.3.13 The gun installation shall be such See also leaflet 5.
that empty cartridge cases or complete rounds and
links, when not collected cannot cause damage to
the aircraft or stores in any configuration of the
aircraft. When empty cartridge cases or complete
rounds and links are collected their collection shall
not prejudice other requirements.
3.3.14 The cartridge case ejection

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REQUIREMENT COMPLIANCE GUIDANCE

tube/system shall be capable of passing complete
rounds.
3.3.15 Collected cartridge cases links and
complete rounds shall be readily removable by
personnel on the ground.

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 DEF STAN 00-970 PART 13/11
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3.4 INSTALLATION OF EXPLOSIVE DEVICES

INTRODUCTION

3.4.1 The requirements of this Clause apply to all installations of explosive devices in aeroplanes. See Leaflet 8 for guidance.

REQUIREMENT COMPLIANCE GUIDANCE

REQUIREMENTS FOR DESIGN
3.4.2 Only explosive devices approved by See also MAP MA 5106.
the DE&S and DOSG shall be used, and the
installations shall be approved by him.
3.4.3 Explosive devices shall be so Consideration shall be given to making the
installed as to be easily accessible and shall not component in which the explosive is installed
require excessive handling during installation. readily replaceable so that the actual handling of
the explosive charge can be done under servicing
bay conditions.
3.4.4 Particular attention shall be paid to Protective sheaths or grommets shall be used
protecting the insulation of circuits connected to wherever vibration or handling can cause
explosives. damage.
3.4.5 Electrical circuits concerned with the
installation of explosive devices shall conform to
Part 1, Section 6, Clause 6.6
TESTS
3.4.6 The Contractor shall carry out such
tests as the DE&S and DOSG may require,
demonstrating the efficiency and safety of the
explosive devices and their detonating systems.

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3.5 IN-FLIGHT REFUELLING

INTRODUCTION

3.5.1 This Clause sets out the design requirements for the installation of 'in-flight refuelling' equipment in both Tanker and Receiver aeroplanes.

Symbol/ Abbreviation Meaning

RR air-to-air refuelling (receiver)
RT air-to-air refuelling (tanker) (see Part 1, Section 2, Leaflet 1, Para 3.1 and Table 2)

REQUIREMENT COMPLIANCE GUIDANCE

BASIC OPERATIONAL REQUIREMENTS
3.5.2 When the Aeroplane Specification Unless otherwise specified, the Probe and
calls for refuelling in flight, it will lay down the rate Drogue system shall be used and the transfer
of flow, the range of fuels to be used, and limits of system assembly shall be in 'package' form.
speed and height at which the operation is to be
carried out.
3.5.3 For aeroplanes which have an Compliance with this requirement shall be
alternative role, the time to change the aeroplane proved by demonstration.
from this role to that of a Tanker or Receiver shall
be as short as possible.
3.5.4 Opening of the refuelling valves in
the Receiver shall be selected by the crew of the
Receiver aeroplane.
REQUIREMENTS FOR DESIGN
DIMENSIONS
3.5.5 The mating dimensions of the
reception coupling and the nozzle/probe mast shall
conform with STANAG 3447.
3.5.6 A clearance space shall be provided
around the nozzle/probe mast installation in
accordance with STANAG 3447.
COCKPIT CONTROLS AND DISPLAYS - See Part 1, Section 4, Clause 4.19
PRESSURES
IN-FLIGHT REFUELLING COCKPIT CONTROLS AND DISPLAYS

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3.5.7 The following controls and
displays shall be provided for in - flight
refuelling operations in receiver Aircraft:

(a) Controls (See Leaflet 9 Para 6.5.1):

(1) Air refuel selector,

(2) Reset button (where applicable),

(3) Disconnect button (where

applicable)

(b) Display:

(1) Tanks full indication,

(2) Disconnect indication (Master

caution) (where applicable)
PRESSURES
3.5.8 The system shall be designed to See STANAG 3447.
transfer fuel at a static pressure (gauge) not
exceeding 55 lbf/in2 (380 kPa) at the coupling,
without damage to the fuel system or tanks; in
either the Tanker or Receiver. The refuelling
pressure shall be capable of being regulated at all
values of Tanker delivery flow-rate.
3.5.9 The system shall be so designed Note: If the design of the system limits the surge
that any surge pressures developing at any stage pressure to a value lower than the maximum
during refuelling operations are limited to 517 kPa, quoted above, then the lower value can be used
except in multi-tank systems, where simultaneous for design purposes.
closure of any combination of tank valves can
occur, the surge pressure shall not exceed 827
kPa. The strength requirements of Part 1, Section
5, Clauses 5.2.202 and 5.2.203

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3.5.10 Ample inward venting shall be See Leaflet 9.
provided in the tanks of the Tanker to prevent any
risk of tanks collapsing Outwards venting in both
the Tanker and Receiver shall be sufficient to cater
for the failure of any one fuel cut-off valve without
endangering the fuel tank or surrounding structure.
Provision shall also be made to prevent the build-
up of excessive pressure in any pipeline.
3.5.11 The requirements of Clauses 3.5.8
to 3.5.10 inclusive apply also to the tanks of a
refuelling package installation.
3.5.12 It shall be possible to transfer fuel at
not less than half of the specification flow rate in
the event of failure of any main transfer pump or
equipment associated with the flight-refuelling
package.
FIRE PRECAUTIONS
3.5.13 The installation of associated
electrical equipment shall comply with the
requirements of Part 1, Section 4, Clauses 4.26
and 4.27 and Part 1, Section 6, Clause 6.6
3.5.14 Electrical connection (to discharge
static) shall be established between the Tanker
and Receiver before fuel is transferred.
3.5.15 There shall be no leakage from the
Tanker installation when the hose is stowed or
trailed. There shall be the minimum possible
spillage when the Receiver makes contact, no
leakage at all whilst in contact, and a minimum
leakage on disconnect. Leakage at disconnect
shall be held to this minimum both under normal or
emergency break-away conditions even with the
most adverse offset of the probe from the drogue.
3.5.16 Fuel pipes shall not run through: Such protection would involve the use of a The aim of this requirement is to minimise the risk
design that contains the fuel in the event of of fire or explosion caused by leakage from the

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REQUIREMENT COMPLIANCE GUIDANCE

a. passenger, leakage from the fuel pipes, however caused, flammable fluid system.
and subsequently vents the fuel into a suitable The drain system should also enable detection
b. crew, drain system. Typical examples of such a design and identification of the general location of any
are double-skinned/walled, jacketed or leakage. For further guidance see also EASA CS
c. cargo or shrouded. 25.863.
Whenever lagging is used in compartments in
d. baggage compartments which pipes, tanks or equipment containing
flammable fluids are installed, suitable
nor in hazardous proximity to potential sources of precautions shall be taken to prevent the wetting
ignition unless they are adequately protected of the lagging by flammable fluids as a result of
against potential sources of ignition and damage. normal operation, damage, failures of the
Any space between a pipe and its protection shall equipment or leakages from joints or unions
be adequately vented and drained. See also Part
1, Section 4.26, Clauses 4.26.32 and 4.26.82.
3.5.17 No fuel shall be left in the refuelling
pipelines after completion of the operation, if it
could constitute a fire hazard.
SAFETY CONSIDERATIONS
3.5.18 No single failure in the Receiver or
in the Tanker refuelling package or the installation
of it in the Aircraft shall endanger the safety of
either Aircraft, and no fuel/vapour shall be released
into the cockpit or cabin.
3.5.19 The equipment in both the Tanker
and Receiver shall be installed in such a position
that, with hose stowed, there is no appreciable
adverse aerodynamic effect on either Aircraft.
3.5.20 The accelerating capability of the
hose drum unit shall be such that the hose will not
"whip" when the Receiver makes contact at any
closing speed up to 2.44 m per second. The probe
shall be provided with a weak link so that the
nozzle will break away in the event of excessive
loads occurring due to instability of the hose or of
failure of the nozzle and coupling to release under

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REQUIREMENT COMPLIANCE GUIDANCE

normal operating conditions. It shall also be
provided with a non-return valve so that, in the
event of such breakage, fuel spillage will not
endanger the Receiver.
3.5.21 Consideration shall be given to
locating the refuelling probe so as to minimise the
risk of spilled fuel obscuring the windscreen,
entering the air intakes or otherwise interfering with
the safe and efficient functioning of the Aircraft or
its equipment.
3.5.22 The probe shall be so installed that .
the pilot of the Receiver has an adequate view, to
achieve accurate alignment and engagement with
the drogue. The nozzle shall normally be aligned
with the flight path of the Receiver at the refuelling
speed.
3.5.23 The system shall be installed so that
the centre of gravity of both Tanker and Receiver
does not move outside authorised limits before,
during and after transfer of fuel.
3.5.24 In the event of failure of the re-
winding gear, it shall be possible to jettison the
hose from the refuelling package. Otherwise it shall
be impossible for the hose to leave the Tanker. If
the equipment is of the "pod" type it shall be
possible, in combat Aircraft, to jettison the
complete assembly. To guard against inadvertent
jettisoning, the operation, if electrical, shall be by a
double pole swatch which is effectively guarded,
and if mechanical or hydraulic, the system shall be
duplicated.
3.5.25 When the refuelling package
incorporates a fuel tank, the fuel in the tank shall
be available for use by the Tanker if required.
3.5.26 Where fuel tanks are installed within

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REQUIREMENT COMPLIANCE GUIDANCE

the Aircraft pressure cabin, these tanks shall be
vented to atmosphere, and they shall be contained
in structural enclosures. Means of detecting leaks
into the enclosure shall be provided, and the
enclosure shall be drainable overboard.
3.5.27 A system safety assessment shall
be carried out demonstrating compliance with the
specification safety requirements.
3.5.28 Radio equipment and aerial
positions shall be capable of providing safe and
effective communications between the Tanker and
Receiver Aircraft, with the Receiver in the refuelling
position and in proximity to the Tanker.
INDICATORS AND LIGHTS
3.5.29 Means shall be provided in the For signal lights see Clauses 3.5.33 to 3.5.35
Tanker to indicate:

(a) the amount of fuel available for

transfer,

(b) when fuel is flowing and the rate of

flow,

(c) the amount of fuel (in kgs) which

has been transferred to the Receiver Aircraft,

(d) the length of the hose that is trailing

(for centre-line units) and,

(e) whether pressure has built up in the

hose such as to prevent successful contact
being established.
3.5.30 When night refuelling is specified, it
shall be possible to illuminate the probe and the
drogue and to floodlight the wings of the Tanker so

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REQUIREMENT COMPLIANCE GUIDANCE

as to provide a datum for the pilot of the Receiver.
3.5.31 The "root end" length of the hose,
which it is necessary to wind in after engagement
before fuel transfer can be initiated, shall be
marked so as to be clearly distinguishable to the
Receiver pilot.
3.5.32 On refuelling packages which
incorporate a fuel tank, an indicator, to show the
range full to empty shall be fitted in such a position
that it can readily be seen by ground personnel
during servicing. If the package is in the form of a
pod carried on the folding portion of the wing, the
indication shall be given whether the wings are
folded or spread.
SIGNAL LIGHTS
3.5.33 Signal lamp systems shall be
duplicated such that no single failure will result in
the inability of the Receiver Aircraft to monitor the
Tanker Aircraft signals.
3.5.34 A system of red, amber and green The lights shall operate in the following order:
signal lights shall be provided at the following
positions in the Tanker: (a) red, when the master switch is
switched on to start the refuelling
(a) on the control panel from which the operation,
refuelling operation is controlled, and
(b) amber, when the hose has
(b) externally in a position readily reached the full trail position,
visible to the pilot, or appropriate
crewmember, of the Receiver. (c) green, when successful contact
has been made and partial wind-in of the
hose under the thrust of the Receiver's
probe has opened the valves permitting
fuel flow,

(d) amber, when fuel flow ceases,

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REQUIREMENT COMPLIANCE GUIDANCE

and

(e) red, when winding in of the hose

has started.
3.5.35 It shall be possible, at any stage, to
flash the red lamps to warn the Receiver pilot to
disengage or stand-off.

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AIR-TO-AIR REFUELLING - FLIGHT

3.5.36 The object of the tests of this Clause is to demonstrate that the handling characteristics of Aircraft are satisfactory.

(a) When operating in the air-to-air tanker role.

(b) When operating in the air-to-air receiver role.

3.5.37 The Clause is divided into two parts:

(a) Part A describes the tests to be made to assess the flying qualities of Aircraft when operating as Air-to-Air Refuelling (AAR) tankers.

(b) Part B describes the tests to be made to assess the flying qualities of Aircraft when engaged in AAR operations as receivers.

REQUIREMENT COMPLIANCE GUIDANCE

TANKERS
3.5.38 The installation of AAR equipment in
the tanker Aircraft will normally either take the form
of a pod for external attachment to a wing pylon, or
a hose Drum Unit (HDU) installation in the rear
fuselage, or be a combination of such systems in a
multi-point tanker.
3.5.39 When an Aircraft is developed as a The tests shall normally include the following: The test methods to be used are described in Part
tanker, or converted to the tanker role, a full 1, Section 2, Clauses 2.2 to 2.17 (as appropriate).
programme of tests will be carried out to assess the (a) Stability and control of the Aircraft Some further tests, specific to the tanker role, will
handling characteristics over its operating with the hose trailed, and when the hose is be required, and this part deals with these tests.
envelope. being extended or retracted. Stability of
the hose and drogue when extended. All
these tests shall be made with the Aircraft
under manual and, where applicable,
automatic control.

(b) Night lighting of the Aircraft and

AAR equipment and line-up markings.

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REQUIREMENT COMPLIANCE GUIDANCE

(c) Failure cases: e.g., handling at
low speed with the hose extended.

(d) Cockpit layout: AAR controls and

indicators.
3.5.40 Any testing relating to the 'hose'
shall include where relevant configuration of any
one, any combination or all of the hoses on a multi-
point tanker.
RECEIVERS
3.5.41 Tests of receiver Aircraft shall
consist of assessments of stability and control
when approaching to make contact, in contact, and
withdrawing from contact with the tanker. Failure
cases, such as the effect of engine failures shall be
considered. Night lighting shall also be assessed.

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PART A - TANKERS

APPLICABILITY

3.5.42 These tests are applicable to all classes of Aircraft (when operating in the AAR tanker role) as defined by Part 1, Section 2, Clauses
2.1.13 to 2.1.16, and to all types of control system as defined by Part 1, Section 2, Leaflet 6, Para 2.

REQUIREMENT COMPLIANCE GUIDANCE

EQUIPMENT
COCKPIT INSTRUMENTS
3.5.44 The test Aircraft shall be fitted with a
normal accelerometer and angle of attack gauge.
The Airspeed Indicator shall have been recently
calibrated.
TEST INSTRUMENTATION
3.5.45 The parameters which should be
recorded during these tests are listed in Leaflet 9
Para 9.
LOADING
3.5.46 The tests are to be made at loadings
such that all forward and aft c of g load conditions
of Table 9 are covered.
GENERAL TEST CONDITIONS
ALTITUDE
3.5.47 The tests shall be made at low,
medium and high altitudes appropriate to the flight
phase of the Aircraft under test.
AIRCRAFT CONFIGURATIONS AND SPEED RANGES REQUIRED FOR THE TESTS
3.5.48 The tests shall be made over the
speed ranges indicated in Table 9, and with the lift
devices (i.e., slats, flaps etc.,) in the position(s)
most appropriate to the flight phase categories
(Part 1, Section 2, Clauses 2.1.17 and 2.1.18)
given in Table 9. Any other intermediate positions

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shall also be considered if these are likely to be
used. Combinations of flaps down/slats in, or vice
versa shall only be considered if these are likely to
be used, or if they result from single failures.
TESTS
PRELIMINARY TESTS
3.5.49 Some preliminary tests can be
carried out on the ground to assess the suitability
of the controls and indicators relating to the AAR
tanker role at the pilots' stations and the refuelling
control panel. An initial assessment of the night
lighting, signal lighting system and visual
references may also be made on the ground before
flight trials are undertaken.
HOSE AND DROGUE STABILITY TESTS
3.5.50 The tests of Clauses 3.5.51 to
3.5.55 shall be made with the hose both empty and
full, over the specified maximum AAR envelope.
3.5.51 Straight and Level Flight - For these The chase or receiver Aircraft is to be positioned
trials a chase or receiver Aircraft equipped with a slightly behind the drogue, a few feet from the
video or cine camera shall be provided. refuelling position. At each test speed and
altitude, with the tanker under manual control,
the behaviour of the hose and drogue shall be
recorded, and any vertical or lateral oscillations
which occur shall be noted. The tests shall be
repeated with the tanker under autopilot control.
In some instances, particularly with a large
receiver Aircraft, the proximity of the nose of the
receiver may cause a disturbance in the airflow
round the drogue, with consequent irregular or
oscillatory hose behaviour. If this tendency is
noted, a separate chase Aircraft shall be
employed to record, using video or cine camera,
the hose behaviour during AAR approaches and
contacts.

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3.5.52 Banked Turns - Gentle turns of up to
30° bank are to be initiated with the tanker under
(a) manual and (b) autopilot control, and the
behaviour of the hose and drogue assembly shall
be recorded at the initiation of, during and on
recovery from turns.
3.5.53 Descents - The behaviour of the
hose and drogue shall be recorded on initiation of,
during and on recovery from descents made with
up to 500 ft/min rate of descents.
3.5.54 Longitudinal Trimming - The effect
of small adjustments to the longitudinal trim of the
Aircraft under manual and autopilot control shall be
assessed by recording the hose and drogue
behaviour during and after the trim changes. Any
long period phugoid tendencies of the Aircraft shall
be noted and recorded.
3.5.55 Induced Hose Oscillations -
Oscillations in pitch shall be initiated by inducing
'doublet' inputs in the pitch control. The behaviour
of the hose and drogue shall be recorded.
NIGHT LIGHTING, VISUAL REFERENCES AND LINE-UP MARKINGS
3.5.56 Chase or receiver Aircraft shall be The suitability of the lighting and the ability to
used in making assessments of the night lighting vary its intensity to cover a range of light
and line-up markings of the tanker. conditions shall be assessed. The presence of
any unwanted reflections in the receivers'
canopies shall be noted. The adequacy and
position of visual references and suitability of the
line-up and hose markings shall be evaluated
over a range of day and night ambient
conditions. The suitability of the signal lights
shall be investigated.
FAILURE CASES
3.5.57 Hose Jettison Tests - These tests Video or cine films shall be taken to provide a If failure to re-wind the hose can result from a
shall be made in association with Engineering record of the trajectory of the hose, following single malfunction, hose jettison tests may be

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Division and shall be undertaken over a specified jettison. Wind-tunnel tests or predictions shall necessary.
dropping zone with a chase Aircraft in attendance. precede the tests to provide guidance on the
preferred jettison speed, and the initial flight
tests shall be made suitably close to this speed.
3.5.58 AAR Operations With Engine Failed The most critical engine shall be shut down and
- Tests shall be made to assess the feasibility of the hose and drogue deployed at an altitude at
undertaking AAR operations with an engine of the which a satisfactory cruising speed (within the
tanker failed, and to explore the flight envelope in normal AAR limits) can be maintained. A
which fuel transfer is possible under these receiver shall then attempt to make a dry
conditions. contact. If contact can be made and held, fuel is
to be transferred. A practical AAR envelope in
terms of speed and altitude for this condition can
then be explored, and the tests are to include
gentle turns of up to 30° bank and descents of
up to 500 feet/min rate of descent whilst in
contact. The refuelling tests shall be made with
the tanker under manual and autopilot control.
COCKPIT LAYOUT, CONTROLS AND INDICATORS
3.5.59 All controls and indicators relating to Accurate knowledge of the fuel contents and
the AAR tanker role at the pilots' stations and the distribution is essential to enable the tanker's
refuelling control panel shall be evaluated, and any mass and c of g position to be readily determined.
shortcomings or deficiencies noted.

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APPLICABILITY

3.5.60 These tests are applicable to all classes of Aircraft (when operating in the AAR receiver role) as defined by Part 1, Section 2, Clauses
2.1.13 to 2.1.16, and to all types of control system as defined by Part 1, Section 2, Leaflet 6, Para 2.

REQUIREMENT COMPLIANCE GUIDANCE

EQUIPMENT
COCKPIT INSTRUMENTS
3.5.61 The Airspeed Indicator shall have
been recently calibrated.
TEST INSTRUMENTATION
3.5.62 The parameter(s) which should be
recorded during these tests are listed in Leaflet 9
Para 9.
3.5.63 In addition to the handling See Part 1, Section 1, Clause 1.2.9 Guidance.
instrumentation specified in Clause 3.5.62, some
structural instrumentation may also be required to
monitor fatigue aspects.
LOADING
3.5.64 The tests shall be made at loadings If the Aircraft is capable of carrying external
such that all forward and aft c of g load conditions stores, these loadings shall include the most
of Table 9 are covered. adverse combinations of the stores in relation
to:

(a) highest mass.

(b) lowest lateral and directional

stability.

(c) highest pitch inertia.

(d) The most aerodynamically

destabilising configuration.

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GENERAL TEST CONDITIONS
ALTITUDE
3.5.65 The tests shall be made at selected
altitudes over the intended AAR envelope, up to the
maximum practical AAR altitude.
AIRCRAFT CONFIGURATIONS AND SPEED RANGES REQUIRED FOR THE TESTS
3.5.66 The tests shall be made over the
speed ranges indicated in Table 10, and with the
high lift devices (i.e., slats, flaps etc.,) in the
position(s) most appropriate to the flight phase
categories (Part 1, Section 2, Clauses 2.1.17 and
18) given in Table 10. Any other intermediate
positions shall also be considered if these are likely
to be used. Combinations of flaps down/slats in, or
vice versa shall only be considered if these are
likely to be used, or if they result from single
failures. High lift devices will normally only be used
if essential for the tanker speed range to be
matched.
3.5.67 For Aircraft which possess variable
wingsweep, the sweep angle to be used is that
most appropriate to the flight phase category, but at
least one other wingsweep angle shall be
evaluated as well.
TESTS
GENERAL
3.5.68 If any significant changes have been e.g., engine handling tests with an AAR probe
made to the Aircraft in converting it to the air-to-air extended, if in front of an engine intake.
Receiver role, a brief assessment shall be made to
ensure that the Aircraft and/or engine handling
characteristics have not been significantly changed
by the conversion.
3.5.69 Except where otherwise stated, the
tests of Clauses 3.5.66 to 3.5.68 shall be made
over the specified AAR envelope.

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PREPARATION
3.5.70 Some preliminary tests can be
carried out on the ground to assess the suitability
of the controls and indicators relating to the AAR
receiver role at the pilots' stations.
INITIAL FLIGHT TESTS
3.5.71 The flight programme shall (a) After making contact with the drogue, the
commence with an initial exploration of the receiver optimum refuelling position shall be achieved,
behaviour when flying in representative positions and the flying qualities assessed.
behind the tanker with the tanker hose wound in.
This should be done at moderate speed, altitude (b) The receiver shall be displaced from the
and weight, where the receiver will have plenty of optimum position (with the hose still attached).
power and manoeuvre capability in hand. In The handling shall be explored within a 'cone' of
addition to exploring the positions likely to be about 15° (if possible) around the optimum
encountered during a good approach and contact, position.
the receiver pilot shall also ensure that the handling
remains acceptable when displaced from the (c) The receiver shall be moved closer to the
correct position. The tanker will then extend the tanker, and the handling assessed at the
hose and the receiver will make the first contacts closest position that is likely to be encountered.
with the drogue. At each speed and altitude the
assessments listed under Compliance shall be (d) In the optimum refuelling position left and
made. right hand turns shall be made by the tanker
and receiver. Bank angles up to about 30° shall
be assessed if possible.

(e) The tanker shall be followed into

descents at rates of up to 500 ft/min to cover the
"toboggan" technique.

(f) Contact shall be broken in the normal

way by gently backing off from the tanker until,
at full extension, the hose disconnects.

(g) Contact shall be broken by emergency

breaks. The use of airbrakes shall be assessed

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during this manoeuvre.

(h) When confidence in the handling aspects

has been gained, further contacts shall be made
and fuel transferred.
EXTENSION AND EXPLORATION OF THE AAR ENVELOPE
3.5.72

(a) Handling assessments shall be (a) With both Aircraft at light weight. Normally the objective of the flight trials
made at the maximum practical speeds, programme is to clear as wide an AAR envelope
Mach No’s and altitudes. (b) With both Aircraft at high weight. for the tanker and receiver combination as
possible.
(b) If any handling problems arise as a (c) With the receiver at the extremes of its c
result of unusual behaviour of the of g envelope.
hose/drogue when making or holding contact,
the tests shall be repeated with video or cine
camera coverage from a chase Aircraft.
3.5.73 In addition to the general handling
considerations covered in Clauses 3.5.71 and
3.5.72 above, the following particular aspects shall
be investigated:

(a) Airframe Buffet. (a) Buffet levels shall be assessed with the
receiver at the normal refuelling position and at
high, low, left and right positions within the 15°
cone, and any marked variation in buffet level
shall be noted.

(b) Engine Handling. (b) AAR operations can demand large

engine power variations to make and then
maintain contact with the tanker. Engine
behaviour in response to normal throttle
variations within the 15° cone shall be noted,
and any tendency of the engine(s) to surge shall
be recorded.

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At the lower altitudes and speeds, and where

appropriate, the use of airbrakes to increase the
drag, so that the engine(s) can be operated at a
higher power setting, with faster response
characteristics, shall be investigated.

(c) Lateral Control Power. (c) The ability to trim out or counteract by
use of lateral controls any tendency to roll
towards the tanker's fuselage when approaching
wing stations, shall be confirmed.

(d) Lateral and Directional Behaviour. (d) The full AAR envelope shall be explored
to ensure that no unacceptable lateral or
directional oscillations occur, and that lateral
and directional control can be maintained at all
flight conditions.

(e) Longitudinal Characteristics. (e) The longitudinal handling characteristics See Part 1, Section 2, Clause 2.21.29
shall be assessed over the full intended AAR
envelope, to confirm that the longitudinal control
is satisfactory, no unacceptable changes of trim
in pitch are experienced, re-trimming can be
effected without difficulty and no excessive short
period pitching oscillations occur.
NIGHT LIGHTING
3.5.74 If any features of the tanker lighting
are unsuitable for the particular receiver Aircraft,
this shall be noted. Also the lighting of the
receiver's cockpit shall be assessed whilst making
and holding contact to ensure that there are no
undesirable reflections or unsuitable lighting
conditions.
FAILURE CASES
3.5.75 Asymmetric Thrust. (a) No significant handling problems are

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generally experienced with asymmetric thrust
following engine failure of a receiver. However,
the loss of thrust will result in a marked
reduction of the available AAR envelope, and
this shall be defined.

(b) For Aircraft in which Part Throttle Reheat

(PTR) is not available, it may be necessary to
operate with fixed reheat thrust following engine
failure, and in this case it may be necessary to
vary the drag by airbrake operation to maintain
station. In these cases tests shall be undertaken
to ensure that adequate control can be
maintained.
3.5.76 Failure of Autostabilisation. Where appropriate, tests are required to assess
the feasibility of making AAR contacts when
either the tanker or receiver has sustained a
failure prior to making contact. The malfunction
may be either a full or partial autostabiliser
failure, depending on the type of Aircraft
involved. Where necessary, a more restrictive
AAR envelope in the presence of such failures
shall be recommended. Where appropriate, any
related effects of partial loss of control power on
the tanker or receiver shall also be considered.
FUEL SPILLAGE
3.5.77

(a) If any fuel spillage occurs on

disconnection from the hose, it is necessary
to confirm that sufficient visibility can be
retained to enable the receiver to break off
safely, and that no fuel ingestion problems
arise.

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(b) If any handling problems arise as a
result of unusual behaviour of the
hose/drogue when making or holding contact,
the tests shall be repeated with video or cine
camera coverage from a chase Aircraft.
COMBINATIONS OF FAILURES OF THE TANKER AND RECEIVER
3.5.78 If failure studies show that in the
case of a particular combination of tanker and
receiver, there is a high probability of a failure (or
failures) being present on the tanker, and an
unrelated failure (or failures) being present on the
receiver, the effect of these failure combinations
shall be investigated with the aim of permitting the
operational use of some of these, and identifying
associated limitations or techniques.
FLIGHT TRIALS
AIR-TO-AIR REFUELLING
3.5.79

(a) Aircraft which have a capability of

receiving fuel during air-to-air refuelling shall
be tested to demonstrate:

(1) Compatibility with in-service UK and

allied tanker aircraft.

(2) That the Aircraft C of G is not

moved significantly due to the uneven
distribution of fuel entering the tanks.

(3) That the surge pressures resulting

from refuelling valve closure are within the
design limits.

(4) That the venting system is adequate

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to prevent an excessive build-up of
pressure in the fuel tanks.

(5) That on separation of the probe

from the drogue the fuel spilt does not
either cause a significant deterioration in
the pilot's view by obscuring the
windscreens or enter the engine intakes in
sufficient quantities to cause the engine(s)
to surge or enter any other parts of the
Aircraft where loading fuel could become
a hazard.

(b) Aircraft which have the capability to

dispense fuel during air-to-air refuelling shall
be tested to demonstrate:

(1) Compatibility with in-service UK and

allied receiver Aircraft.

(2) That the Aircraft C of G is not

moved significantly due to the uneven
distribution of the fuel dispensed from the
individual Aircraft tanks, or that suitable
operating procedures will prevent
unacceptable movement of the Aircraft C
of G.

(3) That the surge pressure which may

result from an emergency receiver
disconnect, receiver refuelling valves
cycling or more than one receiver Aircraft
disengaging at the same time are within
design limits.

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(4) That the venting system is adequate
to prevent an excessive negative pressure
differential across the walls of the fuel
tanks.

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3.6 ARRESTING HOOKS FOR LAND-BASED AEROPLANES

INTRODUCTION

3.6.1 This Clause states requirements which are applicable when an arresting hook is specified. Leaflet 9 gives background and supplementary
information relevant to these requirements.

OBJECT OF TESTS (Clauses 3.6.31 to 3.6.38)

3.6.2 Clauses 3.6.31 to 3.6.38 describe the tests that are required to determine that the aeroplane can successfully and safely engage airfield
hook wire arresting systems and airfield arrestor barrier systems.

3.6.3 The object of the tests of this Clause is to provide information which can be used as a basis for limitations, within which the aeroplane can
be used in Service.

RELEVANT DESIGN REQUIREMENTS

3.6.4 Def Stan 00-970, Part 1, Section 3, Clause 3.10 and Part 1, Section 4, Clause 4.13

APPLICABILITY

3.6.5 The tests described in this Clause are applicable to all new aeroplanes, aeroplane-mounted arresting hooks, and airfield-mounted hook
wire arresting systems and arrestor barriers. The tests are also applicable when modifications have been made likely to affect the results of the
tests unless otherwise stated.

3.6.6 The tests must be conducted on aeroplanes and arresting systems which are fully representative of the Service standard. Consideration
should be given to testing all relevant combinations of aeroplane configuration, mass, speed and off-centre distance (up to 20% of the total span of
the arresting gear).

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GENERAL DESIGN REQUIREMENTS
3.6.7 Where the arresting hook is required The Aeroplane Specification will state whether the
to be used with a number of different arresting arresting hook is to be designed for use with a
gears: particular arresting gear or a number of arresting
gears.
(a) the shape of the hook shall be
optimised for the arresting gear having the
largest diameter hook cable,

(b) the strength of the installation shall

be determined by the arresting gear giving
the greatest hookload.
3.6.8 The length of the arresting hook Note: All these cases will normally be met if the
suspension arm shall be adequate to provide for hook suspension arm is long enough to provide
engagements at all attitudes which could arise for the case where the nosewheel tyre and oleo
during an arrest: are fully compressed and the mainwheel tyres
and oleos are fully extended with the mainwheel
(a) at any time during normal take-off, just touching the runway.

(b) during heavy braking following a

decision to abort a take-off,

(c) at any time during normal landing,

(d) following a heavy landing with

maximum nose-down pitch.
3.6.9 The angle of the axis of the hook
suspension arm to the ground (the trial angle) shall
not exceed 80° in the most adverse attitude during
landing or taxying.
3.6.10 Vertical damping or other means
shall be provided to prevent the arresting hook
skipping over the arresting cable. The following
cases shall be investigated:

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(a) landing at maximum speed at

maximum pitch and roll angles,

(b) taxying at any speed up to the

maximum specified and at any possible pitch
angle over a 12.5 mm (0.49 in) step-up
(chamfered at 45°) or 12.5 mm (0.49 in) (90°)
step-down at a point of 15 m (16.4 yards)
from the arresting wire.
3.6.11 Means shall be provided to prevent
lateral instability of the arresting hook suspension
systems when landing at maximum vertical velocity
with maximum tail-down rate of pitch at an
associated high angle of bank.
3.6.12 It shall be possible to lower the hook
in not more than 5 seconds in flight at all speeds
up to the Maximum Speed or Flight-Undercarriage
Down and in not more than 2 seconds when
taxying at all speeds up to Maximum Design Take-
Off Speed.
3.6.13 With the hook in its stowed position
all attachments shall have proof and ultimate
factors of not less than 1.5 and 2.0 on the
maximum loads arising from the combination of
any flight envelope load with any pre-stressing load
which may be present in the suspension system.
DETAIL DESIGN REQUIREMENTS
RELEASE SYSTEM
3.6.14 The hook release control shall be The hook release control shall be in accordance
operated by the pilot. The pilot shall also be with the requirements of Part 1, Section 4, Clause
provided with an indicator to show that the hook is 4.19
down. Indication of proper engagement of the
uplock shall be provided to the ground crew.
3.6.15 The down stop shall have proof and

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ultimate factors of not less than 1.5 and 2.0 on the
maximum dynamic down load arising from release
of the hook with the aeroplane jacked up to allow
full movement of the hook suspension system.
MAINTENANCE REQUIREMENTS
3.6.16 Unless an internal system is
provided for the purpose it should be possible for
two men to raise and stow the hook manually
without the use of special tools.
3.6.17 Those parts of the hook which
contact the runway at any time shall be designed
for easy replacement.
3.6.18 A ground safety lock with
appropriate flag shall be provided on the hook or
suspension arm. It shall be capable of preventing
inadvertent lowering of the hook from the stowed
position. It should not be possible for the safety
lock to damage the aeroplane in the event of
inadvertent operation of the arresting hook system.
ENVIRONMENTAL EFFECTS
3.6.19 The environment in which the (a) The temperature environment in which
arresting hook and its associated control system the hook and its suspension is normally stowed
are required to operate shall be considered. shall be assessed and appropriate material
strength allowances shall be included in all
strength calculations.

(b) The location of components of the hook

operating and control system shall ensure
proper functioning of these components in the
environment expected.
DESIGN HOOK LOADS
3.6.20 The principal operational failure These will consist of:
cases for which the arresting gear is expected to
be used will be stated in the Aeroplane (a) Failure cases of normal severity -
Specification. those arising from failures under

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conditions where it can reasonably be
expected that the aeroplane mass and
speed will be within normal limits. No more
than minor superficial damage to the
aeroplane will be acceptable even if the
entry is at maximum off-centre distance
which is to be taken as 12 m (40 ft) unless
otherwise stated in the Aeroplane
Specification.

(b) Failure cases of exceptional

severity - those high speed or high mass
cases where some structural damage to
removable parts of the arresting hook and
its suspension system is acceptable and
more severe superficial damage to the
aeroplane is expected.
3.6.21 The Normal Axial Design Hookload See Clause 3.6.23 for assessment of Impact
will be the greatest hookload arising in the failure Hookloads.
cases of normal severity (Clause 3.6.20
Compliance (a)) and the Maximum Axial Design
Hookload will be the greatest hookload arising in
the failure cases of exceptional severity (Clause
3.6.20 Compliance (b)), in the dynamic or steady
braking phases of the arrest.
3.6.22 Axial Design Hookloads shall be Where trials to determine the performance of the
determined according to the best available data for arresting gears specified have been
the arresting gear or gears specified and for each satisfactorily completed, and measured axial
phase of the arrest. Where it is necessary to hookload curves are available for any of the
estimate the hookload theoretically in any phase phases of an arrest, these should be used in lieu
the methods and factors given in Leaflet 8 shall be of estimated data.
used.
3.6.23 Vertical and lateral components of
the Design Hookloads shall be determined by
applying:

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(a) the Normal or Maximum Axial

Design Hookload (as appropriate) at any
angle within a cone of 10° whose axis is the
intersection of the longitudinal vertical plane
through the aeroplane c of g and the plane
containing both the hook pivot point and the
runway edge sheaves and whose apex is at
the hook pivot point,

(b) 60% of the Normal or Maximum

Axial Design Hookloads (as appropriate) at
any angle within a cone of 20° with the same
axis and apex.
3.6.24 Impact Hookloads shall be applied Consideration shall be given to the problems
at the most adverse aeroplane pitch angle and arising from a stress wave front meeting a
suspension arm angle which could arise at impact discontinuity in the structure of the hook,
in any of the failure cases of Clause 3.6.20 suspension, and back-up structure.
3.6.25 The suspension arm and its
attachments (including the vertical and lateral
damping systems) shall be designed to withstand
the effects of vibration caused by impact with the
runway and the angular acceleration immediately
following impact in all cases.
3.6.26 In the case of roll-back with the hook
down and jammed against an obstruction the hook
suspension arm shall fail as a strut or break out
before the aeroplane structure or any non-
removable part sustains permanent deformation.
The rear face of the heel of the hook shall be
designed to minimise the possibility of jamming
during roll-back against a 90° step of 12.5 mm (½
in) depth.
STATIC STRENGTH
3.6.27 The complete aeroplane structure

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and all removable parts shall have proof and
ultimate factors of not less than 1.125 and 1.5
respectively under the application of the Maximum
Axial Design Hookload and the Maximum Impact
Hookload.
3.6.28 The hook beak and associated parts
which are designated as replaceable after a
specified number of arrests (or when bent or worn
out) shall have proof and ultimate factors of not
less than 1.125 and 1.5 respectively under the
Normal Axial Design Hookload and the Normal
Impact Hookload and an ultimate factor of not less
than 1.0 under the Maximum Axial Design
Hookload and the Maximum Impact Hookload.
FATIGUE LIFE
3.6.29 The fatigue life of all components of
the system shall be estimated for the loading
conditions given in the Aeroplane Specification or
subsequently agreed with AD AS DMPS. The fully
factored life of the aeroplane structure and of all
non-replaceable components shall be not less than
the specified life of the aeroplane as a whole. The
fully factored life of replaceable components shall
be at least equal to their specified replacement life.
Consideration shall be given to the effects of
stress-corrosion. The necessity for a fatigue test
shall be discussed with the Project Team Leader
(PTL)
NOSEWHEEL LOADS
3.6.30 Unless the Aeroplane Specification
states otherwise the nose unit and complete
airframe shall have a factor of not less than 1.0 on
the loads arising from the combination of Maximum
Axial Design Hookload with either:

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(a) nosewheel slam arising from an in-
flight engagement in all practical
combinations of vertical and horizontal
velocities from zero to maximum; or

(b) heavy mainwheel braking.

TEST REQUIREMENTS
STATIC TESTS
3.6.31 The necessity for static strength
tests shall be considered.
DEVELOPMENT TESTS
3.6.32 A programme of taxying tests and
flight tests designed to explore as fully as possible
all the factors discussed in Leaflet 10 shall be
agreed with the Project Team Leader (PTL)
3.6.33 A test aeroplane shall be fitted with
adequate instrumentation to permit measurement
of axial and component hookloads in each of the
phases of arrest.
3.6.34 Within speed limitations which
provide an adequate margin of safety for the tests,
at off-centre distances up to 20% of the total span
of the arresting gear, sufficient data shall be
obtained to show by extrapolation that the
specified requirements have been met.
EQUIPMENT FOR TESTS
INSTRUMENTATION
3.6.35 The test aeroplane shall be fitted
with adequate instrumentation to permit
measurements of the axial and component hook
loads, alighting gear loads, airframe structure
loads, etc. The instrumentation parameters shall
be agreed with the Project Team Leader (PTL)
VISUAL MONITORING EQUIPMENT

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3.6.36 Visual monitoring equipment shall
be mounted on the aeroplane and on the ground,
viewing the arrestor system so that the dynamic
behaviour of the arrestor system can be monitored.
TEST DETAILS
3.6.37 The items listed under compliance (a) The evidence from static strength tests Full details of the tests required will be issued
shall be considered prior to test: and preliminary ground rig tests, later. Pending the issue of the complete Clause,
further details may be obtained from the Project
(b) The standard of the arrestor system and Team Leader (PTL).
its suitability for the test proposal,

(c) The operational use of the arrestor

system.
3.6.38 The tests to be carried out should (a) Hook wire system Trampling tests along centreline and offset from
normally include those listed under compliance. the runway centreline,
Hook/wire engagement tests along centreline and
offset from centreline over relevant speed range.

(b) Barrier System Tow-in tests along centreline and offset from
centreline.

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3.7 INSTALLATIONS FOR EMERGENCY RECOVERY FROM STALL AND SPIN

INTRODUCTION

3.7.1 The requirements of this Clause are applicable to installations designed to assist recovery after loss of control in stalling and spinning flight
trials and, in investigations of aeroplane behaviour at high angles of attack.

3.7.2 The requirements of Clauses 3.7.11 to 3.7.47 below are based on the assumption that the installation is required to function at the relatively
low airspeeds experienced in a stall or in an erect or an inverted spin.

3.7.3 Where it is necessary to consider other types of uncontrolled motion, (e.g., autorotation in roll at high speed) the design requirements shall
be agreed with the Project Team Leader (PTL).

REQUIREMENT COMPLIANCE GUIDANCE

BASIC REQUIREMENTS
3.7.4 In accordance with the requirements See Leaflet 11.
of this Clause, a parachute or rocket installation
shall be fitted to all aeroplanes on which stalling
and spinning flight trials and investigations of
aeroplane behaviour at high angles of attack are to
be made, unless confidence is established before
or during the trials that its omission would not give
rise to unacceptable hazards. The fitment of a
parachute or rocket installation shall also be
considered for any trial where there is a possibility
of the loss of control.
3.7.5 Equipment to permit transmission of
information relevant to the flight trial to a
monitoring ground station shall be provided, unless
otherwise agreed with the Project Team Leader
(PTL), with whom the data to be transmitted shall
be agreed. The aeroplane communications
equipment shall permit direct two-way
communication between the ground monitoring
station and the aeroplane.
3.7.6 The suitability for stalling and

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spinning trials of the information provided for the
pilot by the normal instrument displays shall be
assessed and additional instrumentation shall be
provided where necessary.
3.7.7 Means shall be provided as
necessary to offset any expected adverse effect on
recovery that can result from system malfunction
arising from flight conditions experienced during
stalling and spinning trials.
DESIGN REQUIREMENTS FOR PARACHUTE AND ROCKET INSTALLATIONS
3.7.8 The installation shall be designed to If so required by the Project Team Leader (PTL), See also Leaflets 15 and 16
provide sufficient forces and moments to ensure flight tests shall be made to demonstrate
recovery from the most adverse flight conditions compliance
anticipated.
3.7.9 The installation shall be designed to
have as little effect as possible on the normal
contours of the airframe and on the mass
distribution of the aeroplane.
3.7.10 The installation shall be provided
with means of protection against the,
environmental conditions likely to be encountered
during the period of the flight trials, both on the
ground and in the air.
PARACHUTE INSTALLATIONS
INSTALLATION
3.7.11 The installation, including the control See also Leaflet 12.
and operating system, shall be designed so that as
far as practicable no single fault or likely
combinations of faults shall result in failure to
operate when required, or in inadvertent operation.
3.7.12 The installation shall incorporate a
design feature such that if for any reason the
parachute leaves the aeroplane before the pilot's
control is operated, there is no possibility of the
parachute or cable remaining attached to the

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aeroplane.
3.7.13 If any explosive device is
incorporated, the system shall conform with RAE
Specification WE659 and the need for other safety
precautions shall be agreed with the Project Team
Leader (PTL)
CONTROL
3.7.14 The control for streaming shall be
combined with the control for jettisoning in such a
manner, that jettison cannot be selected before the
parachute has been streamed.
3.7.15 The control shall be designed to
obviate inadvertent operation of either streaming or
jettisoning.
3.7.16 The control shall be readily
accessible to the pilot's throttle hand with harness
locked.
3.7.17 If a second pilot is to be carried
during the trials, a duplicate control of like design
shall be provided at the second pilot's station
unless he has ready access, with harness locked,
to the first pilot's control.
STREAMING
3.7.18 The design aim should be for the The design feature which ensures compliance
parachute to be fully inflated within 3 seconds of with this requirement shall not depend solely on
the operation of the pilot's control. aerodynamic forces.
3.7.19 The stability of the Parachute when
streamed in a smooth airflow shall be such that the
cable attaching it to the aeroplane does not
oscillate outside a cone whose semi-angle is 5°.
3.7.20 The design of the installation shall
include provisions to ensure that fouling any part of
the aeroplane during or after streaming, to the
extent that recovery is jeopardised, is avoided.

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3.7.21 The design of the installation shall
ensure that risk of the streamed parachute and
cable obstructing the emergency escape path of
any crewmember shall be as small as is
reasonably practicable.
JETTISON
3.7.22 It shall be possible to jettison the
parachute and cable after the parachute has been
streamed, whether or not the parachute has
inflated correctly, under loads up to the ultimate
design load of the parachute or the operating load
of the automatic jettison device (see Clause
3.7.24) whichever is the less, and under the most
adverse combination of load and angle that can
occur in the stall or spin.
3.7.23 It shall be possible to make a
before-flight check that the jettison mechanism is in
the operational position.
3.7.24 An automatic jettisoning device, or
weak link, which functions independently of the
pilot, shall be provided to protect the aeroplane
structure from excessive parachute loads. This
device shall be designed so that the parachute and
cable are jettisoned automatically at a load
corresponding to the proof strength of the structure
to which the parachute cable is attached, within the
limits +0%, - 10%.
STRENGTH OF THE PARACHUTE
3.7.25 The parachute canopy, rigging lines
and cables shall have an ultimate factor of not less
than 1.5 on the maximum loads sustained during
towing under the conditions of Clause 3.7.26
below.
3.7.26 The parachute should be capable of The speed Vp is defined as the maximum
being streamed and inflated at a speed of Vp at equivalent airspeed along the flight path liable to

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any altitude at which it is to be used. The streamed occur in either the stall or spin condition, up to the
parachute shall be capable of being towed at a point where recovery can safely and reliably be
speed of 1.3 Vp at any altitude at which it will be completed by normal use of flying controls.
used for a period of 30 seconds without suffering
significant damage.
STRENGTH OF THE AEROPLANE STRUCTURE
3.7.27 The aeroplane structure shall have
proof and ultimate factors not less than 1.125 and
1.5 respectively on the maximum load derived from
the aerodynamic and inertia forces, and the forces
imposed by the opening of the parachute at the
most adverse combination of speed and angle
when the parachute is streamed in a stall or a spin.
3.7.28 The aeroplane structure shall also See Clause 3.7.26
have proof and ultimate factors not less than 1.0
and 1.2 respectively on the maximum load derived
from the aerodynamic and inertia forces arising
during recovery from the stall or the spin, and the
steady force imposed by the streamed parachute
at a speed equal to 1.3 Vp, and at the most
adverse cable angle.
GROUND TESTS
3.7.29 Before the light tests of Clause
3.7.30 are made, ground tests of each installation
shall be carried out to demonstrate that:

(a) The part of the installation which

initiates the streaming of the parachute
operates satisfactorily and ensures extraction
or ejection to the desired distance.

(b) The parachute assembly will not

jettison inadvertently during steaming.

(c) The parachute assembly will jettison

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satisfactorily under loads up to the ultimate
design load of the parachute or the operating
load of the automatic jettison device,
whichever is the less, and under the most
adverse combination of load and angle that
can occur in the stall or spin.

(d) The parachute assembly will jettison

under the low loads it will experience if the
parachute has failed or if it has not inflated
correctly.

(e) The automatic jettison device

operates within the design load range (see
Part 1, Section 4 Clause 4.4.3)

(f) The functioning of the streaming

and jettisoning assembly is not affected by
the operation of any electrical or electronic
systems likely to be used during the flight
trials.
FLIGHT TESTS
3.7.30 Before stalling or spinning flight Note: If an aeroplane is subsequently altered so
trials begin: that the associated structure and/or the parachute
installation are changed, then the tests described
(a) The Contractor shall undertake such in (b) shall be repeated.
towing trials as are necessary to satisfy the
Project Team Leader (PTL) that the
parachute meets the requirement of Clause
3.7.26.

(b) A parachute shall be satisfactorily

streamed and afterwards jettisoned in flight
from the first aeroplane to be used in the
trials, and from as many individual

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aeroplanes of the same type as deemed
necessary by the Project Team Leader
(PTL); the test shall be done at a speed of
1.05 Vp, with low engine power settings, and
at the altitude at which the parachute would
normally be streamed for the initiation of
recovery from stalling and spinning trials.
3.7.31 If so required by the Project Team
Leader (PTL), flight tests shall be made to
demonstrate compliance with Clause 3.7.8
ROCKET INSTALLATIONS
INSTALLATION
3.7.32 The installation, including the control See Leaflet 13.
and operating system, shall be designed so that as
far as practicable, no single fault or likely
combination of faults shall result in failure to
operate when required, or in unwanted operation.
3.7.33 A means of checking the
serviceability of the rocket initiating system before
and during flight shall be provided.
3.7.34 The characteristics of the rocket
installation, in particular the magnitude and
duration of thrust, shall be agreed with the Project
Team Leader (PTL)
3.7.35 The rocket installation shall be
designed so that, as far as possible, there shall be
no damage to the aeroplane caused by heat,
debris or other effects of the rocket burning.
CONTROL
3.7.36 The rocket system shall be initiated
by the operation of a suitable control, readily
accessible to the pilot's throttle hand with harness
locked.
3.7.37 The control shall be designed to

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obviate inadvertent operation.
3.7.38 If a second pilot is to be carried
during the trials, a duplicate control of like design
shall be provided at the second pilot's station
unless he has ready access, with harness locked,
to the first pilot's control.
OPERATION
3.7.39 Each part of the rocket installation
shall develop its full thrust within 3 seconds of
being brought into action by the operation of the
pilot's control.
STRENGTH OF THE AEROPLANE STRUCTURE
3.7.40 The aeroplane structure shall have
proof and ultimate factors not less than 1.125 and
1.5 respectively under the maximum load derived
from the aerodynamic and inertia forces and the
forces produced by the rocket system during the
stall or spin.
STRENGTH OF THE ROCKET INSTALLATION
3.7.41 The flight envelope, including
environmental conditions for which the rocket
installation is designed, shall be agreed with the
Project Team Leader (PTL)
ELECTROMAGNETIC COMPATIBILITY
3.7.42 The installation shall be safe against
stray electrical and radio hazards and the initiating
arrangement shall be designed in accordance with
the requirements of RAE Specification WE 659.
FIRE HAZARD
3.7.43 The aeroplane shall be suitably
protected from the effects of fires which might be
caused by the rocket installation; warning of such
fires in the air shall be provided to the pilot.
GROUND TESTS

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3.7.44 Before any of the flight tests of
Clauses 3.7.46 and 3.7.47 are made, the rocket
shall be fired in a test rig which is representative of
the aeroplane installation, to indicate the efflux
pattern, component temperature levels, correct
functioning of the firing circuits and the
effectiveness of any fire warning and fire protection
measures.
3.7.45 Tests shall be carried out to ensure
that the functioning of the rocket installation is not
affected by the operation of any electrical or
electronic systems likely to be used during the
flight trials.
FLIGHT TESTS
3.7.46 If so required by the Project Team
Leader (PTL), the rocket installation shall be
demonstrated in straight flight.
3.7.47 If so required by the Project Team
Leader (PTL), flight tests shall be made to
demonstrate compliance with Clause 3.7.8

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3.8 TARGET-TOWING INSTALLATIONS

INTRODUCTION

3.8.1 This Clause states the design requirements for target towing aeroplanes. The requirements relate to the aeroplane itself and are not
necessarily those of the towed equipment.

3.8.2 The towing duties may fall within the following categories:-

(a) Anti-aircraft gunnery/missile practice

(b) Anti-missile gunnery/missile practice

(c) Combat decoy target deployment

(d) Surveillance device deployment

(e) Deception device deployment

(f) Radar target deployment

3.8.3 The requirements cover both ground launched and air launched systems.

3.8.4 The towing of manned aircraft (i.e., gliders or sailplanes) is not covered by these requirements.

3.8.5 The specification of the towed assembly to be used with the installation shall be approved by the Project Team Leader (PTL).
Consideration shall be given to the intended operational use of the towed assembly when defining its expected durability. The towed assembly
shall be sufficiently robust to withstand all phases of deployment and, if required, subsequent recovery/retrieval.

3.8.6 Definitions and a Glossary of terms used in this Clause are given in Part 0 Definitions.

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OPERATIONAL REQUIREMENTS
3.8.7 Flight envelopes appropriate to
carriage, launch, tow, recovery and release of the
towed equipment shall be established in
accordance with the Aeroplane Specification.
3.8.8 The carriage, launch, tow, recovery The trajectory adopted by the target/tow line Consideration shall also be given to the effects of
and release of the towed equipment shall not affect during launch, tow, recovery and release of the jettisoning the equipment, frangible covers and
the operational function or safety of the aeroplane. towed equipment shall not interfere with the protective devices, including the random impact of
adequate function of the various aeroplane discarded pieces on the aeroplane structure.
flying control surfaces.
3.8.9 With recoverable systems provision See Part 1, Section 4, Clause 4.13 "Ground
shall be made for landing and, when applicable, Clearance".
arresting the aeroplane with the target stowed
(carriage).
3.8.10 It shall be possible to jettison safely, The details shall be discussed and agreed with
within an appropriate envelope, all external the Project Team Leader (PTL)
equipment that could be critical for operational or
flight safety reasons.
AERODYNAMIC AND FLYING QUALITIES
3.8.11 The requirements of Part 1, Section (a) The effects of internal and external See Leaflet 15.
2 shall apply for such symmetric and asymmetric equipment and of the intentional release/jettison
combinations of internal and external equipment as of such equipment on the mass and its
stated in the relevant aeroplane specification. The distribution and on the aerodynamic
Contractor shall decide, in conjunction with the characteristics of the aeroplane shall be
Project Team Leader (PTL), the equipment load considered.
combinations which are critical for flying qualities
design and demonstration. (b) In deciding which store combinations are
critical, Aeroplane Normal States embodying
deliberate asymmetries, and any shorter-term
asymmetries, which may be experienced, shall
be considered for each mission flight phase.
3.8.12 When the equipment contains
expendable components, the requirements of Part
1, Section 2 shall apply throughout the range of

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component loadings.
3.8.13 The flight envelope for change of
mass distribution or release of components shall
be governed by the most restrictive of the flight
envelopes before, during and after the event.
3.8.14 For flying qualities of Level 1, See Part 1, Section 2, Clause 2.1.19
operation of movable parts of the equipment shall
not cause flutter, buffeting, trim changes, or other
related dynamic instabilities which impair the role
effectiveness of the aeroplane under any pertinent
flight condition.
3.8.15 The intentional release of any
equipment shall not result in objectionable flight
characteristics for Levels 1 and 2 and shall never
result in dangerous or intolerable flight
characteristics in Level 3.
3.8.16 Target towing equipment, which is
fitted to an aeroplane to, facilitate an additional
role, or to introduce a new role within which the full
operational flight envelope is not required, may be
designed to lower flight envelope limits and/or
other operational limitations to be agreed by the
Project Team Leader (PTL)
ENVIRONMENTAL CONDITIONS
3.8.17 The fixed target towing installation
shall be capable of operating under the
environmental conditions quoted in Part 1, Section
7, Clauses 7.1 and 7.2 or as modified by the
Aeroplane Specification as appropriate to this role.
3.8.18 Where the Target Towing Aeroplane See Leaflet 14, Para 4.
Specification allows less arduous conditions to be
used for the operation of the target towing
installation than for the original aeroplane then
consideration shall be given to the effects of
carriage of the target towing installation under the

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more arduous conditions.
STRUCTURAL INTEGRITY
3.8.19 The target towing installation and
the associated aeroplane structure shall be
designed to the requirements of Part 1, Section 3
and capable of withstanding all loads generated by
the towed target system throughout the approved
flight envelope for the system and at all phases of
deployment.
3.8.20 Particular attention shall be given to
the effects of shock loads which may occur notably
during the deployment phases of launch or ground
snatch.
3.8.21 Where the intrinsic strength of the
towed assembly exceeds that of the target towing
installation or associated aeroplane structure
provision shall be made to relieve the aeroplane of
excessive loads.
3.8.22 The effects of loads arising from the
target towing system shall be considered in terms
of potential fatigue life of the aeroplane.
INSTALLATION REQUIREMENTS
CONSTRUCTION
3.8.23 The construction of the fixed target
towing installation shall conform to BS 3G
100(Multipart) "General requirements for
equipment for use in aircraft".
3.8.24 Where the target towing installation
takes the form of an external store. MIL-Std-8591
"Airborne Stores, Suspension Equipment and
Aircraft-Store Interface (Carriage Phase)"
POSITION OF TARGET TOWING INSTALLATIONS
3.8.25 The position of the target towing
installation shall be such that the carriage,

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launching, towing, recovery and shedding (as
appropriate) of the towed assembly does not cause
interference with any part of the aeroplane.
3.8.26 Particular attention shall be paid to the
relationship between the aeroplanes attitude, and
the trajectory of the towed assembly during launch.
3.8.27 The proximity of the aeroplanes
engine(s) to the target towing installation shall be
examined to ensure that:-

(a) There is no danger of entrainment See Leaflet 15, Para 1

or entanglement of the towed assembly,
including any discarded portions thereof, with
the engine intake or propeller.

(b) There are no adverse effects

caused by the impingement of engine efflux
or propeller slipstream on the fixed
installation or towed assembly.
3.8.28 The presence of the target towing
installation shall not impede the aircrew's means of
escape from the aeroplane in an emergency.
GROUND CLEARANCE
3.8.29 The requirements of Part 1, Section
4, Clause 4.13 "Ground Clearance" shall be met.
USE OF ARMAMENT STATIONS
3.8.30 Where an aeroplane carries target See Clause 3.1
towing equipment as an alternative store on an
armament installation it shall be confirmed that the
integrity of the armament installation is not
compromised and that reversion to the original
installation is readily achievable. Under these
circumstances it may be necessary to restrict the
choice of target towing controls to the existing
armament controls.

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USE OF AEROPLANES POWER SOURCES
3.8.31 Where the towed target installation
draws any form of power (electric, hydraulic,
pneumatic etc.,) from the aeroplanes systems it
shall be possible for the aircrew to isolate the
target towing installation from the aeroplane's
system.
POWER LOSS
3.8.32 Towed target installations shall be
designed to a "fail safe" philosophy so that in the
event of power failure (aeroplane sources or self-
generated), the installation shall shut down without
endangering the aeroplane.
3.8.33 In the event that the power loss is
irretrievable it shall still be possible to jettison the
towed assembly.
FITTING AND LOADING THE INSTALLATION
3.8.34 The target towing installation shall
be designed to allow easy fitting to and removal
from the aeroplane of the fixed installation by
personnel wearing appropriate protective clothing.
3.8.35 Adequate clearances must be Loading methods should be as simple and direct
provided around the fixed installation to allow all as possible.
necessary servicing and inspection, including
loading of tow lines and targets, subsequent
arming of targets (where required) and final
inspection by the aircrew.
3.8.36 Where the fitting or removal of the Wherever possible existing in-service equipment
fixed installation, tow line or target necessitates the shall be used.
handling of a heavy or awkward load provision
shall be made for use of hoisting equipment.
ROLE CHANGE
3.8.37 Where the target towing is an
additional role for the aeroplane the reversion to
the operational role(s) should be readily

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achievable.
TOWED ASSEMBLY ATTACHMENT
3.8.38 Where the attachment of the towline
or target is undertaken on the airfield runway the
following points shall apply:-

(a) It shall be possible to couple the

towed assembly to the attachment point
without the use of tools.

(b) Indication that the towline

attachment is correctly cocked shall be
evident to the ground crew.

(c) It shall be possible to uncouple the

towed assembly from the attachment point
without access to the cockpit controls and
without the use of tools.
CONTROLS AND INDICATORS
COCKPIT CONTROLS
3.8.39 The installation shall incorporate
sufficient controls to ensure the safe operation by
the aircrew, of the target towing system at all
phases of the sortie.
3.8.40 Required controls are:-

(a) Launch (air launched systems)

(b) Discard, or recover, towed

assembly

(c) Jettison Towed Assembly by:-

(1) Pilot’s emergency circuit selection

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(2) Ejection synchronised selection

(3) Operator's selection

(d) Jettison External Installation (where

permitted) by: - Pilot's emergency circuit
selection.

(e) Stop streaming (variable tow length

systems only)
COCKPIT INDICATORS
3.8.41 The installations shall incorporate The extent of the indicators will depend on the See Leaflet 18, Para 2.
sufficient indicators to ensure the safe operation, complexity of the installation.
by the aircrew, of the target towing system at all
phases of the sortie.
SYSTEM USAGE
INTRODUCTION
3.8.42 There is a paramount requirement
for target towing aeroplanes either to recover or to
shed any towed assembly prior to landing.
Consequently, the reliability of techniques and
devices to achieve this end must be proven to the
satisfaction of the Project Team Leader (PTL)
TARGET RECOVERY
3.8.43 Provision shall be made on Aspects to be assessed are:-
recoverable systems to assess the condition of the
returning target prior to commitment to complete (a) Target stability
the recovery sequence.
(b) Target position relative to the
towing aeroplane

(c) Target damage state

3.8.44 Provision shall be made for the safe
removal on the ground of a recovered target which
may be in a damaged state.

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TERMINATION OF TOWING OPERATION
3.8.45 The aeroplane shall be provided One of these options shall be the jettison of the
with at least two options for terminating the towing towed assembly.
operation.
3.8.46 It shall be possible to shed the See Leaflet 17.
towed assembly from the towing aeroplane for both
discardable and recoverable systems in all states
of towline tension and velocity. Tensions shall
range from minimal values representing only a
remnant of the towline surviving to maximal values
representing the Ultimate Tensile Strength of the
towline. In order to cater for a worst case condition
the presence of weak links in the towed assembly
shall be discounted. Velocities shall include
overspeed conditions.
3.8.47 Recoverable systems shall
incorporate at least one means of shedding the
towed assembly in addition to their normal method
of recovery.
3.8.48 Towline cutting devices shall not
pose a hazard to the aircrew or to ground
personnel.
3.8.49 Discardable systems shall
incorporate at least two means of shedding the
towed assembly. These means shall be powered
from independent sources so that the loss of one
source does not prevent the operation of the
release. One of these means may be the towline
jettison selection.
3.8.50 Variable tow length towed
assemblies shall be capable of being shed at any
point along their length, i.e. cut.
3.8.51 Winch systems shall not have their
towline secured permanently to the reel; it shall be
possible to stream the entire towline clear of the

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aeroplane if necessary.
3.8.52 Towline cutting devices shall be
capable of severing the towline at any speed
attainable by that towline, including runaway
conditions in excess of normal speeds.
CONTAINMENT OF UNCONTROLLED SITUATIONS
3.8.53 On installations involving the See Leaflet 15, Para 2.
streaming, or recovery of the towlines the
aeroplane shall be protected against the
consequences of uncontrollable operation.
3.8.54 Consideration shall be given to the
protection of the aircrew where target towing
equipment is fitted inside the aeroplane cabin.
VIBRATION
3.8.55 The presence, and operation, of the (a) See Part 1, Section 1, Clause 1.1.6
target towing installation shall not cause
unacceptable vibration in the aeroplane. (b) Rotating masses, especially variable
masses such as towlines on reels, need particular
attention to ensure that no critical periods are
created at any achievable combination of mass
and rotational speed.
RADIO/TELEMETRY LINKS
3.8.56 Where there is a
command/telemetry link between the towed target
and the aeroplane or any other station, it shall be
confirmed that this link does not cause any
interference with any radio or avionics system in
the aeroplane.
JETTISON
3.8.57 Target towing aeroplanes have
particular jettison requirements due to the nature of
their tasks. It is imperative that the aeroplane be
able to rid itself of any towed assembly which it
cannot recover before the aeroplane itself is forced

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to land. The jettison actions are in addition to
normal modes of recovery or shedding of towed
assemblies.
3.8.58 Jettison selections are part of the
"Emergency" procedures; as such it shall be
possible to operate the jettison circuits despite a
complete failure of the "Normal" supply system.
3.8.59 Towed assemblies shall be capable
of being jettisoned separately from any other
equipment and at any phase of the towing
operation.
3.8.60 The pilot shall be provided with the
means of jettisoning the towed assembly and also,
where permitted, the fixed installation.
3.8.61 Where jettison of the fixed
installation, or any part of it, is permitted
consideration shall be given to all system states
including:-

(a) Store fully loaned with target and

towline stowed.

(b) Store operational with target and

towline streamed.

(c) Store empty with target and towline

already discarded.
3.8.62 On long deployment/recovery period
systems it shall be possible to jettison the towed
assembly whilst the towline is running, in either
direction, at any speed attainable by the installation
including overspeed conditions, and at any point
along its length.
3.8.63 In the event of the aircrew having to
eject from the aeroplane before being able to

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jettison the tow assembly by other means a jettison
of the streamed tow should be activated by the
ejection system.
EXPLOSIVES AND PYROTECHNICS
ELECTRO-EXPLOSIVE DEVICES
3.8.64 The requirements of Clause 3.4
"Installation of Explosive Devices" shall be met for
those towed target installations incorporating such
devices.
3.8.65 Cable terminations in connector
systems, which activate EEDs should not occupy
connections adjacent to other connections which
carry standing voltages or which can become
energised during the normal operation of the target
towing system.
PYROTECHNIC DEVICES
3.8.66 Where the towed assembly
incorporates pyrotechnic devices the ignition of
such devices in the stowed condition shall not be
possible.
3.8.67 Where radio frequency
transmissions are used to activate the pyrotechnic,
or similar hazardous devices, safeguards shall be
established to ensure safe handling of the
pyrotechnics whilst being loaded or unloaded.
ELECTROMAGNETIC COMPATIBILITY
3.8.68 It shall be confirmed that the EMC of Particular attention shall be given to installations
the aeroplane is not compromised by the presence incorporating active electromagnetic devices. In
of the target towing installation. such cases the EMC state of the aeroplane shall
be assessed in the following conditions:-

(a) Target stowed. (a) The consequence of inadvertent EM

activity shall also be considered.

(b) Target streamed in close (b) Consideration shall be given to the

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proximity (both launch and recovery requirements of Part 1, Section 6, Clause 6.10
phases as appropriate) "Electromagnetic Compatibility of Safety Critical
Systems" and to Def Stan 59-411
"Electromagnetic Compatibility"
(c) Target streamed to operational
tow lengths.
TESTS
3.8.69 On prototype aeroplanes (and on (a) The Contractor shall demonstrate that
the first of all aeroplanes where a new installation the installation functions satisfactorily and meets
is used, or use of special equipment has been the requirements of the relevant Aeroplane
approved) the Contractor, on completion of the Specification.
installation shall subject it to such tests as
considered necessary by the Project Team Leader (b) Ground tests shall demonstrate that, in
(PTL) compliance with the aeroplane specification:-

(1) The replaceable components of

the tow system can be loaded and
unloaded when the aeroplane is at rest on
the ground.

(2) Examination of the installation, in

the form of a ground functions check
satisfactorily confirms that the installation
is fit for flight test.

(3) The role change capability, where

required, is feasible.

(4) The jettisoning arrangements will

function satisfactorily.
FLIGHT TESTS
3.8.70 Flight trials shall provide a sufficient The flight trials shall provide evidence of The flight test objective is to demonstrate that the
basis, together with ground test and calculation to satisfactory carriage, launch, tow, recovery and installation is safe and satisfies the operational
show that the design of the aeroplane is sound and release of the target and jettison of the requirement for the aeroplane concerned for the
that the limiting environmental and performance equipment. Thus the contractor's trials must carriage and deployment of the towed assembly.

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factors are not exceeded. explore the extremities of the flight envelope if
there is any doubt as to flight safety. The Project
Team Leader (PTL) in conjunction with the
aeroplane designer is responsible for
recommending the limits to which service pilots
may fly. The flight trials shall cover:-

(a) CARRIAGE - Carriage flights are

to be flown within the required operational
envelope of the aeroplane. Sufficient in-
flight data is to be obtained to validate the
design. Any adverse effects on aeroplane
handling or performance shall be noted.

(b) LAUNCH TOW, RECOVERY

AND RELEASE - The evidence provided
by the target launches, tows, recoveries
and releases must be sufficient to validate
the design.

(c) JETTISON - Jettison tests shall

be made. The jettison evidence must be
sufficient to validate design.
CROSS REFERENCES
3.8.71 A number of requirements which
influence the installation of target towing
equipment appear elsewhere in this publication
and the most significant are listed in Table 3.

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3.9 REDUCTION OF VULNERABILITY TO BATTLE DAMAGE

INTRODUCTION

3.9.1 This Clause contains design aims and requirements which enhance the survivability of the aeroplane and its crew by reducing their
vulnerability to battle damage, consistent with overall survivability requirements. They are applicable to all types of aircraft and RPAS except
research, primary trainer, and basic trainer aircraft unless required by the Airplane Specification.

3.9.2 Aircraft should be designed to reduce their vulnerability in operations to as low a value as possible, consistent with overall survivability
requirements. The vulnerability of an aircraft is influenced by a large number of aircraft design features, for example, whether or not the aircraft
contains duplicated and separated flight controls and whether there is a single or twin engine configuration. Likewise, whether there is a single pilot
or a pilot and co-pilot also has an influence on the aircraft’s vulnerability. (Also see Clause 3.10 on Crew Protection). Vulnerability also depends
upon the threat weapon in terms of it damaging effects, as well as the weapon’s attack direction distribution around the aircraft. Two documents
that are relevant in this connection are the Project Threat Statements and CONOPS. Because aircraft vulnerability is affected by so many of these
and other aspects, this Def Stan clause does not attempt to specify kill probabilities that have to be achieved. Instead, it is recommended that
vulnerability assessments are undertaken so that the influence of these different influential aspects can be investigated and their contribution to the
aircraft’s vulnerability reduced. Overall, this should result in an optimised aircraft design that minimises aircraft vulnerability, while still keeping the
solution within the other design constraints that are acting, such as maximum weight, centre of gravity limits, etc. The different types of kill
categories are defined in this document for reference purposes and they should be agreed with the PT Leader along with the appropriate kill
category time-frame (Refer to Leaflet 19, Section 1.2).

3.9.3 See Leaflet 19 for general background information on Reduction of Vulnerability to Battle Damage, Design Aims, Vulnerability Analysis,
Protection Measures, Battle Damage Repair, and Kill Categories. Refer to Part 1, Section 4, Clause 4.22 for Crash Landing and Ditching
requirements and Clause 3.10 for Protection Systems for Aircrew.

3.9.4 The design requirements below should be applied in conjunction with the Project Threat Statement provided by the PT Leader. (Refer to
Definitions in Leaflet 21). It is recommended that the threat definitions are referenced in relation to the particular project.

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DESIGN
3.9.5 Based on the defined Threat List (a) The vulnerability of all systems shall be Consideration of systems vulnerability requires a
and CONOPS, the principal design aims shall be considered. The vulnerability of critical systems vulnerability analysis. A vulnerability model should
to: shall be reduced by application, where relevant, be agreed between the Design Organisation for
of the following techniques: the Aircraft and the PTL.
(a). Maximise the probability that no
single Threat Effect will degrade the Flying (1) Configuration adjustment. See Leaflet 19 on how to identify critical systems.
Qualities of the Aircraft below Level 3 of Def
Stan 00-970 Part 1, Section 2, Clause 2.1.19 (2) Redundancy.
(fixed wing aircraft) and Part 7, Section 6,
Leaflet 600 (rotary wing aircraft). (3) Separation.

(b). Minimise the overall aircraft Pk|H i.e. (4) Change of dimensions.
minimise the probability that the aircraft fails
to maintain controlled flight or to complete its (5) Additional protection measures.
mission, (consistent with overall survivability
requirements). (6) Use of components designed to
tolerate battle damage.

(b) Consideration shall be given to isolation See Part 1, Section 5, Clause 5.2 and Part 1,
and suppression of fire and sources of ignition Section 4, Clause 4.26
shall be separated effectively from flammable
fluids and gases.

(c) Consideration shall be given to the

following when assessing any material for use in
the design:

(1) Nuclear, biological and chemical

effects.

(2) Reparability.

(3) Residual structural strength.

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(d) Consideration shall be given to the See Part 1, Section 6, Clause 6.14
probability of a hit on each Pressurised Gas
Storage Vessel or Gas/Oil Hydraulic
Accumulator. A hit could make a hole in the
vessel thereby releasing energy, which may
damage structure and systems in the vicinity.
Also, fragments could be generated by the
disintegration of the pressure vessel’s casing.
Such fragments could be lethal to critical
systems, unless containment or protection
measures are taken.
VULNERABILITY ANALYSIS
3.9.6. The Design Organisation shall The Vulnerability Analysis shall include a See Clause 3.10
consult with the Project Team Leader (PTL) and Casualty Reduction Analysis.
establish whether, and how, the vulnerability of the
aircraft to Defined and Specified Threat Effects will
be assessed, and consider how consequent design
changes, if any, will be introduced.
BATTLE DAMAGE REPAIR
3.9.7 The designer shall consider and
provide for the repair of structure and of flight and
mission critical systems following battle damage
and shall incorporate such design features as will
facilitate battle damage repair.

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3.10 PROTECTION OF AIRCREW AGAINST CONVENTIONAL WEAPONS1

GENERAL

3.10.1 This Clause contains requirements for the provision of efficient ballistic protection systems for aircrew, particularly armour, and is additional
to those requirements relating to aircrew protection contained in Clause 3.9. (Crew-related protection requirements in Clause 3.9 refer to the
minimisation of the aircraft Pk|H and the maximisation of the probability that the post-damage flying qualities will be above a minimal level). For a
side-by-side, twin-crew aircraft, the design aim from Clause 3.9 would be to avoid the loss of the “crew system” so that flight control of the aircraft
can be maintained. In this case, the design aim would try to minimise the probability of both crew members being killed from the same weapon
interaction. One possible solution for a side-on attack direction would be to place an armour panel between the two crewmembers. This solution
would not be ideal for protecting the individual crew members but it is ideal for protecting the “crew system”. In contrast, Clause 3.10 covers the
protection of the aircrew as individual beings and is a different consideration and would lead to different armour solutions. However, the
importance of the aircraft’s survival cannot be overlooked because the survival of the individual crewmembers relies on the aircraft maintaining
controlled flight. For this reason, it is anticipated that a balance will have to be struck between the level of “crew system” protection (i.e. aircraft
protection, upon which the individual aircrew survival depends) and the level of individual aircrew protection.

3.10.2 The requirements are applicable to all types of aircraft (excluding RPAS) except research, primary trainer, and basic trainer aircraft unless
required by the Aeroplane Specification (Leaflet 20).

REQUIREMENT COMPLIANCE GUIDANCE

DESIGN
3.10.3 Structure and non-critical See ASCC Air Standard 61/102/14A (1) (Armour
components shall be used as much as possible for Protection for Aircrew) & Leaflet 20.
shielding the aircrew. In addition to providing
protection from the threat, protection should also
cover fragmentation and over-pressurisation
effects from onboard energetic sources. (Clause
3.9). As far as possible, energetic sources should
be placed as far as possible from crew members to
minimise adverse effects.
3.10.4 Multiple pilot stations shall be
separated as much as practicable and
consideration shall be given to the provision of
shielding between them.
3.10.5 Aircrew protection systems shall not
restrict aircrew mobility and vision, nor access to

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controls to the extent that the accomplishment of
the mission is affected. They shall also not
preclude the use of survival equipment nor
Nuclear, Biological and Chemical protection
equipment.
3.10.6 Aircrew protection systems shall not
interfere with Normal Entrance and Exit (Part 1,
Section 4, Clause 4.20), nor Emergency Escape
(Part 1, Section 4, Clause 4.23), nor constitute a
hazard in the event of a Crash Landing or Ditching
(Part 1, Section 4, Clause 4.22)
3.10.7 Materials used for armour protection See Leaflet 20 Para 3.
or fragment suppression shall be selected to
provide the most suitable characteristics for their
purpose at each location.
3.10.8 The Vulnerability Analysis of Clause Where tests are considered to be necessary, a Consideration shall be given to the need for tests
3.9 shall include a Casualty Reduction Analysis. programme shall be agreed with the Project to verify the protection provided against
Team Leader (PTL) Conventional Defined and Specified Threats
(Clause 3.9)
ARMOUR
3.10.9 Where armour is provided for
protection of the crew, the following requirements
apply:

(a) Where the armour is an integral part

of the aircraft structure, the structure
including the armour shall meet all relevant
design requirements.

(b) Where the armour is an integral part The forces on the armour attachments and back-
of the seat, it shall be capable of sustaining up structure arising from direct projectile impact
the loads applied to the seat, and the seat, shall also be considered.
including the armour, shall meet all relevant
design requirements.

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(c) Where the armour is attached to the
aircraft structure, it shall be treated as an
item under Part 1, Section 4, Clause 4.22

(d) Where the armour is attached to the

seat or mounted on the pilot or crewman, the
mass of the armour shall be accounted in the
mass of the seat in all other design
requirements such as crashworthiness, when
designing to meet the requirements of Part 1,
Section 4, Clause 4.22

(e) Materials which generate spall

particles shall not be used unless suitable
provision is made to suppress the spall, to
avoid injury to aircrew and to prevent
increase of vulnerability to an unacceptable
level.

(f) Body armour shall be designed for

quick release in an emergency.
REDUCTION OF VULNERABILITY TO BATTLE DAMAGE
3.10.10 Any equipment specifically required
to protect the crew against Defined and Specified
Threat Effects will be listed in the Aeroplane
Specification. Where space provision, fixed fittings
and/or support systems for this equipment are
required, this also will be stated in the Aeroplane
Specification.

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3.11 PROTECTION FROM THE EFFECTS OF NUCLEAR EXPLOSIONS, LASER WEAPONS, CHEMICAL AND BIOLOGICAL WARFARE
AGENTS

INTRODUCTION

3.11.1 This Clause specifies the design requirements, which will enable the Aircraft to survive and operate in Nuclear, Biological, Chemical
(NBC) and/or laser environments and their associated decontamination requirements where relevant. Relevant definitions are given in Leaflet 21.

3.11.2 When there is a requirement for NBC and/or laser hardening; this Clause specifies the requirements to be applied for the protection of
the Aircraft and crew both in the air and on the ground in an NBC/laser environment which is considered survivable.

3.11.3 Several of the reference documents quoted in this Clause are classified and are only available on a need-to-know basis. Defence and
protection against these threats is a specialised subject, and advice and information will or may need to be obtained from the Project Team
Leader.

REQUIREMENT COMPLIANCE GUIDANCE

GENERAL
3.11.4 The Aircraft and its installed When defining the form of aircraft CB hardening The definition of NBC and laser protective
equipment shall be designed to be operated by to be adopted consideration shall be given to the clothing for the flying crew of an Aircraft will be
personnel wearing full NBC and laser protective following for all materials and design features through the Project Team Leader (PTL) or FAP
clothing. If full NBC protective clothing is not worn incorporated in the Aircraft and its equipment: 108B-0001-1(Aircrew equipment assemblies). For
by the aircrew then the Aircraft shall be equipped further guidance see Leaflet 23.
with a system capable of supplying suitably (a) Immediate damage effects.
pressurised and filtered air to the crew such that
the crew are properly protected against the effects (b) Prolonged contamination.
of NBC hazards (CW agent liquid and vapour, BW
agent in the form of aerosol and nuclear hazards in (c) The long-term consequences of
the form of dust). Appropriate levels of protection short-term contamination.
are required for the eyes, the skin, and the oro
nasal tract. These protection levels will be advised (d) Decontaminants and
to the contractor as required by the Project Team decontaminability.
Leader.
The acceptability of the interfaces to the AEA NBC AEA and aircraft interfaces are a specialist
Where NBC protection is worn by flying personnel, and the flying personnel shall be by design area. Guidance shall be sought from the Project
the life support system (where fitted) and any review and test where this is required by the Team Leader.
aircraft interfaces (including ejection seats for Project Team Leader.

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aircraft so equipped) shall be compatible with
providing the required levels of protection to the
personnel (NBC and non-NBC)
3.11.5 The design of the Aircraft and Electrical and other performance and / or safety Defence standard 08-11, 08-41 and 00-72 provide
equipment shall be such that maintenance, critical interfaces shall be designed in additional guidance.
replenishment and rearming tasks can be accordance with extant CB hardening
performed by personnel wearing full NBC requirements. Specific requirements will vary by the platform,
protective equipment. The design of the Aircraft and guidance shall be obtained through the
shall be such that CW agents, in particular in liquid Design review shall be used to determine Project Team Leader.
form, will not affect the performance, function, or compliance, with test as required by the Project
safety of critical interfaces such as weapon and Team Leader.
other hardpoints.
NUCLEAR ENVIRONMENT REQUIREMENTS
GENERAL
3.11.6 The basic nuclear survivability aim The application environment will be defined by (a) Typically, this environment will be in
for military aircraft weapon and ground systems, the PTL. accordance with that given in Def Stan 08-4 Part
and installations shall be to comply with the 4/2 Chapter 04-05, Table 1.
application environment.
(b) The factors which should be considered
during the design phase to reduce vulnerability to
the effect of nuclear weapons explosions are
described in Def Stan 08-4 Part 4 and Leaflet 22.
The effects of contamination from nuclear fallout
should also be considered in the design process.
DESIGN
3.11.7 The design objective for the nuclear Environmental criteria are laid down in Def Stan
hardening of the Aircraft, its installed equipment 08-4 Part 4/2 Chapter 04-05 Table 1. Table 6 of
and installed weapons, shall be compliance with this document is typical of the parameters
the environmental criteria and that associated with involved in the Nuclear Hardening design.
the nuclear threat by friendly weapons.
3.11.8 The principal design aim shall be to
maximise the probability that the threat effects
defined in Table 1 or in the Aircraft Specification
will not degrade the flying qualities of the Aircraft
below Level 3 of Part 1, Section 1, Clause 2.1.19

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3.11.9 During the Concept Phase the (a) The results of the initial Nuclear
Design Organisation shall undertake initial Survivability Feasibility Studies shall be
feasibility studies of the design to quantify its submitted to the Project Team Leader (PTL) for
Nuclear survivability level in relation to technical acceptance of the estimated hardness levels of
achievability and all the penalties incurred. the proposed design relative to the balanced
Nuclear Hardness Criteria referred to in Clause
3.11.7 with special consideration to any
penalties as mentioned in Clause 3.11.8
If the estimated hardness levels quoted in the
initial Nuclear Survivability studies are not
acceptable to the Project Team Leader (PTL), it
may be necessary to redefine the environment
criteria. The Design Organisation shall
incorporate design changes aimed at achieving
the values of the amended criteria and a further
Nuclear Survivability Study Report shall be
submitted to the Project Team Leader (PTL)

(b) The acceptability of the design will then

be considered by the Project Team Leader
(PTL) using scenario studies when appropriate.
3.11.10 Consideration shall be given to It shall be assumed that essential aircrew are
damage to the Aircraft caused by its own or protected against flashblindness.
friendly forces weapons.
OPERATIONAL CONDITIONS
3.11.11 (a) The flight conditions and configuration of
the Aircraft, e.g., speed, altitude, attitude,
autopilot engagement, and stores configuration,
immediately prior to exposure to the nuclear
burst will be defined by the Project Team Leader
(PTL)

(b) Similarly, when the Aircraft is on the

ground e.g., parked in the open, sheltered, or in
a Hardened Aircraft Shelter (HAS) the conditions

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will be defined by the Project Team Leader
(PTL)
CHEMICAL AND BIOLOGICAL ENVIRONMENT REQUIREMENTS
GENERAL
3.11.12 The level of chemical hardening to The design shall consider the need for The properties of chemical warfare (CW) agents
be embodied in the Aircraft, its installed equipment survivability and operability in a Chemical and and decontaminants together with the factors
associated armament, interfaces (including Biological warfare environment based on criteria which should be considered during the design
electrical) and ground systems shall be determined laid down by the Project Team Leader (PTL) phase to reduce vulnerability of equipments to the
by the environment detailed in the relevant which will permit Aircraft operations to be carried effects of such agents and decontaminants were
Specification. out on an airfield which is or has been subjected previously contained in CDE Technical
to attack by CW agents. Memorandum No 79. Additional and up-to-date
guidance can be sought through the Project Team
Design review shall be used to determine Leader. Reference should be made to Defence
compliance, with test as required by the Project Standards 08-11, 08-41 and 00-72, which provide
Team Leader. specific guidance and advice on hardening and
protection measures where these are required.

It should be recognised that the majority of

decontaminants that are in service cannot be
used on aircraft because of their corrosive effects.
The Design Organisation shall assume that
decontamination, where required, must be
achievable using weathering and washing, using
hot water and approved detergents only. The
Design Organisation shall ensure that the Aircraft
design is compatible with decontamination using
these techniques, where decontamination is a
requirement.
DESIGN
3.11.13 The design objective for the (a) Aircraft components and equipments See also Leaflet 22.
Chemical and Biological hardening of the Aircraft is shall be categorised as follows:
that its installed equipment and weapons and their Guidance should be sought from Defence
interfaces (including electrical) shall be compliant (1) those items which are likely to be Standard 08-41. This defines those materials
with the criteria laid down by the Aircraft subjected to direct chemical attack. which are degraded / damaged by CW agents
Specification or Project Team Leader (PTL) and those which are not. The defence standard

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(2) those items not subject to direct also provides design advice for compatibility with
chemical attack but which may become CB hardening and decontamination. Advice on
contaminated due to ingestion. any materials that are not contained in the
defence standard can be sought through the
(b) Aircraft components and equipment in Project Team Leader.
the interior of the Aircraft shall not be considered
to be subject to direct attack unless they can be Defence Standard 08-11 contains additional
contaminated due to ingestion of CW agents via advice and guidance, including in relation to
external apertures, air systems, etc. It should be interior / internal contamination and means to
recognised that CB agents can also enter the address this.
aircraft through the engines and APU, should an
APU be installed. These agents will then Defence Standard 00-72 describes the
contaminate the interior of the Aircraft, including requirements for CW agent resistant paints and
its environmental control system, should one be finishes.
installed.

Design review shall be used to determine

compliance, with test as required by the Project
Team Leader.
3.11.14 The principal design aim shall be to Design review shall be used to determine Refer to Defence standards 08-11 and 08-41.
maximise the probability that the threat effects compliance, with test as required by the Project Further advice should be sought through the
defined in Table 5 or in the Aircraft Specification Team Leader. Project Team Leader.
shall not degrade the flying qualities of the Aircraft
below Level 2 of Part 1, Section 2, Clause 2.1.19
3.11.15 The Aircraft shall be capable of safe Design review shall be used to determine Refer to Defence standards 08-11 and 08-41.
return to base after chemical (or biological) compliance, with test as required by the Project Further advice should be sought through the
contamination and without decontamination Team Leader. Project Team Leader.
measures being applied it shall be possible to
utilise the Aircraft for a specific number of missions
involving repeated contamination or for a period
defined by the Project Team Leader (PTL) without
there being a requirement to undertake any
unscheduled maintenance as a result of such
contamination.
3.11.16 All components and equipment Design review shall be used to determine Refer to Defence standards 08-11, 08-41 and 00-

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subjected to direct attack shall be designed to compliance, with test as required by the Project 72. Further advice should be sought through the
resist the damaging effects of the specified CW Team Leader. Project Team Leader.
agents and their decontaminants and to facilitate
the rapid reduction or elimination of the hazard
from any chemical contamination which may be
found in the Aircraft or equipment. This includes all
electrical interfaces and connectors.
3.11.17 All components and equipment Design review shall be used to determine See Clause 3.11.4 for guidance on definition of
subjected to direct attack shall be designed to compliance, with test / demonstration as equipment and clothing standard.
allow Service personnel dressed in full Individual required by the Project Team Leader.
Protective Equipment to handle, decontaminate, or Further advice should be sought through the
service the equipment. Project Team Leader.
3.11.18 During the Project Definition stage The results of the feasibility study shall be As part of this study, the contractor should review
the contractor shall undertake initial Chemical and submitted to the Project Team Leader (PTL) for and take into account the information contained in
Biological survivability feasibility studies of the acceptance of the proposed design relative to Defence standards 08-11 and 08-41.
design of the structures, system and components the criteria referred to in Clause 3.11.13 with
in relation to technical achievability and the special consideration to the options set out in Additional specialist advice and information may
penalties incurred in performance, reliability, Clause 3.11.18 be obtained through the Project Team Leader.
maintainability, cost limitations and to the
disposition of component/system within the Aircraft
based on the specified criteria.
3.11.19 The design, structure and Design review shall be used to determine Where a high degree of hardening is required
equipment shall prevent as far as practical the compliance, with test / demonstration as hermetic sealing of equipment is preferable and a
ingress of liquid chemical agents or their required by the Project Team Leader. positive internal pressure is also desirable.
penetration into crevices, joints, slots, etc.
Defence standard 08-41 contains specific
information and advice in relation to methods for
CW hardening and materials of use. Advice on
any materials that are not contained in the
defence standard can be sought through the
Project Team Leader. The same is applicable
should the contractor seek specific design advice
or information.
3.11.20 Design criteria for the environmental The filtration requirements for environmental Defence standard 08-11 contains relevant
control system shall consider the installation of a control systems shall include the rejection of information and advice in relation to these

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filter or a method to completely shut off ambient air particles greater than 0.2 microns diameter. aspects.
and allow any contaminated air trapped in the
system to be purged prior to the subsequent use of The filtration requirements for environmental Any system that shuts off the ambient air supply
the cockpit/cabin system. control systems shall include CW agent vapour to the environmental control system and the crew
protection. The protection requirements shall be spaces shall be considerate of avionics cooling
defined by the Project Team Leader. and other requirements by design. The system
must remain shut off until the Aircraft reaches a
Design review shall be used to determine safe altitude above ground (preferably
compliance, with test / demonstration as automatically). This altitude will be advised by the
required by the Project Team Leader. Project Team Leader.
3.11.21 All-weather seals shall be designed Design review shall be used to determine Contamination of equipment compartments could
to also preclude the entrance of CW or BW compliance, with test / demonstration as be a threat to ground maintenance personnel.
contaminants into not only the cabin but also other required by the Project Team Leader.
equipments. Defence standards 08-11 and 08-41 provide
specific advice and information. Further advice
should be sought through the Project Team
Leader.
3.11.22 Suitable particulate filtration shall be The filtration system shall remove liquid droplets Specific information and advice may be obtained
used in air-cooled equipment. and particulate matter down to 1-micron from the Project Team Leader.
diameter.

Design review shall be used to determine

compliance, with test / demonstration as
required by the Project Team Leader.
3.11.23 Materials used in positions liable to The contractor shall demonstrate the ability of Defence Standards 08-11, 08-41 and 00-72
contamination shall: materials which are subjected to direct attack to contain specific advice and information.
withstand CW agents and decontaminants.
(a) Minimise the penetration/absorption Further advice and information may be obtained
of contaminants thereby reducing the Design review shall be used to determine from the Project Team Leader.
residual vapour hazard following compliance, with test / demonstration as
decontamination. required by the Project Team Leader.

(b) Not be damaged by liquid

contaminants or solid or liquid
decontamination agents to an extent that will

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impair the performance of the Aircraft or its
equipments within the limitation specified
under Clause 3.11.18
CHEMICAL AND BIOLOGICAL TESTING
3.11.24 The manufacturer shall supply to Materials lists shall be provided for specialist This is a specialised matter, and advice should be
Dstl, or other Authority designated by the Project review. Where materials cannot be cleared by sought through the Project Team Leader, in
Team Leader (PTL), a representative sample or analysis or similarity, test shall be required. relation to materials usage and test requirements.
samples, in a finished condition, including any
conformal coatings, of material that shall be Compliance will be based on the compatibility of Defence standard 08-41 contains a list of
defined by the Project Team Leader to require test. the material with CW agents being confirmed compatible and incompatible materials. However,
This applies when the manufacturer is using a (no damage or degradation outside of new materials and processes / finishes are in
material, process or finish which has not been acceptable levels that shall be agreed) continual development, requiring the Design
previously tested or which cannot be cleared for Organisation to seek specialist advice where this
use by analysis. is not contained in the referenced Defence
standard.
LASER REQUIREMENTS
GENERAL
3.11.25 Laser range finders and target designators have
and are being deployed throughout the world and
further development involving the use of exotic
laser materials has resulted in the diversity of
operating wavelengths such that optical/electro -
optical component damage can be inflicted at
battlefield ranges. Further development of the
laser in terms of increased range and power has
led to the emergence of the High Energy Laser
weapon, which when fully developed, will prove a
significant threat to airborne weapon systems.
Since the turn of the century, the widespread
availability of cheap portable and relatively
powerful hand held lasers has produced a
widespread dazzle threat.
3.11.26 The basic aim for laser survivability The factors which shall be considered during the
of military aircraft weapon and ground systems, design phase to reduce susceptibility to the
and installation shall be compliance with the effects of laser weapons are described in Leaflet

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application environment detailed in the relevant 24.
Specification.
DESIGN
3.11.27 The design objective for the laser Laser threat parameters involved in the laser
hardening of the aircraft, its installed equipment hardening design are given in Table 7.
and installed weapons systems shall be
compliance with the criteria detailed in the
appropriate Specification.
3.11.28 The principal design aim shall be to
maximise the probability that the threat effects
defined by the Project Team Leader (PTL) or in the
Aircraft Specification will not degrade the flying
qualities of the Aircraft below Level 2 as defined in
2.1.19
3.11.29 During the feasibility study stage the The results of the initial laser survivability
Design Organisation shall categorise those items feasibility studies shall be submitted to the
which are likely to be subjected to direct laser Design Organisation for acceptance of the
attack and shall subsequently undertake studies of estimated hardness levels of the proposed
the design to quantify its laser survivability level in design relative to the laser hardness criteria
relation to technical achievability and all the referred to in Clause 3.11.27 with special
penalties incurred. consideration to any penalties as mentioned in
Clause 3.11.28
3.11.30 Consideration shall be given to
minimising the damage to the Aircraft caused by
friendly forces laser weapons. It shall be assumed
that essential aircrew are adequately protected.

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3.12 AIRCREW EQUIPMENT ASSEMBLIES

INTRODUCTION

3.12.1 This Clause contains the requirements for the integration of the Aircrew Equipment Assemblies (AEA) with the aircraft.

3.12.2 The specification for each type of aircraft will list the AEA to be used. This is defined in FAP 108B-0001-1. The individual items are
described in the FAP108F series and in AP3456.

REQUIREMENT COMPLIANCE GUIDANCE

3.12.3 Failure of any or all the engines in It shall be demonstrated by test that the design Various items of AEA require power or service
flight shall result in the crew being able to operate of the installation is such that the correct power from the aircraft and include those listed in Table
those systems for the AEA which are essential to or service required by the AEA is delivered at 8.
the making of a descent and emergency landing of the point of connection of the AEA to the aircraft
the aircraft or for escape of the crew from the service.
aircraft.

Where the requirement exists to provide flying Personal NBC protection equipment shall be This is a specialised matter, and advice should be
personnel with NBC protective garments and / or shown to meet the requirement including those sought through the Project Team Leader.
assemblies, these shall fully and safely integrate associated with NBC agent protection.
with the AEA, the aircraft, its systems and
subsystems. This protection will be required during
ground operations and in flight, and potentially
after emergency escape, including from aircraft
fitted with ejection seats.

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3.13 BRAKE PARACHUTE INSTALLATIONS

INTRODUCTION

3.13.1 The requirements of this Clause apply to parachute installations in which a parachute or cluster of parachutes is provided to retard
the aeroplane during its landing run.

REQUIREMENT COMPLIANCE GUIDANCE

BASIC OPERATIONAL REQUIREMENTS
3.13.2 It shall be possible to stream the
parachute at any time during the landing run or
preferably just before touchdown so that it is fully
inflated, or in the case of a constant drag
parachute, in its operating condition, when the
wheels touch the ground.
3.13.3 The parachute shall be fully inflated,
or in the case of a constant drag parachute, in its
operating condition, as soon as possible, but within
5 seconds of operation of the pilot's control.
3.13.4 Parachute assemblies and cables
shall be capable of quick replacement.
3.13.5 Parachutes shall be attached as
close to the plane of symmetry of the aeroplane as
practicable and shall not foul any part of the
aeroplane during deployment. Streaming of the
parachute should not cause undue change of trim.
3.13.6 The parachute streaming and
jettisoning mechanism shall be unaffected both in
flight and on the ground by rain, snow and ice, and
shall be capable of operation after prolonged flying
at low temperature conditions.
3.13.7 It shall not be necessary to jettison
or detach any component or part of the aeroplane
during normal operation of the parachute(s)
3.13.8 The control for streaming and

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jettisoning shall be in accordance with Part 1,
Section 4, Table 19, Item 4.
SAFETY
3.13.9 It shall be possible for the pilot to
jettison the streamed parachute and cable with a
delay of not more than one second when landing
and taxying in winds up to 30 knots (55.6 km/h) in
any direction. Parachute instability shall be taken
into account by assuming the parachute to cone
through an included angle of 5°.
3.13.10 Parachutes and cables shall be
provided with a device such that if streaming
occurs before the pilot's control is operated, there
is no possibility of the parachute and cable
remaining attached to the aeroplane.
3.13.11 An over-riding jettisoning device or An effort should be made to reduce the scatter on
weak link, whose functioning is independent of the the strength of the overriding jettisoning device or
pilot, shall be provided to protect the aeroplane weak link to a minimum, and the aim should be to
structure from excessive parachute loads. This design the device so that its maximum estimated
device shall be designed to jettison the parachute failing load will be equal to the proof strength of
and cable at the proof load of that part of the the fuselage.
structure which carries the parachute load, within
the limits +0% -10%.
STRENGTH OF AEROPLANE STRUCTURE
NORMAL LANDING
3.13.12 The aeroplane structure shall have
proof and ultimate factors not less than 1.5 and 2.0
respectively on the combination of aerodynamic
loads and the average opening load (as defined in
Leaflet 58 Para 6) of the parachute when it is
streamed at the recommended approach speed
just before touchdown in a normal landing, at
landing weight, or at a speed 1.2 Vs1 if this is
greater, where Vs1 = stalling speed at landing

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weight with flaps and undercarriage down, engines
throttled.
EMERGENCY LANDING
3.13.13 The aeroplane structure shall have See Leaflet 25, Para 6.
proof and ultimate factors not less than 1.0 and 1.2
respectively on the combination of aerodynamic
loads and the maximum opening load when the
parachute is streamed just before touchdown in an
emergency landing at a speed of 1.2Vs or higher
emergency speed, if appropriate, where Vs =
stalling speed at the maximum weight for
emergency landing with high lift devices
inoperative, engines throttled.
STRENGTH OF PARACHUTE
NORMAL LANDING
3.13.14 The parachute canopy, rigging lines
and cables shall be capable of withstanding the
load applied in the normal landing case of Clause
3.13.12 for the number of times stated in the
Aeroplane Specification.
EMERGENCY LANDING
3.13.15 The parachute canopy, rigging lines
and cables shall have an ultimate factor not less
than 1.4 on the maximum loads applied during the
emergency landing of Clause 3.13.13
GROUND TESTS
3.13.16 Tests shall be made to demonstrate The type and number of tests, on the general See also Leaflet 25, Para 7.
that the streaming and jettisoning mechanism lines of those given in Leaflet 25, Para 7, shall
functions with a degree of reliability such that it is be agreed with AD AS DMPS.
unlikely to fail during the life of the aeroplane.
FLIGHT TESTS
3.13.17 Landing trials shall be performed on See also Leaflet 25, Para 7.3
a prototype or development aeroplane and the
parachute loads measured to establish that the

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installation is satisfactory and that the design loads
are not exceeded.
3.13.18 The landing tests shall be made
with the parachute streamed at the normal landing
speed of Clause 3.13.12 and at the highest speed
consistent with safety to establish as far as
possible compliance with the requirements of
Clause 3.13.13
3.13.19 The functioning of the jettisoning
mechanism shall be demonstrated at high and low
speeds by jettisoning in some of the tests early in
the landing run and in other tests at the end of the
landing run.

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3.14 INTEGRATION OF STORES

3.14.1 This Clause states the requirements for the structures and loads aspects of the integration of external stores, which in the context of
this Clause includes stores carried in internal weapons bays and Defensive Aids Suites (DAS) that eject expendable stores, onto British military
registered air systems. (Other requirements such as handling qualities are not addressed here). These requirements apply to all external stores
which may be carried, whether for release or long term carriage, and those carried in internal weapons bays, as well as DAS that eject expendable
stores. Advice on the compliance with these requirements is given in Leaflets 26 and 27. However, all compliance procedures and judgements are
subject to the approval of the relevant MoD PTL which in the case of stores integration is the aircraft PTL (herein referred to as ‘PTL’). In this
clause, the term ‘store’ refers to any launchers and / or adaptors required to carry and launch the store, including the attachment of these to the
pylon or installation attaching to aircraft structure (in the case of a DAS), as well as the store itself. The pylon / weapon carrier is assumed to be
part of the air system structure and the qualification of the pylon with the store is the responsibility of the Designer/Design Organisation
(Designer/DO) for that pylon, if different from the aircraft Designer/DO. Qualification of the DAS installation on the aircraft structure is the
responsibility of the Designer/Design Organisation. It is assumed that both the aircraft and store will have been designed by competent and
approved organisations, and in accordance with the relevant standards (e.g. Def-Stan 00-970 for the aircraft, Def-Stan 07-85 for the store).

REQUIREMENT COMPLIANCE GUIDANCE

APPROACH
3.14.2 The air system, comprising the The technical requirements of the system shall See Leaflet 26 – Integration Methodology.
aircraft sub-system and stores sub-systems, shall be defined by an overall systems integration
be qualified using an integrated systems approach. plan. A Qualification Programme Plan (QPP)
shall be produced for each sub-system, defining
the requirements for achieving airworthiness
clearance.
3.14.3 The interfaces and interactions The Integrating Authority (the PT, Designer/DO See Leaflet 26 – Integration Methodology.
between the structures system and all other sub- or other party with overall responsibility for the
systems within the air system shall be considered, integration exercise) shall provide a definition of
as well as the load interaction between the store the interface, and maintain the Interface Control
and the aircraft. Document (ICD), describing the mechanical,
electrical/software and environmental conditions
between the aircraft sub-system and the store
sub-system.
3.14.4 The carriage of the store on the air In the assessment of this, all significant loading See Leaflet 27 – Description of Design
system, in all locations to be cleared for service actions (both on the store imparted by the Considerations & Loading Actions.
use and in all combinations with other stores which aircraft, and on the aircraft imparted by the
are required for service clearance, shall not impact store) occurring as a result of normal and
the safe operation of the air system. emergency operations throughout the required

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envelope shall be considered. This assessment
should also include environmental loads and
effects, such as plume effects, vibration and
acoustic loading. Loads resulting from store
emergencies and failures, such as birdstrike or
failure to release when a rocket motor is fired,
shall also be addressed. The factors to be
considered to achieve structural clearance of a
new aircraft / store combination include, but are See Leaflet 27 for all items.
not limited to:

 Static loading (See Clause 3.1 of Def

Stan 00-970 Part 1, Section 3 - Structure)
 Fatigue loading (See Clause 3.2 of
Def Stan 00-970 Part 1, Section 3 -
Structure)
 Mass and the effect of variable store
centre of gravity (c of g).
 Environment
 Birdstrike
 Lightning Strike
 Plume assessment
 Hammershock loading – Unsteady
pressure pulses emanating from a jet
engine due to a surge condition and
passing through the duct and out of the
intake system (See Def Stan 00-970 Part
11, Section 3 E660);
 Scheduled and emergency jettison;
 Interaction with other stores, including
their release;
 Dynamic loading – carriage (Buffet)
and release.

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A flutter assessment must be carried out to
cover the new store / aircraft combination and all
agreed stores configurations.
Additionally for helicopters the effects of the
additional mass of the weapons on the airframe
dynamics shall be assessed. A ground
resonance clearance shall be provided and a
flight survey carried out to check that airframe
vibration levels are acceptable.
3.14.5 The impact on the performance of Any limitations on air system operation (e.g.
the air system due to carriage of the store shall be speed, fatigue life or ‘g’ limits) must be identified
agreed with the PTL. and agreed with the PTL.
3.14.6 The level of in-service monitoring of See Leaflet 27 – Description of Design
carriage of stores on the air system shall be Considerations & Loading Actions - Section 3.14
agreed with the operator, the Designer/DO for the and Def Stan 00-970 Part 1, Section 3 –
store / pylon / air system, and the PTL. Structure - Leaflet 38 – Service Monitoring.
QUALIFICATION PROGRAMME
3.14.7 Structural qualification and A programme of tests will normally be carried See Leaflet 27 – Description of Design
certification shall be achieved by means of a out to validate the structural analysis; the tests Considerations & Loading Actions – Section 2
staged, integrated programme of structural shall be sufficiently instrumented to enable and Def Stan 00-970 Part 1, Section 3 –
analysis, structural tests, and flight and ground validation. Structure - Clauses 3.1 and 3.2 as well as Leaflet
loads measurement, to be agreed with the PTL. Where representative testing is not carried out, 35 – Safe–Life Substantiation.
additional safety factors shall be applied. This
shall also be the case where fatigue loads are
unmonitored in service.
FLIGHT AND GROUND LOADS MEASUREMENT
3.14.8 Measurements to derive loads, It will be necessary to ensure that any It may be necessary to employ instrumented wind
temperatures and other relevant parameters during instrumented stores are sufficiently tunnel models. It may be acceptable to obtain the
flight and on the ground, to an extent to be agreed representative (e.g. in terms of mass required data by means of read-across to similar
with the PTL, shall be made on one or more distribution, inertia, stiffness) of a service air systems / stores combinations. However,
representative air systems, as required by the standard store. Measurements shall be made at evidence of the similarity shall be provided to
appropriate Clauses of this publication. sufficient stations to establish the loads in the support such a read-across.
relevant parts of the structure to a level of The number and position of stations will depend
accuracy agreed with the PTL. on the project concerned, on the quantities being
measured and the practical aspects of

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accessibility
For example, careful positioning of
instrumentation will be needed to ensure that the
normal modes of vibration are not affected
unduly, whilst still being measured accurately.
See Def Stan 00-970 Part 1, Section 3 –
Structure - Clauses 3.1.23, 3.1.24, 3.1.25 and
3.1.26 as well as Leaflet 27 – Description of
Design Considerations & Loading Actions –
Sections 3.1 and 3.2
3.14.9 Measurements made to derive The magnitude, distribution and time history of See Def Stan 00-970 Part 1, Section 3 –
loads, temperatures and other relevant parameters the loads, temperature and vibration levels shall Structure - Clause 3.1.21
during flight and on the ground shall be used to be determined in the agreed flight conditions
validate the design assumptions and to confirm the and appropriate air system configurations. In
structural clearance of the air system plus store selecting the conditions, the cases to be
combination. considered shall include but not be limited to:

(a) symmetric manoeuvres,

(b) asymmetric manoeuvres,

(c) continuous turbulence, where

practicable,

(d) special flight manoeuvres, such

as in-flight refuelling, or supply dropping,
and any other manoeuvres that are likely
to be performed during the specified use
of the air system, for example, terrain
following in still and turbulent air,

(e) the take-off and landing cases of

Clause 4 and other specified ground
manoeuvres,

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(f) stores ejection loads,

(g) internal weapons bay operation,

acoustic noise and vibration environment,
and any other static or fatigue loading
cases arising from the particular role of the
air system.

In all manoeuvre cases, the loads and their

distributions and time histories shall be
determined over a range of severity of the
manoeuvres up to the design conditions or any
lower limit set by safety or other considerations.
3.14.10 Thorough inspection of the structure Structural inspection methods shall be sufficient See Def Stan 00-970 Part 1, Section 3 –
shall be made after flight tests. to determine that there are no signs of structural Structure - Clause 3.1.22
damage and/or deformation.

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3.15. Not Issued

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3.16 DEFENSIVE AIDS SYSTEMS (DAS)

INTRODUCTION

3.16.1 Defensive Aids Systems are intended to reduce the susceptibility of the aircraft and its crew to attack by conventional guided and
unguided weapons. This is achieved by improving situational awareness, enabling avoidance of threat, and by providing timely warning of attack
such that countermeasure responses can be deployed effectively.

3.16.2 Defensive Aids Systems vary considerably from aircraft to aircraft in terms of the subsystems fitted, their integration and control. Most
DAS will comprise some or several of the following items:

(a) Radar Warning Receiver (RWR) and/or Electronic Support Measures (ESM)
(b) Missile Warning System (MWS)1 {encompassing Missile Approach Warning System and Missile Launch and Approach Warning
Systems}
(c) Hostile Fire Indication / location (HFI)1
(d) Laser Warning System (LWS)
(e) Countermeasure Dispensers or Countermeasure Dispensing System (CMDS)
(f) InfraRed Jammer (IRJ)
(g) Directed InfraRed Countermeasure system (DIRCM)
(h) On board Radio Frequency Jammer (OBJ)
(i) Towed Radar Decoy (TRD)
(j) Expendable Active Decoys (EAD)
(k) Integration via a DAS Controller (DASC)
(l) Crew Interfaces
3.16.3 This Clause sets out the design requirements for the installation of DAS equipment. The requirements are applicable to all types of
aircraft, except those designated for trials or other research activities.

1
MWS incorporating HFI is sometimes referred to as a Threat Warning Systems (TWS); such a system may also incorporate LWS and RWR
functionality.

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GENERAL
3.16.4 The general requirements of Part 1,
Section 6 on avionic equipment installations apply
to DAS.
3.16.5 Installation of DAS elements shall The performance parameters of each DAS
primarily be driven by performance considerations. element and of the system as a whole will be
specified in the SRD to provide effective
protection against threats.
The specified performance of the DAS will be
driven by the aircraft signature. This must be
considered if signatures are changed significantly
during the life of the aircraft.
3.16.6 All elements of the DAS, and the A schedule of testing shall be presented in the In order to facilitate acceptance testing, recording
integrated DAS as a whole, shall be characterised form of an Integrated Test, Evaluation and equipment will be required. The design of the
and tested. Installed performance shall be proven Acceptance Plan (ITEAP), and agreed with the DAS and each of its elements must include
in compliance with the requirements of the aircraft Project Team Leader (PTL). Flight tests shall be provision for flight data recording. Such recording
SRD. made to demonstrate compliance. may also be used in service.
Recognising that only a limited subset of all
responses to threats and threatening
circumstances can be proven through flight trial, a
comprehensive and validated model /prediction
toolset2 of the DAS3 should be procured by the
PT, and maintained though the life of the aircraft.
3.16.7 The DAS shall be re-programmable Threats will evolve and new threats appear
to function effectively against new or upgraded through the life of the aircraft.
threats. A comprehensive and validated model2 of the
DAS3 should be procured by the PT, and
maintained though the life of the aircraft. This

2
See also “Modelling, Simulation and Synthetic Environments Policy, information and guidance on the Modelling, Simulation and Synthetic
Environments aspects of UK MOD Defence Acquisition”, version 1.0.0 - March 2008
3
And other relevant survivability elements such as aircraft signatures and flight performance.

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model can be used to predict performance
against new or upgraded threats, and to develop
countermeasures.
3.16.8 DAS sensor and effector aperture Permissible increase in RCS or other signatures This requirement is of particular importance to low
installations shall not adversely affect the radar will be defined in the SRD, and shall be – observable aircraft.
cross section (RCS) or other key signature demonstrated by measurement on aircraft.
elements of the aircraft.
3.16.9 The integrated DAS as a whole shall Design must, for example, preclude any DAS
be designed such that no single fault or failure failure commanding or permitting unsafe firing of
shall adversely affect the safety of the DAS, nor stores, unsafe RF, laser or other EM
operation of any other aircraft system, nor the transmissions in air or on the ground. DAS units
safety of the aircraft as a whole. must not be able to overheat, emit fumes, catch
fire etc, nor cause failures in other aircraft
systems including power supply.
See also Part 1, Section 6.
3.16.10 The integrated DAS as a whole shall Consideration should be given to but not limited
be designed such that no single fault or failure to;
shall adversely affect the operation of other
elements of the DAS. - Dual redundancy in the DAS system.

- Provide isolation between elements of the

DAS.

- Provide alternative means of operating in

the event of any failure or damage.

- Not to allow a fault in one databus or direct

link to prevent the operation of the DAS
system (but possibly accepting slightly
degraded performance).
3.16.11 The integrated DAS shall comply
with the requirements for Static Strength and
Deformation, detailed in Part 1, Section 3.1
3.16.12 The integrated DAS shall comply
with the requirements for Fatigue Strength,

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detailed in Part 1, Section 3.2
3.16.13 The integrated DAS shall comply
with the requirements for Aero-elasticity, Flutter
and Vibration, detailed in Part 1, Section 4.8
RADAR WARNING RECEIVER (RWR) AND/OR ELECTRONIC SUPPORT MEASURES (ESM)
3.16.14 Receiving apertures (antennas) Airframe obscuration and reflections shall be Sensors may be adversely affected by sources of
shall be mounted with consideration of airframe minimised. heat, sources of RF energy (e.g. on-board
obscuration, airframe reflections, and meeting the Maps of receiver sensitivity against angle transmitters), and by close firing of weapons or
system field of view required in the SRD. (azimuth and elevation), and waveband shall be stores.
validated by measurement on aircraft. See also Part 1, Section 6.1
See Def Stan 59-411 for EMC test requirements.
Note that Part 7 Section 7 Leaflet 707/3m does
not necessarily apply.
3.16.15 Receiving apertures (antennas) See also Part 1, Section 1, Clause 1.1.15.
shall be mounted with consideration of airframe Precise harmonisation4 may be of high
structural flexibility. Provision shall be made for importance in direction – finding ESM. In some
bore-sighting and harmonisation4 in order to meet systems separate sensors may cover overlapping
the angle of arrival accuracy requirements of the regions of the RF spectrum. In such cases testing
SRD. should prove performance of the system against
emitters in such overlap bands.
3.16.16 Installed performance parameters of Parameters will be defined in the SRD. The The RWR/ESM is a sensitive RF receiving system
the RWR/ESM shall be proven. ability of the RWR/ESM to detect, characterise which is required to operate in the EMI
and prioritise threat emitters shall be proven. environment of the aircraft. Consideration must be
The ability of the RWR/ESM to link multiple given to the effects of on board transmitters; the
emissions from a single threat shall be design of the RWR/ESM must allow for
demonstrated. False alarm rate and multiple satisfactory receive performance in this
declaration rate shall be demonstrated. Flight environment.
testing is required.
MISSILE WARNING SYSTEM (MWS)
3.16.17 Sensor apertures shall be mounted Airframe obscuration and reflections shall be Some types of MWS may be adversely affected
with consideration of airframe obscuration, minimised. by engine exhaust plumes, sources of heat, oil

4
Harmonisation here refers to the relative boresighting of separated sensors (or effectors) both statically and dynamically in the face of airframe
structural flexibility in flight.

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airframe reflections, contaminants, and meeting Coverage maps shall be provided by spray, sources of light (e.g. landing lights) or other
the system field of view required in the SRD. measurement on aircraft. EM energy, and by close firing of weapons or
stores. Some types may be particularly sensitive
to RF interference, and to the aircraft’s vibration
environment.
3.16.18 Sensor apertures shall be mounted Precise harmonisation is often of great
with consideration of airframe structural flexibility. importance to MWS operation, as is proving of
Provision shall be made for bore-sighting and performance against threat objects which fall in
harmonisation4 in order to meet the angle of arrival overlap regions, or pass from one sensor field of
accuracy requirements of the SRD. view to another.
See also Part 1, Section 1, Clause 1.1.15
3.16.19 Installed performance parameters of Parameters will be defined in the SRD. The
the MWS shall be proven. ability of the MWS to detect and characterise
and prioritise threats shall be proven.
False alarm rate and multiple declaration rate
shall be demonstrated.
Flight testing is required.
HOSTILE FIRE INDICATION (HFI)
3.16.20 Sensor apertures shall be mounted Airframe obscuration and reflections shall be HFI may be adversely affected by engine exhaust
with consideration of airframe obscuration, minimised. plumes, sources of heat, oil spray, sources of light
airframe reflections, contaminants, and meeting Coverage maps shall be provided by (e.g. landing lights) or other EM energy (e.g. on-
the system field of view required in the SRD. measurement on aircraft. board lasers), and by close firing of weapons or
stores.
3.16.21 Sensor apertures shall be mounted Harmonisation4 may be important to HFI
with consideration of airframe structural flexibility. operation, as is proving of performance against
Provision shall be made for bore-sighting and threat objects which fall in overlap regions, or
harmonisation4 in order to meet the angle of arrival pass from one sensor field of view to another.
accuracy requirements of the SRD. See also Part 1, Section 1, Clause 1.1.15
3.16.22 Installed performance parameters of Parameters will be defined in the SRD. The
the HFI shall be proven. ability of the HFI to detect and characterise and
prioritise threats shall be proven.
False alarm rate and multiple declaration rate
shall be demonstrated. Flight testing is required.
LASER WARNING SYSTEMS (LWS)

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3.16.23 Sensor apertures shall be mounted Airframe obscuration and reflections shall be LWS may be adversely affected by engine
with consideration of airframe obscuration, minimised. exhaust plumes, sources of heat, oil spray,
airframe reflections, contaminants, and meeting Coverage maps shall be provided by sources of light (e.g. landing lights) or other EM
the system field of view required in the SRD. measurement on aircraft. energy (e.g. on-board lasers), and by close firing
of weapons or stores.

3.16.24 Sensor apertures shall be mounted Performance must be demonstrated against

with consideration of airframe structural flexibility. threat emitters which fall in sensor overlap
Provision shall be made for bore-sighting and regions5, or pass from one sensor field of view to
harmonisation4 in order to meet the angle of arrival another.
accuracy requirements of the SRD. See also Part 1, Section 1, Clause 1.1.15
3.16.25 Installed performance parameters of Parameters will be defined in the SRD. The
the LWS shall be proven. ability of the LWS to detect and characterise and
prioritise threat emitters shall be proven. The
ability of the LWS to link multiple emissions from
a single threat shall be demonstrated.
False alarm rate and multiple declaration rate
shall be demonstrated. Flight testing is required.
COUNTERMEASURE DISPENSING SYSTEM (CMDS)
3.16.26 Dispenser locations and directions Modelling and simulation are required to prove Stores fired from the dispensing system will be
of fire shall primarily be determined by the the effectiveness of dispenser locations and specified in terms of performance against threats.
effectiveness of the countermeasure response. firing directions. The decoy signatures provided by stores will be
Flight testing is required to validate modelling related to the signatures of the protected aircraft6,
results. and the nature of the threat.
3.16.27 Fixed or movable dispenser The effects of the installation on the aircraft
installations shall not prejudice aircraft structure, aerodynamics and the ease of servicing and
other aircraft systems, or aircraft operation. replenishment of dispensers should also be
carefully considered.
See also Clauses 3.16.11, 3.16.12 and 3.16.13
3.16.28 The firing of stores from dispenser Compliance shall be demonstrated on a working Maximum reaction loads due to stores firing must
installations shall not prejudice aircraft structure, rig of the installation or by ground firing trials, be considered. Stores natures are likely to
other aircraft systems, or aircraft operation. and by air firing trials. change though the lifetime of the aircraft.

5
Geometrically, or in terms of waveband, where separate sensors cover overlapping spectral regions
6
Hence the importance of measuring and understanding signatures, and maintaining such understanding through the life of the aircraft.

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Any adverse effects, however brief or minor, Guidance must be sought from the Project Team
shall be quantified. Leader (PTL) as to design margins required.
The trajectories of dispensed stores must be
considered with regard to airframe strike, strike of
rotors or propeller blades, strike of towed or
underslung loads, engine ingestion, engine surge
or flameout, and contamination of other apertures
or stores. Worst case errors in trajectory,
including those due to uncommanded or partial
firing of eject charges, must be considered.
Stores firing should not damage or contaminate
sensor apertures. Effects such as dazzling of
sensors shall be minimised.
Safety of parachutists, safety during pallet drop
and safety during aerial refuelling must be
considered.
See also Clauses 3.16.11, 3.16.12 and 3.16.13
3.16.29 Countermeasure stores and/or their See also Part 1, Section 6.10, and Part 1, Section
ejection charges are regarded as explosive stores, 1, Clause 1.1.28
and a CMDS as an armament system, thus Part
13, Sections 3.1, 3.2 and 3.4 apply.
3.16.30 It shall not be possible to mis-load The use of idents, dowels, and/or component
dispensers. shapes is required to prevent loading of incorrect
stores into dispensers and to prevent mis-
orientation.
3.16.31 The dispensing system shall keep A misfire procedure must be established to
count of stores loaded and expended, and of provide clear instruction to ground crews on safe
misfires. A misfire procedure shall be established. unloading of misfired stores.
3.16.32 Safety of aircrew and ground Appropriate procedures must be defined. Ground handling will involve explosive stores.
personnel must be ensured. Appropriate tools must be provided.
Also consider crew entry and exit from the
platform, including emergency escape.
3.16.33 Installed performance parameters of Parameters will be defined in the SRD. Flight See also Part 1, Section 1, Clause 1.1.15
the CMDS shall be proven. testing is required. The ability of the CMDS to
dispense a commanded sequence of stores and

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store types from designated dispensers and
within specified timing parameters shall be
demonstrated.
INFRARED7 JAMMER (IRJ)
3.16.34 IRJ transmitter apertures shall be Airframe obscuration and reflections shall be Apertures may be adversely affected by engine
mounted with consideration of airframe minimised. exhaust plumes, sources of heat, oil spray, and
obscuration, airframe reflections, contaminants, Coverage maps shall be provided by by close firing of weapons or stores.
and meeting the system field of effect required in measurement on aircraft.
the SRD.
3.16.35 Installed performance parameters of Parameters will be defined in the SRD. Flight
the IRJ shall be proven. testing is required.
3.16.36 IRJ transmissions shall not prejudice Any adverse effects, however brief or minor, Effects such as dazzling of sensors shall be
other DAS elements, other aircraft systems, or shall be quantified. minimised.
aircraft operation. Safety of parachutists, safety during pallet drop
and safety during aerial refuelling must be
considered.
3.16.37 Safety of aircrew and ground Appropriate procedures must be defined. High power lamp sources can cause injury,
personnel must be ensured. including eye injury. An IRJ may run very hot and
remain hot for some time after shut-down.
Consider safety of maintainers etc., also consider
crew entry and exit from the platform, including
emergency escape.
DIRECTED INFRARED COUNTERMEASURE SYSTEM (DIRCM)8
3.16.38 DIRCM sensor and transmitter Airframe obscuration and reflections shall be Apertures may be adversely affected by engine
apertures shall be mounted with consideration of minimised. exhaust plumes, sources of heat, oil spray,
airframe obscuration, airframe reflections, Coverage maps shall be provided by sources of light (e.g. landing lights) or other EM
contaminants, and meeting the system field of view measurement on aircraft. energy (e.g. on-board lasers), and by close firing
/ field of effect required in the SRD. of weapons or stores.
3.16.39 DIRCM apertures shall be mounted Harmonisation4 is typically of importance to
with consideration of airframe structural flexibility. DIRCM operation, as is proving of performance
Provision shall be made for bore-sighting and against threat objects which fall in overlap

7
Or other EO-band
8
Including visible or other optical directed countermeasures.

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harmonisation in order to meet the angle of arrival regions, or pass from one sensor/effector field of
accuracy requirements of the SRD. view/effect to another.
See also Part 1, Section 1, Clause 1.1.15
3.16.40 Installed performance parameters of Parameters will be defined in the SRD. Flight
the DIRCM shall be proven. testing is required. The ability of the DIRCM to
detect, characterise, prioritise and jam threats
shall be proven.
If DIRCM can directly or indirectly command the
firing of expendable stores, this function must be
tested and proven.
False alarm rate and multiple declaration rate
shall be demonstrated.
3.16.41 Requirement Removed by NPA
2013-005
3.16.42 DIRCM transmissions shall not Any adverse effects, however brief or minor, Effects such as dazzling of sensors shall be
prejudice other DAS elements, other aircraft shall be quantified. minimised.
systems, or aircraft operation. Safety of parachutists, safety during pallet drop
and safety during aerial refuelling must be
considered.
3.16.43 Safety of aircrew and ground Appropriate procedures must be defined. High power lamp and laser sources can cause
personnel must be ensured. injury. Uncommanded operation on the ground
must be prevented. The eye-hazard range of such
systems can be very large, and must be
understood by maintainers and testers.
Consider safety of maintainers etc., also consider
crew entry and exit from the platform, including
emergency escape. See JSP 390 for Military
Laser Safety requirements.
ON BOARD RADIO FREQUENCY JAMMER (OBJ)
3.16.44 OBJ sensor and transmitter Airframe obscuration and reflections shall be Apertures may be adversely affected by sources
apertures shall be mounted with consideration of minimised. of heat, sources of RF energy (e.g. on-board
airframe obscuration, airframe reflections, Maps of receiver sensitivity, and of transmitted transmitters), and by close firing of weapons or
contaminants, and meeting the system field of view power against angle (azimuth and elevation), stores.
/ field of effect required in the SRD. and waveband shall be provided by
measurement on aircraft.

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3.16.45 Apertures (antennas) shall be
mounted with consideration of airframe structural
flexibility. Provision shall be made for bore-sighting
and harmonisation in order to meet the angle of
arrival accuracy requirements of the SRD.
3.16.46 Installed performance parameters of Parameters will be defined in the SRD. Flight Some types of OBJ incorporate their own receiver
the OBJ shall be proven. testing is required. The ability of the OBJ to elements; others will rely upon an RWR/ESM to
detect and characterise and prioritise threat detect, characterise and prioritise threat emitters
emitters shall be proven. The ability of the OBJ and initiate a jamming response.
to link multiple emissions from a single threat In either case, interoperability of receiver and
shall be demonstrated. The ability of the OBJ to transmitter must be demonstrated.
transmit the commanded jamming waveform(s)
shall be demonstrated. False alarm rate and
multiple declaration rate shall be demonstrated.
3.16.47 OBJ transmissions shall not Any adverse effects, however brief or minor, See Def Stan 59-411 for EMC test requirements.
prejudice other DAS elements, other aircraft shall be quantified. See also Part 1, Section 6.10.
systems, or aircraft operation. Safety of parachutists, safety during pallet drop
and safety during aerial refuelling must be
considered.
3.16.48 Safety of ground personnel must be Appropriate procedures must be defined. High power RF sources can cause injury.
ensured. Uncommanded operation on the ground must be
prevented.
Consider safety of maintainers etc., also consider
crew entry and exit from the platform, including
emergency escape.
TOWED RADAR DECOY (TRD)
3.16.49 Receiving apertures (antennas) Airframe obscuration and reflections shall be TRDs may incorporate receiving elements as part
shall be mounted with consideration of airframe minimised. of the towed body, other types will use dedicated
obscuration, airframe reflections, and meeting the Maps of receiver sensitivity against angle receivers (which may be shared with an OBJ),
system field of view required in the SRD. (azimuth and elevation), and waveband shall be others will rely upon an RWR/ESM to detect and
provided by measurement on aircraft. characterise and prioritise threat emitters and
initiate a jamming response.
Sensors may be adversely affected by sources of
heat, sources of RF energy (e.g. on-board
transmitters), and by close firing of weapons or

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stores.
3.16.50 Apertures (antennas) shall be
mounted with consideration of airframe structural
flexibility. Provision shall be made for bore-sighting
and harmonisation4 in order to meet the angle of
arrival accuracy requirements of the SRD.
3.16.51 A TRD is regarded as a towed body, See also Part 1, Section 6, Clauses 6.1.20 to
thus the requirements of Section 3.8 apply. 6.1.21
3.16.52 Installed performance parameters of Parameters will be defined in the SRD. Flight TRDs may incorporate receiving elements as part
the TRD shall be proven. testing is required. The ability of the TRD to of the towed body, other types will use dedicated
detect and characterise and prioritise threat receivers (which may be shared with an OBJ,
emitters shall be proven. The ability of the TRD other will rely upon an RWR/ESM to detect and
to link multiple emissions from a single threat characterise and prioritise threat emitters and
shall be demonstrated. The ability of the TRD to initiate a jamming response.
transmit the commanded jamming waveform(s) In either case, interoperability of receiver and
shall be demonstrated. False alarm rate and transmitter must be demonstrated.
multiple declaration rate shall be demonstrated.
Streaming, recovery (if applicable), flight stability
and jettison of the towed body shall be
demonstrated (see Clause 3.8.70)
3.16.53 TRD deployment, transmissions, Any adverse effects, however brief or minor, See Def Stan 59-411 for EMC test requirements.
recovery and/or jettison shall not prejudice other shall be quantified. See also Part 1, Section 6.10
DAS elements, other aircraft systems, or aircraft Safety of parachutists, safety during pallet drop
operation. and safety during aerial refuelling must be
considered.
3.16.54 The TRD system shall keep count of
TRD bodies loaded, streamed, recovered and
jettisoned, and of any which fail to deploy.
3.16.55 Safety of ground personnel must be Appropriate procedures must be defined. High power RF sources can cause injury.
ensured. Uncommanded operation on the ground must be
prevented.
Consider safety of maintainers etc., also consider
crew entry and exit from the platform, including
emergency escape.
EXPENDABLE ACTIVE DECOYS (EAD)

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3.16.56 Clauses 3.16.26 to 3.16.33 apply. An EAD is regarded as an expendable store,
which may incorporate explosive charges.
3.16.57 Installed performance parameters of Parameters will be defined in the SRD. Flight
the EAD shall be proven. testing is required. The ability of the EAD to
transmit the commanded jamming waveform(s)
shall be demonstrated. Flight parameters of the
EAD shall be demonstrated.
INTEGRATION VIA A DAS CONTROLLER (DASC)
3.16.58 Integrated threat detection shall be Parameters will be defined in the SRD. Flight This refers to the association of several sensed
demonstrated. testing is required. The ability of the DASC to entities (e.g. radar tracker + radar illuminator +
accept threat detections, characterise and laser rangefinder + missile warning) into a single
prioritise threats shall be proven. threat system where appropriate, depending on
accuracy and resolution of the estimated locality
of threat. This is also referred to as “weapon-
linking”.
The system will need to differentiate between
different operating modes of an identified threat
system.
3.16.59 Integrated command of Parameters will be defined in the SRD. The This refers both to the countering of several
countermeasure shall be demonstrated. ability of the DASC to receive and process all modes of a single threat (e.g. requiring RF
relevant information and to command expendables, EO expendables and RF jamming),
countermeasure responses shall be proven. and to the countering of multiple threats of diverse
Safe and effective interoperation of nature.
countermeasures shall be proven. The DASC may need to receive aircraft data
Flight testing is required. inputs such as attitude, altitude, and wind velocity.
Some paths of countermeasure command may
bypass the DASC.
The DAS may be required to command other
aircraft systems, for example radio frequency
transmitters, in response to certain threats9, if so
specified in the SRD.
3.16.60 The system shall allow manual The SRD will define the detail of each mode of For example: Manual - the crew member(s) select
selection of the mode of response to declared operation. and initiate countermeasures to threats they are

9
E.g. threats which may home on the aircraft’s own RF / radar transmissions

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threats as follows: aware of (by whatever means) or as a
preventative measure;
Manual Automatic - the system selects and initiates
countermeasures, excluding manoeuvre,
Automatic according to its own situational awareness and
Mission Threat Data;
Semi-automatic Semi-Automatic - the system selects
countermeasures according to its own situational
awareness, but the aircrew must authorise
initiation.
3.16.61 Any requirement for automated The SRD will define any automated manoeuvre Automated manoeuvre is not generally permitted
manoeuvre will be subject to thorough flight responses to threat. in manned aircraft, but may be an option in
testing. remotely piloted platforms.
CREW INTERFACES
3.16.62 The DAS shall provide adequate See also Part 1, Section 4, leaflet 64
crew displays, controls and audio warnings for and Def-Stan 00-250
situational awareness, responses to threat and of
the condition of the system. Consideration must be given to the interoperation
and prioritisation of DAS – derived voice or audio
warnings, against those from other avionic
systems.
This requirement applied to manned aircraft.
Guidance must be sought from the PT leader as
to the applicability to remotely piloted air vehicles
(e.g. in ground-stations), and to autonomous
remotely piloted air vehicles.

Clause Subject
1.1 Alignment of directionally sensitive weapons
1.1 Jettisoning of stores
7.1 and 7.4 Operation in various climatic conditions

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7.3 Operational colouring and markings

4.15 Colour of illuminated indicators
4.19 Cockpit controls and instruments
Design criteria for aircraft equipment, systems and
installations
3.11 Picketing
4.13 Ground clearance
6.6 Electrical installations
4.27 Bonding and screening
13.3.3 Gun installations
6.4 Magnetic Compass Installations
13.3.11 Protection from the effects of nuclear explosion
Section 4 Leaflet 55 Doors and locks
Leaflet 7 The effect of gun firing on turbine engines
Leaflet 5 Armament Installations Fixed Guns

TABLE 1 - LIST OF OTHER IMPORTANT REQUIREMENTS - (Clause 3.1)

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Clause Subject
Section 1 Alignment of directionally sensitive equipment
1.1 and weapons
Section 4 Control column
4.19
Section 4 Armament controls
4.19
Section 4 Protective Treatment
4.3
3.1 Armament installations
Section 4 Fumes and vapour seals
4.26
Section 6 Magnetic compass installations
6.4
3.11 Protection from the effects of nuclear
Explosion
Section 4 Routine Servicing and Turn Round
4.4
3.3 Gun Installations
Leaflet 5 Armament installations Fixed guns

TABLE 2 - (Clause 3.2)

LIST OF OTHER IMPORTANT REQUIREMENTS

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CLAUSE SUBJECT
1.1 Vibration
1.1 Jettison of Stores
7.1 and 7.4 Operation in Various Climatic Conditions
4.19 Jettison of Stores
4.13 Ground clearance
2.1 Flight Phase Categories
2.1 Levels of Flying Qualities
13.3.1 Armament Installations
13.3.4 Installation of Explosive Devices
6.10 Electromagnetic Compatibility of Safety Critical Systems

TABLE 3 - LIST OF OTHER REQUIREMENTS (Clause 3.8)

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PROBABILITY DIRECTIONAL
FORM OF OCCURENCE QUALIFIERS
THREAT SOURCE (Hit density, impact BY AEROPLANE
area, velocity, mass). CLASS (NOTE 2) GENERAL
I II III IV Elevation Azimuth EFFECT
 (deg)  (deg)
a INERT GUNS 0.2/m2; over aeroplane; 0. 0. 0. 0. +5 -15 +6 -60 PENETRATION
BULLETS 600m/s; 7g 40 30 05 05 0
b INERT MISSILES 20/m2; over 2 m x 7m; 0. 0. 0. 0. (See Note 3) PENETRATION
FRAGMENTS 2000m/s; 5g 05 15 40 60

Group 1 c SHELL GUNS 0.05/m2; over 0. 0. 0. 0. +5 -10 +6 -60 DISRUPTION

Threats aeroplane; 23mm) 40 40 40 15 0
(Conventional) (Note 1)

d WARHEAD MISSILES Fragments and blast 0. 0. 0. 0. +1 -10 +1 -170 DISRUPTION

(Note 1) 15 15 15 20 0 70
e INCENDIARY GUNS (23mm) one part in four - - - - - - FIRE
BULLETS of threats
a and c

TOTAL FOR GROUP 1 THREATS 1. 1. 1. 1.

0 0 0 0
HEAT
f - DIRECTED LASER See Aeroplane 0. 0. 0. 0. +0 -90 +1 -180 PENETRATION
Specification 0 2 3 2 80
g - GENERAL NUCLEAR " " " ) DEGRADATION
Group 2 h NEUTRON NUCLEAR " " " ) DISRUPTION
Threats FLUX
(Non- i GAMMA NUCLEAR " " " ) 0. 0. 0. 0. DEGRADATION
Conventional) RADIATION 8 4 5 3
j ELECTROMA NUCLEAR " " " ) DISRUPTION
GNETIC
k BLAST NUCLEAR " " " ) DISRUPTION
l BIOLOGICAL - ) See NGSAR NO. 562 0. 0. 0. 0. INCAPACITATION

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for crew and 1 2 1 25

m CHEMICAL - ) personnel protection 0. 0. 0. 0. INCAPACITATION
1 2 1 25

TOTAL FOR GROUP 2 THREATS 1. 1. 1. 1.

0 0 0 0

TABLE 4 - TABLE OF DEFINED THREAT EFFECTS (Clause 3.9)

NOTES:
1. See RAE Report No. 79123.
2. See Leaflet 24 and Part 0 for definitions.
3. From any direction perpendicular to the axis of the aeroplane at any point within the length of the aeroplane.

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THREAT SOURCE FORM; i.e. density/dose, area PROBABILITY OF DIRECTIONAL QUALIFIERS GENERAL
exposed, power density, etc OCCURRENCE BY EFFECT
AEROPLANE CLASS
(NOTE) 1

I II III IV ELEVATION(o) AZIMUTH(o)

Directed EMR Laser See Aeroplane Specification 0.0 0.2 0.3 0.2 +0 - 90 + 180 - 180 Degradation
Energy
Thermal Nuclear “ “ “ ) Degradation
Neutron Flare Nuclear “ “ “ ) Disruption
Gamma Radiation Nuclear “ “ “ ) 0.8 0.4 0.5 0.3 Degradation
Electromagnetic Nuclear “ “ “ ) Disruption
Blast Nuclear “ “ “ ) Disruption
Biological - Liquid, Aerosol and Vapour 0.1 0.2 0.1 0.25 Note ( Incapacitation /
Death / Loss of
Aircraft
Chemical - See NGASR No 562 for crew and 0.1 0.2 0.1 0.25 2 ( Incapacitation /
personnel protection. Death / Loss of
Aircraft
TOTAL FOR GROUP 1.0 1.0 1.0 1.0
2 THREATS

TABLE 5 - DEFINED THREAT EFFECTS (Clause 3.11)

Notes:
1. See Leaflet 24 and Part 0 for definitions.
2. See AP 71328 series. (AC 71328)

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AIR BLAST THERMAL RADIATION

REF PARAMETER SYMBOL VALUE REF PARAMETER SYMBOL VALUE
1a Incident Peak Static p 2a Total Thermal Fluence Q
Overpressure
1b Peak Dynamic q 2b Time to maximum irradiance t max
Pressure
1c Dynamic Pressure Iq 2c Maximum irradiance Q max
Impulse
1d Arrival Time ta 2d 80% energy delivery time S
1e Static Overpressure tp 2e Pulse width at half Q max S
Duration
1f Static Overpressure Ip
Impulse
1g Peak Underpressure p neg
INITIAL NUCLEAR EMP
RADIATION
REF PARAMETER SYMBOL VALUE ENDOATMOSPHERIC EMP EXOATMOSPHERIC EMP
3a Total dose (tissue) Dt REF PARAMETER ELECTRIC MAGNETIC REF PARAMETER ELECTRIC MAGNETIC
FIELD FIELD FIELD FIELD
3b Maximum neutron Dn 4a Peak Kv/m A/m 5a Peak Kv/m A/m
contribution Amplitude A Amplitude A
3c Maximum gamma Dy 4b t1 (0.9A) ns ns 5b t1 (0.9A) ns ns
contribution
3d Maximum combined Di 4c t2 (A) ns ns 5c t2 (A) ns ns
ionising dose
3c Maximum neutron Fn 4d t3 (0.83A) ns ns 5d t3 (0.5A) ns ns
Fluence
3f Peak gamma dose . 4e t4 (o.65A) ns ns 5e t4 (0.1A) ns ns

rate
4f t5 (0.1A) ns ns

TABLE 6 - AEROPLANE NUCLEAR HARDENING - TYPICAL PARAMETERS (Clause 3.11)

Notes:

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1. The numerical values associated with Air blast, Thermal Radiation, Initial Nuclear Radiation and EMP parameters will be specified
or agreed with the Project Team Leader (PTL) (see Clauses 3.11.6 and 3.11.9).
2. The number of exposures to the above parameters will be defined by the Project Team Leader (PTL).
3. The above parameters are to be considered separately except where combinations are considered significant to the design. In this
case the Combinations to be considered together with relevant separation timescales will be defined by the Project Team Leader
(PTL).
4. The Kill Levels to be associated with the threat values given in the above table will be that of Mission Completion (see Leaflet 22).

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LASER OPERATING MODE - PULSED LASER OPERATING MODE - CW

REF PARAMETER UNIT VALUE REF PARAMETER UNIT VALUE
1a WAVELENGTH m 2a WAVELENGTH m
1b FLUENCE Kj/cm2 2b IRRADIANCE kW/cm2
1c PULSE WIDTH s or ns 2c BEAM SPOT mm
DIAMETER
1d BEAM SPOT DIAMETER mm 2d AVERAGE POWER W
1e EXPOSURE DURATION S
1F PRF* Hz
1g PEAK POWER MW
1h AVERAGE POWER W

TABLE 7 - AEROPLANE LASER HARDENING - TYPICAL PARAMETERS (Clause 3.11)

NOTES:
1. The numerical values associated with all nominated pulse and CW laser weapon parameters will be specified or agreed with the
Aircraft Design Organisation (Clauses 3.11.27 and 3.11.29).
2. The number of exposures to the above parameters will be defined by the Aircraft Design Organisation.
3. The kill levels to be associated with the threat values given in the above table will be that of mission completion (see Leaflet 22).
4. *Pulse Repetition Frequency.

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Service Garment Reference

28V DC Socks and gloves AP3456 Part 1 Section 2 Chap 2 Para 16
electrically heated Para 146
Personal ventilator - - - Guidance shall
Tactical ventilator be sought from
the Project
Team Leader
Compressed Anti-G Suit " " " Chap 2 Para 23
Air (or
Oxygen)
Oxygen Oxygen regulator " " " Chap 1 Para 30
Pressure jerkin (see " " " Chap 2 Para 67
Section 4, Leaflet 3)
Filtered Blown Aircrew Respirator - - - Guidance shall
Air be sought from
the Project
Team Leader

TABLE 8 - AEROPLANE SERVICES FOR AEA (Clause 3.12)

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Aircraft Configuration
Flight Phase
Ref Lift Devices Under- Airbrake Category Speed Range
carriage Applicable
3.5.48 Cruise Up In B As defined in Specification
for AAR
3.5.48 Manoeuvre Up In B As defined in Specification
for AAR
3.5.48 Approach Down In and Out C Min approach to max
permissible
3.5.48 Landing Down In and Out C Min approach to max
permissible

TABLE 9 - AIRCRAFT CONFIGURATIONS AND SPEED RANGES - (Tankers)

Aircraft Configuration
Flight Phase
Ref Lift Devices Under- Airbrake Category Speed Range
carriage Applicable
3.5.66 Cruise Up In & Out A Specified speed range for
AAR, or max range
achievable in relation to
3.5.66 Manoeuvre Up In and Out A tanker’s performance

TABLE 10 - AIRCRAFT CONFIGURATIONS AND SPEED RANGES – (Receivers)

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MILITARY SPECIFIC SYSTEMS - LEAFLETS

1 References. 2
2 Armament Installations – Weapon Release and Fuzing. 11
3 Armament Installations - Jettison Systems. 15
4 Armament Installations - The Effect of Firing Air Weapons on the Behaviour of
Turbine Engine Aircraft. 17
5 Gun Installations - General Recommendations. 20
6 Gun Installations - Gun Gas Concentrations. 24
7 Gun Installations - Gun Blast: The effect of Gun Firing on Turbine engines. 25
8 Installation of Explosive devices - General Recommendations. 27
9 In-Flight Refuelling Systems - General Recommendations. 30
10 Arresting Hooks for Land-Based Aeroplanes. 47
11 Installations for Emergency Recovery from Stall and Spin - General Information
and Recommendations. 60
12 Installations for Emergency Recovery from Stall and Spin - Parachute
Installations. 64
13 Installations for Emergency Recovery from Stall and Spin - Rocket Installations. 67
14 Target Towing Installations - Definitions and Glossary. 69
15 Target Towing Installations - General and Operational Requirements. 72
16 Target Towing Installations - Aerodynamic and Flying Qualities. 74
17 Target Towing Installations - Loading and Shedding. 76
18 Target Towing Installations - Cockpit Controls and Indicators. 78
19 Reduction of Vulnerability to Battle damage - General Requirements. 81
20 Protection of Aircrew against Conventional Weapons - General Requirements. 89
21 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents - Definitions. 91
22 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents - Nuclear Weapon Effects on Aircraft. 98
23 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents - General Recommendations - Chemical and
Biological Warfare Agents. 107
24 Protection from the Effects of Nuclear Explosions, LASER Weapons, Chemical
and Biological Warfare Agents - LASER Weapon Effects on Aircraft. 117
25 Brake Parachute Installations - Safety and Strength Recommendations. 125
26 Integration of Stores - Integration Methodology. 129
27 Integration of Stores - Description of Design Considerations & Loading Actions. 131

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LEAFLET 1

REFERENCES

ARMAMENT INSTALLATIONS

AGARD Reports

AGARD-AR-107 Drag and other aerodynamic effects of external stores.

AGARDOGRAPH202 Store separation.

Air Publications

JSP 482 MOD Explosives Regulations.

A R C Reports

R and M 3438 Wind tunnel experiments on the flow over rectangular

cavities at subsonic and transonic speeds. RAE report
dated 1964

R and M 3503 Aerodynamic Loads on external stores. A review of

experimental data and method of prediction.

Armament Design Memoranda

Arm DM24 Strength requirements for bombs and similar stores and
their carriers, release mechanisms and ejector units for use
in aircraft.

Defence Standards (Def Stan)

13-37 Pt 1 (Cancelled 1982) Airborne weapon fuzing systems (implements STANAG

3525 Cancelled 1997).

59-36 Pt 2(Superseded by IEC TS 62239) Electronic components for defence purposes -

procedure for selection and specification.

59-411 Electromagnetic compatibility.

59-71 Pt 1 Crimped electrical connectors.

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 DEF STAN 00-970 PART 13/11
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MOD(PE) Publications

AvP 118 (S/S by Def Stan 59-411) Guide to electromagnetic compatibility in aircraft Systems.

Ordnance Board Procedures (OB Procs)

41273 Principles of design and use for electrical circuits

incorporating explosive devices.

41754 Safety of fuzing systems. Design safety principles.

RAE Reports

Aero 2511 Low speed wind tunnel tests on the flow in bomb
bays and its effect on drag and vibration. Report on
Canberra bomb bay dated 1954

Tech Report 68086 The measurement of fluctuating pressures in and behind

the bomb bay of a Canberra aircraft.

RAE Specifications

WE 659 Design requirements for aircraft systems fitted with electro-

initiated explosive devices.
STANAGs

STANAG 3575 Aircraft Stores Ejector Racks.

STANAG 3441 Design of aircraft stores.

STANAG 3558 Location of Electrical Connectors for Aircraft Stores.

STANAG 3576 Electrical Connector for Dispensers and Internal

Intervalometer Type Rocket Launchers for Aircraft.

STANAG 3605 Compatibility of Arming Systems and Expendable Aircraft

Stores.

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GUN INSTALLATIONS

RAE TECHNICAL REPORTS

72216 Evaluation of the French AME M621 Model Fl 20 mm Gun.

Helicopter, dated 1972
72163 Ground Trials on the Oerlikon 304 RK 30 mm gun. dated
1972
RAE TECHNICAL NOTES

Arm 1329 Gun, Aden, 30 mm Recoil and Runout loads. Aden 30 mm

no longer in inventory.

Arm 545 Gun, Aden, 30 mm mounting loads.

Arm 589 The reduction of blast pressures from Aden guns by the
use of obstructions in the path of gun gases.

Structures 890 Blast from moving guns.

RARDE TECHNICAL MEMORANDA

28/70 A study of gun blast in relation to that from a moving

explosion. Report on moveable guns dated 1970

29/70 Loads on surfaces due to gun blast. Rifle used as basis for
loads on nearby surfaces and mounted guns dated 1970

34/72 Experiments on gun blast shields to reduce impulsive loads

on nearby surfaces. Dated 1972

17/74 A theoretical model of the blast from stationary and moving

guns.

Note: The foregoing documents are government sponsored; the

list is not exhaustive and additional information may be
available from AD AS DMPS and other sources.

INSTALLATIONS FOR EMERGENCY RECOVERY FROM STALL AND SPIN

A & AEE Notes

2113 Test methods and flight safety procedures for aircraft trials
which may lead to departures from controlled flight.

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 DEF STAN 00-970 PART 13/11
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AGARD

Flight Test Manual Vol.2 Chap.8 Pt.II Dec.1963.

RAE Technical Reports

67197 Low speed wind tunnel tests on the effects of tailplane and
nacelle position on the superstall characteristics of
transport aircraft. (Also available as ARC R&M 3517).

RAE Technical Notes

Aero 1323 Note on recovery from spins by tail parachutes.

Aero 1576 Model spinning data affecting strength requirements.

Mech.Eng.202 The effect of stress waves on the strength of brake

parachute cable.

Mech.Eng.311 The minimum parachute size and maximum towing speed

in relation to strop characteristics.

Mech.Eng.318 The towing characteristics of stable parachutes at various

altitudes.

Mech.Eng 390 The drag and opening characteristics of parachutes of

several designs when towed behind an aircraft; their
consistency and comparison with previous data.

RAE Technical Memoranda

Aero 968 Feasibility of rocket assisted recovery from the deep stall.

RAE Specification

WE 659 Design requirements for aircraft systems fitted with electro-

initiated explosive devices to minimise hazards from
environmental electromagnetic fields.

TARGET TOWING INSTALLATIONS

AGARD Reports

AGARD-AR-107 Drag and other aerodynamic effects of external stores.

5
 DEF STAN 00-970 PART 13/11
SECTION 4

AGARDOGRAPH 202 Store separation.

A R C Reports

R and M 3503 Aerodynamic loads on external stores. A review of

experimental data and method of prediction.

British Standards

BS 3G 100(Multipart) General requirements for equipment for use in aircraft.

Def Stan

05-123 (S/S by MAP RAs) Technical procedures for procurement of Aircraft, Weapons
and Electronic Systems.

Military Specifications (MIL-SPEC)

MIL-A-8591 (S/S by MIL-STD-8591) Airborne stores, suspension equipment and aircraft-store

interface (carriage phase); general design criteria for.

MOD (PE) Publications

AvP 118 Guide to electromagnetic compatibility in aircraft Systems

(S/S by Def Stan 59-411)

Ordnance Board Procedures (OB PROCS)

41273 Principles of design use for electrical circuits incorporating

explosive devices.

RAE Specifications

WE 659 Design requirements for aircraft systems fitted with electro-

initiated explosive devices.

STANAGs

STANAGs (3441, 3558 and 3576) Dimensional requirements of airborne stores.

PROTECTION FROM THE EFFECTS OF NUCLEAR EXPLOSIONS, LASER WEAPONS,

CHEMICAL AND BIOLOGICAL WARFARE AGENTS

Note: See also Leaflets 21 to 24.

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 DEF STAN 00-970 PART 13/11
SECTION 4

NWE General

HTI-R-78-109 Nuclear Weapons Effects Programs; User's Manual.

Horizons Technology
Inc
San Diego,
California

MSDS T(F) M805 MSDS/NS Guide 5: Designers' Guide to the Nuclear

Pilkington FB Survivability of Electro-Optical Components.
MSDS (NSG) Frimley
* (Restricted)

RMCS Shrivenham Nuclear Weapon Effects Course.

Course Notes

NH 9006 Jan 1983 Nuclear Vulnerability Handbook - A Handbook

BAe Dynamics Group and Guide for Designers of Electronic Equipment.
Bracknell Division,
Bracknell, Berks.

NWE Blast

MSDS/NSI/R32 The Effects of Thermal Radiation and Blast from a

MSDS(HSG) Frimley Nuclear Explosion on some Glass Reinforced Plastic
* (Restricted) Components.

RMCS TN Struct, 2 Report on the Sixth International Symposium

RMCS Shrivenham on Military Applications of Blast Simulation.
Smith PD, Pennelegion L.

SID 63-656 A Study of High Altitude Nuclear Burst Data.

North American Aviation Inc.
Downey, California, USA.
Hodden DT, Fowler RA.

AFSWP-877 (NAVORD 4486) The Effect of Altitude of Detonation in the

Naval Ordnance Lab, First Thermal Pulse and Early Shock Pressures
White Oak, Maryland, USA. from an Atomic Bomb.
Willet JE.

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SECTION 4

NWE Thermal
Radiation

SR1 Proj IMU-4021 Survey of the Thermal Threat of Nuclear

Stanford Research institute, Weapons.
Mento Park, California,
Roger JC, Miller T.

WADC TR-54-103 Behaviour of Magnesium and Fiberglass Panels

Write Air Dev Centre. subjected to Thermal Radiation.
Wright-Patterson AFB, Ohio.
Feigen M, Ambrosio A

AMRL-TR-64-139 Experimental Determination of the Maximum Safe

Aerospace Medical Research Lab, Thermal Radiation Loads for a Fighter Bomber
Wright-Patterson AFB, Ohio. Cockpit.
Kaufmann WC.

AWRE SLF N 20/81 Thermal Tests on Military Aircraft

AWRE, Aldermaston, Transparencies.
Balderston J.

MOD (PE) WD (NUC) 2/4 Thermal Data Book.

D/DP (N) 21/5/17
MOD, Whitehall, London.

NWE Radiation

Shell P-602 Effects of Nuclear Radiation on Jet Fuels.

Shell Development Company,
Emeryville, California, USA.
Nixon AC, Thorpe RE, Monor HB.

RMCS TN PD/28/82 and 83 Effects of Radiation on Optical Fibres: CVD

RMCS Shrivenham Contract No RV 26-4

ASD-TDR-63-893 The Effects of Nuclear Radiation on Explosive

Aeronautical Systems Div, Solids.
Eglin Air Force Base, USA.

STL-GM-TR-0165-00358 Effects of Nuclear Radiation on some

Space Technology Labs, Materials and Electronic Components.
Los Angeles, USA.

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 DEF STAN 00-970 PART 13/11
SECTION 4

BAe NH9008 iss 2 Transient Radiation Effects on Electronic

BAe Dynamics Group, Components and Circuits - Guidelines for Avionic
Bracknell Div Engineers.
*(Restricted)

AWRE SSWL 2/76 A Guide to Radiation Effects on Electronic

AWRE Aldermaston Equipment at the Tactical Level.
(Restricted) NWE EMP

A&AEE Note 3205 The Procurement of Aircraft for Survival in

Aeroplane and Armament an Electromagnetic Pulse Environment.
Experimental Establishment,
Boscombe Down, Salisbury.

AFWL IN 315 EMP Penetration through Advanced Composite

F Casey Skin Panels.
EMtec Inc, Los Angeles,
California.
(AFWL Note).

Chemical and Biological

Weapon Effects (C&BE)
General

Stanag 2133 Vulnerability Assessment of Chemical and

Nuclear, Biological and Biological Hazards. (Cancelled)
Chemical Operational
Interservice Working
Party (NBCOIWP)

AC/225(Panel V11/ASP) D/24 Chemical Agents Removal Using Sacrificial

NAAG Coatings.

AC/225(Panel V 11/ASP) WP/56 Resistance of Canopy Materials to Chemicals.

NAAG

AEP 7 Chemical Defence Factors in the Design of Military

NATO Panel V 11 Equipment.
*(NATO Restricted)

AC225 (Panel V 11/ASP)WP/16 Probability of Contamination of Aircraft

NAAG Cockpits during Flying or Taxying in a Chemical
Environment.

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 DEF STAN 00-970 PART 13/11
SECTION 4

AC225(Panel V11/ASP)WP/28 Use of Dilution Pressure Demand Oxygen

NAAG Systems - Protection against CW Agents.
* (NATO Restricted)

Chemical Research & Dev Centre NBC Contamination Survivability. A

US Army Armament, Munitions Handbook for Development/Management of Material
and Chemical Command, Programs.
Aberdeen Proving Ground
Maryland, USA.

CRDEC-CR-87033, December 1986 Nuclear, Biological and Chemical Contamination

Battelle Columbus Laboratories, Survivability Methodology. A Manual for
USA. Equipment Developers, Contractors, and
Bailey P, Hill T E, McNeeley Government Combat and Material Developers.

Documents marked thus * are classified and may be available only on a need-to-know basis.

BRAKE PARACHUTE INSTALLATIONS

R.A.E. Reports

Aero 2379 The use of landing parachutes for aircraft.

R.A.E. Technical Notes

Mech.Eng.202 The effect of stress waves on the strength of brake

parachute cables.

Mech.Eng.311 The minimum parachute size and maximum towing speed

in relation to strop characteristics.

Mech.Eng.318 The towing characteristics of stable parachutes at various

altitudes.

Mech.Eng.390 The drag and opening characteristics of parachutes of

several designs when towed behind an aircraft; their
consistency and comparison with previous data.

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LEAFLET 2

ARMAMENT INSTALLATIONS

WEAPON RELEASE AND FUZING SYSTEMS: FACTORS INFLUENCING SAFETY AND

RELIABILITY OF ELECTRICAL CIRCUITS

1 INTRODUCTION

1.1 The dangers associated with the carriage of weapons on aircraft demand that special
attention be paid to the electrical release and fuzing circuits. The problems are accentuated on
aircraft that can carry nuclear weapons. The aim must be that conditions can never arise which
could result in accidental release of weapons, live or safe, or, on the other hand, in failure to
release the weapons, live or safe as selected. This is the objective of the requirements of
Section 3 Clause 3.2.4 and the purpose of this leaflet is to provide a guide to the design and
engineering of armament electrical systems to meet these requirements, and to indicate
acceptable practices. The requirements apply to all aircraft.

2 TYPES OF FAILURE

2.1 The types of failure that need to be guarded against may be sub-divided as follows:

(a) Faults in equipment and components:

(1) those which make equipment operate inadvertently,

(2) those which prevent equipment operating,

(3) those which cause equipment to operate incorrectly.

(b) Faults in cables and connectors:

(1) short circuits,

(2) open circuits,

(3) cross connections during assembly and servicing,

(4) low insulation resistance between circuits,

(5) high resistance connections.

(6) variation of insulation and contact resistance with Environment and age.

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 DEF STAN 00-970 PART 13/11
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(c) Faults in power supplies:

(1) voltage below acceptable limits,

(2) voltage above acceptable limits

(d) Malfunctions from:

(1) electrical transients in excess of acceptable limits,

(2) induced currents,

(3) conducted and radiated interference.

3 AIRFRAME, PYLON AND CARRIER CIRCUITS

3.1 Each circuit of a duplicated system should be independent and connected to its own
feeder bar, so that a failure of either circuit will leave the other fully operative.

3.2 Care should be taken to ensure that each side of a duplicated system (e.g. fuzing,
release) is connected via separate plugs and sockets and that each side is connected to the
appropriate feeder and independent frame connection (see Section 3, Clause 3.2.4). Circuits
supplied from different feeders should not be run in the same loom.

3.3 To minimise the risk of failure by inadvertent damage or enemy action duplicate services
should be separately routed wherever practicable, subject to the limitations imposed by aircraft
size and design. Note that Section 3 Clause 3.2.8 requires the run of the cables associated with
the circuits for firing, fuzing, release and jettison to be identical in all aircraft of the same type,
mark, role and modification standard. When it becomes necessary to run a cable loom carrying
armament circuits alongside other aircraft service looms, the length of runs in close proximity
should be kept to a minimum and additional protection provided at the points where they touch.
If however, general services circuits are encased, these may run alongside armament circuits in
separate encasements without further protection.

3.4 It is desirable that cables controlling the weapon fuzing, and release systems should not
be loomed together. Individual conduits may, if necessary, be laid up alongside one another
provided there are no overriding requirements in the weapons specification.

3.5 Clause 3.2.4 of Section 3 requires that cables are encased; however, exceptionally
(subject to agreement of the Project Team Leader) encasement may not be necessary in the
following instances:

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 DEF STAN 00-970 PART 13/11
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(a) Where physical constraints imposed by the aircraft structure severely limit the
space available, the wiring is well protected mechanically by the structure itself and the
area is free from the risk of contamination.

(b) Where encasement would introduce a servicing penalty disproportionate to the

protection provided.

The omission of encasement under the terms of this paragraph constitutes a Design Deviation,
any requests for which should be made to DSA-MAA-Cert-ES4-ArmSys@mod.uk in accordance
with the Technical Procedures applicable to the aircraft.

3.6 Junction boxes used for armament wiring should meet the following requirements:

3.6.1 Internal wiring should be loomed and secured such that hinge points do not occur at
cable terminations.

3.6.2 Cable end fittings should be of an approved type of crimped terminal.

3.6.3 Junction boxes and terminal blocks should be so designed and mounted that there is
adequate protection against the ingress of fuel oils and moisture or other contaminants (see Part
1 Section 6 Leaflet 14, Para 6)

3.6.4 The location of junction boxes and terminal blocks within the aircraft structure should be
such that they are at all times easily accessible for servicing.

3.6.5 Ideally, when the M.A.S.S. is set to safe, there will be no power supplies to any
armament junction box or unit, to avoid the possibilities of faults developing between the power
input and the outputs. When this is not possible, for example Logic supplies are required when
the M.A.S.S. is set to safe, special precautions should be taken to ensure that these supplies
cannot be short-circuited to the outputs from the junction box or unit. There is no objection to
mounting the M.A.S.S. relays in a junction box which acts as a power distribution box, provided
that the outputs to release units etc are not subsequently routed through that box.

3.6.6 The wires carrying the standing voltage supplies to the M.A.S.S. should not be formed
into any loom which includes release or other sensitive circuits.

3.7 The size of the conductors in cable runs to individual store stations, in addition to having
the required electrical characteristics, should be such that the cables have adequate mechanical
strength in their operational environment. Voltage drop in distribution circuits should be within
the limits specified in the Aircraft Specification or BS 3G 100 Part 3(Multipart) as appropriate.

3.8 Particular care should be taken to protect cables, connectors, plugs and sockets from
damage likely to result from adjacent moving parts and the movements of aircrew and servicing
personnel.

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3.9 Connector systems used for inter-connecting looms in the airframe should be of a type
approved by DSA-MAA-Cert-ES4-ArmSys@mod.uk Crimped cable connections should be
employed. Glands should be used rather than bulkhead connectors to minimise the number of
discontinuities in the circuits. Where plugs and sockets are used these should be to the
requirements approved by DSA-MAA-Cert-ES4-ArmSys@mod.uk

3.10 Approved connector modules carrying armament services should be segregated from
modules used for non-armament services. Where space and weight consideration make it
impracticable to mount the armament modules in a separate frame they should be kept apart
from the non-armament modules by a spacer and clearly labelled "ARMAMENT".

3.11 Inter-connections between aircraft and carriers or pylons should be by connectors of

types complying with Para 3.9 above. Construction of the interconnections should be in
accordance with those requirements approved by DSA-MAA-Cert-ES4-ArmSys@mod.uk may
be used. Cable connections between airframe and carriers should be positioned and secured so
as to prevent damage.

3.12 Care should be taken to ensure adequate flexibility of cable assemblies where physical
movement takes place in flight or during servicing.

3.13 The number of connections of all types in any one circuit should be kept to a minimum.

3.14 For non-nuclear weapons there should, ideally, be no standing voltage supplies in the
pylon carrier or station until such time as the release switch guard is raised. Low voltage or
current limited sources should be used for Bomb-on-Station signalling, Built In Test Equipment
(BITE) and similar circuits so that in the event of a fault there will be insufficient power to operate
the release unit. For Guided or other weapons, where it is not possible to meet these
requirements, there must be complete physical separation between the circuits associated with
these weapons and the normal release and fuzing circuits, and separate connectors should be
provided for the two groups of circuits.

3.15 The use of components with exposed connection tags is to be avoided whenever
possible.

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LEAFLET 3

ARMAMENT INSTALLATIONS

JETTISON SYSTEMS

1 INTRODUCTION

1.1 Design features associated with any jettison facility can only be considered in conjunction
with the particular Aeroplane Specification. However, some features are common to most
jettison systems and these are discussed in the following paragraphs.

2 GUIDED WEAPONS

2.1 These are normally jettisoned safe unless otherwise specified, either by firing the motor
of the weapon, by ejection or by gravity release.

2.2 Rail launched missiles can often be jettisoned by firing the rocket motor with the missile
in a safe configuration. Alternatively it may be possible to jettison the missile and launcher as a
single unit. For emergency jettison, the second method is likely to be the quickest and safest,
but for selective jettison, there are advantages in being able to fire the missile as the first option,
whilst retaining the ability to jettison the missile complete with launcher as a secondary method,
should the rocket motor fail to fire.

3 EMERGENCY JETTISON

3.1 The emergency jettison of launchers/carriers alone or with pylons may be required, in
addition to the separate release of the stores themselves. In such cases consideration should be
given to the feasibility of jettisoning multi-store carriers complete with stores.

Note: It will be necessary to consider the consequences of jettisoning both fully and partly
loaded carriers.

3.2 The definition of Emergency Jettison, Part 0, ANNEX E, requires that the aeroplane be
cleared of all non-nuclear stores as rapidly as possible and without danger to the aeroplane. It is
therefore desirable that the stores, carriers and pylons where necessary, be jettisoned
simultaneously. However, simultaneous jettison may endanger the aeroplane for any one or all
of the following reasons:

(a) The jettison of all stores together may result in aerodynamic disturbances leading
to undesirable behaviour of the aeroplane or even temporary loss of control.

(b) Unacceptably high reaction loads may be generated in the aeroplane structure.

(c) The stores may contact the airframe after leaving the carriers.

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 DEF STAN 00-970 PART 13/11
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(d) The stores may collide with each other after leaving the carriers.

(e) Jettisoned pylons, and jettisoned carriers complete with stores, may behave
violently after release due to their aerodynamic characteristics. Consequently it may be
necessary to release the stores (including carriers and pylons if appropriate) in a
predetermined safe order. In this case control of the order of release must be automatic,
and require no more than a single operation of the emergency jettison switch for complete
jettison.

3.3 In some configurations it may be desirable to jettison the store complete with its carrier or
launcher although release of the unloaded carrier or launcher may endanger the aeroplane. In
such cases means must be provided to inhibit automatically the jettison of the unloaded carrier
or launcher.

3.4 In the case of internally carried stores, operation of the emergency jettison control should
release these stores only if the weapon bay doors are open. It is not considered necessary to
provide automatic opening of the weapon bay doors specifically for emergency jettison, unless
otherwise specified. If, however, an automatic door opening system is provided for normal
operation this could be used for emergency jettison if necessary.

3.5 It is desirable that emergency jettison should be possible at any point within the flight
envelope. However, in practice, there may be limitations and it is desirable that these be
determined at an early stage and that DSA-MAA-Cert-ES4-ArmSys@mod.uk be notified of any
such limitations.

3.6 Considerable electrical power can be required to operate a comparatively large number
of release units simultaneously. Consideration should therefore be given to the capacity of the
aeroplane emergency battery to ensure that this is sufficient to perform an emergency jettison in
the event of a complete aeroplane power generation system failure.

4 SELECTIVE JETTISON

4.1 The selective jettison facility may form part of the stores management system provided
that the mandatory requirements are met. The operation may be controlled by the pilot or a
second crew member, as appropriate; the RPAV operator or for the case of a highly/fully
autonomous vehicle the vehicle management system.

4.2 As stated in Para 3.1 there is sometimes a requirement to jettison carriers or carriers and
pylons in addition to the stores themselves.

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LEAFLET 4

ARMAMENT INSTALLATIONS

THE EFFECT OF FIRING AIR WEAPONS ON THE BEHAVIOUR OF TURBINE ENGINE

AIRCRAFT

1 INTRODUCTION

1.1 This leaflet discusses those factors which are known or suspected of causing propulsion
system malfunctioning when guns, rockets or guided weapons are fired. In addition to the
information contained herein much of the content of Leaflet 7 is also applicable to the firing of
rockets and guided missiles. It must be considered in conjunction with this leaflet.

1.2 Malfunctioning depends upon the type of weapon, its installation and proximity to the
engine air intake system, the propulsion system and its associated air intake and the flight and
propulsion system conditions under which the weapons are fired.

2 CAUSES OF MALFUNCTION

2.1 Propulsion system malfunction is known to be caused by a multiplicity of factors which

include the entry of hot gases from the rocket/propulsion motor efflux into the air intake system,
the variations in total pressure at propulsion system entry, the total temperature distortion at the
propulsion system entry and ingestion of weapon launch hardware. The problems are extremely
complex and although the actual mechanism of the subsequent effect is not understood the
individual effects in a total complex set of effects are (see Leaflet 7 Para 3.1)

2.2 There can be adverse effects due to the chemical composition of the weapon gases,
(e.g. chemical reaction on propulsion system components or aspiration of inert gases and see
Para 4.3 (e) on fouling of optical or IR measuring systems)

3 FLIGHT CONDITIONS

3.1 The flight conditions under which the weapons are fired (i.e. altitude, attitude, forward
speed, throttle position, aeroplane incidence, side slip and ambient temperature) affect the
problem in various ways in that they:

(a) can have a critical influence on the expansion of the weapon gases and hence
the characteristics of the pressure wave,

(b) can have a critical effect on the efficiency of the air intake system directly and
also indirectly by their influence on aeroplane attitude; and

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(c) can have a critical effect on the propulsion system compressor surge margin.
(The surge margin at a particular engine speed although basically a function of engine
design will vary with change in altitude, forward speed, ambient temperature and air intake
performance)

3.2 In general, increasing altitude and decreasing speed will increase the likelihood of
propulsion system malfunctioning because the reduction of air inlet mass flow causes a higher
proportion of weapon exhaust to propulsion system airflow to be ingested.

3.3 Manoeuvres aggravate the problem in their effect on attitude and hence on air intake
performance.

3.4 During supersonic flight conditions blast waves from a fired weapon can modify
instantaneously the organisation of the aeroplane forebody / propulsion system inlet airflow and
cause propulsion system malfunction.

4 ROCKET INSTALLATIONS

4.1 The remarks in this paragraph apply equally to unguided rockets and to guided weapons.

4.2 The type of propellant used in a rocket motor, its rate of burning and the flight conditions
under which it is launched are the important factors in determining the chemical composition,
concentration, temperature gradient and pressure wave characteristics of the efflux gases.

4.3 Characteristics of the overall weapon design which can affect the propulsion system
problem are:

(a) acceleration - in determining the time the rocket efflux may act within the vicinity
of the air intake system;

(b) aerodynamic stability and/or guidance system - in determining the motion of the
weapon in the early stages of its trajectory. (Any motion which causes the efflux to be
directed towards the intake system is undesirable);

(c) the number and size of rockets fired at any one time and firing sequence (i.e.
ripple or salvo);

(d) the positioning of the weapon installation in relation to the air intake system. (This
is clearly a most critical factor);

(e) the base material of the propellant, for example it is known that an aluminium
base can cause serious propulsion system turbine blading contamination, deterioration
and fouling of optical and IR sensors.

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5 GUN INSTALLATIONS

5.1 The type of propellant used in gun ammunition and flight conditions under which it is fired
are important factors in determining the chemical composition, concentration and
pressure wave characteristics of the discharge gases. See also Leaflet 5, Para's 4.5 and
4.6.

5.2 Characteristics which can affect the propulsion system are:

(a) muzzle blast pressure - may create a shock wave which is critical to the intake
airflow;

(b) firing rate - the rate and rhythm of firing may be critical to the intake airflow;

(c) the positioning of the gun and/or gas discharge port(s);

(d) the composition of the propellant - it is known for example that use of
phosphorous compounds as a flash suppressant can cause contamination of turbine
blades and fouling of IR and optical sensors;

(e) vibration - excessive vibration may upset airframe mounted propulsion system
control systems to the extent that physical damage can occur which may then result in
malfunctioning of the propulsion system.

6 RECOMMENDATIONS

6.1 The aim should be to separate the area of weapon disturbances from the propulsion
system air intake as far as possible, thus avoiding weapon effects.

6.2 For rocket propelled weapons, the maximum separation between weapons and air
intakes should be provided. In addition, the initial trajectory should be studied in relation to the
air intake system, so as to avoid directing the rocket efflux towards the intake if it can be
avoided.

6.3 Where design necessitates a close relationship of the weapons and propulsion system
intakes, due consideration should be given to the fact that it will not be possible to design an
propulsion system with enough basic surge margin to cater for serious weapon wake ingestion,
although the characteristics of the propulsion system and the intake and fuel systems should be
such as to minimise the effects of that ingestion.

6.4 An unambiguous statement of weapon firing operational envelope requirements must be

prepared during the initial weapon system design study. It must for example include statements
as to whether propulsion system acceleration/deceleration, supersonic firing, and aeroplane
incidence are relevant factors.

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LEAFLET 5

GUN INSTALLATIONS

GENERAL RECOMMENDATIONS

INTRODUCTION

1.1 This leaflet gives general recommendations on both fixed and free gun installations.

2 GUN MOUNTING

2.1 LOCATION

2.1.1 For fixed guns, fuselage mounting is preferred. This location is conducive to accuracy
and a satisfactory projectile pattern, it can also ease the problems of ammunition stowage and
gun accessibility and it avoids conflict with wing aerodynamic demands. Location should be
chosen after full consideration of the effect of gun vibration on avionic and other systems, and
indirectly, by changes to the aeroplanes (or equipment's) permanent magnetism, or compass
detector units. The effect of gun blast on propulsion system airflow must be considered and the
effect of vibration induced by blast should be minimised; the blast signature varies with height
and aeroplane speed and theoretical methods of predicting these variations are becoming
available. Muzzle flash should not be visible from the cockpit.

2.2 ACCESSIBILITY

2.2.1 Accessibility, and its effect on maintainability, are of prime importance. Adequate space
should be provided to enable adjustments (e.g., harmonisation) and clearance of stoppages to
be made without difficulty. The fitting and removal of guns, including barrel removal, will
generally require the provision of access doors or panels of substantial size. Hinged doors, with
a single fastening device (e.g., shoot bolts) are preferred.

2.3 LIFTING REQUIREMENTS

2.3.1 The weight of guns generally precludes simple manhandling during fitting or removal
operations. Provision for the attachment of suitable lifting equipment should be made.
Permanently installed lifting equipment is preferred.

2.4 MOUNTING STIFFNESS

2.4.1 The stiffness of the mounting has a direct effect on force transmission, gun kinematic
functioning, and accuracy and the characteristics of the transmitted force can be substantially
altered by various damping techniques. The mounting design should aim to avoid effects, during
gun firing, detrimental to gun functioning and accuracy and to equipment by ensuring maximum
separation of fundamental and harmonic frequencies of vibration of the guns and of the
mounting structure.

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3 AMMUNITION CONTAINERS

3.1 FITTING AND REMOVAL

3.1.1 Removable ammunition containers are recommended; they should be readily detachable
and easily secured. The number of containers depends primarily on ease of handling and the
operational environment. If their (fully loaded) weight prevents simple manhandling, provision for
the attachment of suitable lifting equipment should be made. Permanently installed lifting-
equipment is preferred.

3.2 CONTAINER CONSTRUCTION

3.2.1 Ideally, the container should be made in stainless steel, but at least the face of the tank
in contact with the cartridge nose should be in this material. No paint, varnish or similar material
capable of peeling or chipping should be used.

3.2.2 Internal surfaces of the container should be smooth and free from irregularities; bolt and
rivet heads should be avoided. Vertical dividers may be necessary to prevent bunching of the
ammunition and the distance between the top of such dividers and the container lid should be
minimised to avoid flailing of the belt.

3.3 CONTAINER MARKING

3.3.1 The position of the ammunition in the container should be indicated in the following way:

(a) Internally. An Engraved diagram of a round and link should be placed

precisely at the spot where the first round to be flaked must be positioned showing it
pointing in the correct direction. The round should be coloured red and the link black.

(b) Externally. An engraved plate, or engraved diagram, should be placed on an

outer face clearly showing the manner in which the ammunition must be flaked and any
leading link requirement. The flaking should be shown as viewed from the side of the
container on which the diagram is situated.

4 FEED AND EJECTION

4.1 INSTALLATION EFFICIENCY

4.1.1 Maximum gun reliability will only be achieved by complete control of belt feed, and empty
case and link ejection during gun functioning.

4.2 FEED CHUTES

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4.2.1 Flexible ducts may, be used, but these should be generously supported over their length.
Rollers may be provided at changes of direction to minimise drag on the belt, including inside
the tank and at the tank exit. Joints in the chute system should overlap in the direction of feed so
that no edges face oncoming rounds. It should not be possible for projections (e.g., bolt and rivet
heads) to make contact with the moving ammunition or belt.

4.3 CHUTE BREAKS

4.3.1 No opening in the chute for loading, or making and breaking of links, should present any
hindrance to the moving ammunition or links, particularly the last one. A break in the chute, giving
good access to the belt, should be provided as close to the gun as possible.

4.4 BELT DRAG

4.4.1 To minimise belt drag due to friction and inertia, the outlet from the tank should be as
close to the gun as possible. Where long feed runs are unavoidable, the use of feed assistors
may be necessary, but the need for their introduction depends on the relative value of gun pull to
belt drag (under all conditions of flight).

4.5 EJECTION GENERALLY

4.5.1 The energy in an ejected case can be high. Chutes should be rigid and without severe
bends to prevent jamming. Wherever possible, links, cases and complete rounds should be
collected - preferably links should be collected separately. Links require guide rails right up to
the exit from the link chute.

4.6 EJECTION OVERBOARD

4.6.1 When cases, or links, or complete rounds, are ejected overboard, chutes should be
directed downwards, and entry into the airstream should be arranged to avoid not only bunching,
particularly of the low density links, but also to prevent the debris from being ingested by the
propulsion system.

5 VENTILATION

5.1 Ventilation of the gun installation to reduce gas concentration to an acceptable level
should also include ventilation of any tank in which empty cases are collected.

6 BLAST TUBES

6.1 The dimensions of the blast tube and blast deflectors should be sufficient to
accommodate the effects of gun jump and projectile dispersion at any harmonisation setting.

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7 FIRING CIRCUIT

7.1 In addition to the safety arrangements specified in Section 3, Clause 3.3.5, it should be
possible, before beginning maintenance work, to break the firing circuit as close to the gun as
possible. The preferred method is a plug and socket arranged so the firing lead, disconnected
from the gun and connected only to the aeroplane, may hang in full view and prevent the closing
of the gun compartment door.

8 FREE GUNS - ADDITIONAL CONSIDERATIONS

8.1 GENERAL REQUIREMENTS

8.1.1 Free gun installations require considerations additional to those for their fixed gun
counterparts. Adequate attention should be given to the gunner's ability to search for and
acquire the target. Gunner comfort and hence ease of operation is of prime importance.

8.2 FIRING INTERRUPTORS

8.2.1 With free guns, the possibility of self-inflicted damage to the aeroplane should be
eliminated by suitably programmed firing interrupters. The arrangement should operate through
all aeroplane configurations and under all conditions of flight.

8.3 LINK AND CASE DISPOSAL

8.3.1 The method of link and case disposal (ejection overboard or collection) should be
adequate over the full range of gun movements. In general, collection is preferred.

8.4 SIGHTING

8.4.1 There is no restriction on the location of the sighting device except that its mounting
arrangements should be such as to minimise the effects of gun vibration.

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LEAFLET 6

GUN INSTALLATIONS

GUN GAS CONCENTRATIONS

1 INTRODUCTION

1.1 When an aeroplane gun is fired, roughly half the products of propellant combustion are
themselves combustible, i.e. carbon monoxide and hydrogen. For example the Mauser 27 mm
gun produces about 1.5 litres at Normal Temperatures and Pressures (NTP) of this combustible
gas mixture for every round fired. On aeroplane installations which are part of the aeroplane
structure or are podded, ventilation should be provided when the gun is firing to ensure that
these combustible gases are sufficiently diluted to prevent the build-up of a potentially explosive
material. The design certificate for this aspect of gun installations is quoted in Para 2 below.

2 DESIGN CRITERION

2.1 Section 3, Clause 3.3.7 states that flammable gas from a gun installation shall not
present a hazard to the aircraft. The design of the gun installation and its associated air purging
system shall be such that the concentration of flammable gun gas within the gun/ammunition
compartment and, where applicable, the compartments immediately surrounding the blast tube
and barrel, is not permitted to rise to a potentially dangerous level. While the concentration of
gas in the region immediately adjacent to the leak source on the gun will be high for as long as
the gun is firing, purging air should be arranged to dilute the gas as close as possible to the
source, so that the concentration in all other areas never exceeds 80% of the Lower limit of
Aircraft Hazard (LLAH), as defined in Ref 1. Particular attention should be paid to achieving low
concentrations at any potential sources of ignition. The installation is to be capable of
withstanding, without damage, the effects of any transient ignitions that may occur at
concentrations below the LLAH.

3 VENTILATION

3.1 A guide to ventilating air mass flow requirements is given by Oelman in Ref 2 but this
should be updated by information from the gun designer who will measure the volume of
exhaust products from a representative quantity of guns. It may be possible to dump part of
these gasses directly overboard so reducing the ventilating air mass flow required through the
gun installation.

References: 1 RAE Tech Note Arm 654 - J C Iredale-Williams

2 RAE Tech Memo Arm 1552 - S H Oelman Dated 1954

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LEAFLET 7

GUN INSTALLATIONS

GUN BLAST: - THE EFFECT OF GUN FIRING ON TURBINE ENGINES

1 INTRODUCTION

1.1 Blast, unburnt propellant ingestion and an increase of gas mass flow from aeroplanes
guns can severely affect the behaviour of gas turbine engines. The guiding principles for
reducing the effect of gun blast and gas on engines are given below.

2 CAUSES OF ENGINE MALFUNCTION

2.1 The primary cause of engine malfunction is the entry of the gun exhaust gases and gun
blast pressure wave into the engine intake system. The subsequent airflow disturbance leads to
compressor surge and, in some cases, flame-out.

2.2 The flight conditions under which the gun is fired (i.e., altitude, speed, aeroplane
incidence, ambient temperature) affect the problem in various ways in that:

(a) they influence the blast wave development and the degree of gas ingestion and
hence their effect within the intake,

(b) they affect the efficiency of the air intake system directly, and also indirectly
through their influence on aeroplane attitude,

(c) they affect the engine compressor surge margin.

2.3 High altitude and low speed tend to increase the likelihood of an engine malfunction due
to gun blast and/or gun gas ingestion. Additionally, aeroplane manoeuvre aggravates the
problem by its effect on attitude and hence on engine intake performance.

3 GUN INSTALLATION EFFECTS

3.1 The magnitude and frequency of the blast pressure wave are influenced by the following
parameters:

(a) calibre, muzzle velocity, and muzzle pressure of the gun,

(b) the number of guns (affecting each intake),

(c) the rate of fire,

(d) the size and design of any blast deflector.

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3.2 The positions of the gun gas exits (usually the blast tube mouths) relative to the engine
intakes are the greatest single factor in the influence of gun blast and gas on engine behaviour.
The effect of the disturbance will to some extent be dependent on the air intake design, but the
greater the separation of the blast tube exits from the engine intakes, the smaller will be the
effects of the gun. Positioning of gun gas exits forward of the intakes should be avoided or the
maximum possible separation achieved; failing this, effective blast deflectors should be fitted.

3.3 When pressure disturbances and/or gas do reach the engine, the extent of their effect is
a function of' compressor design and compressor surge margin.

4 RECOMMENDATIONS

4.1 POSITIONING

4.1.1 The gun gas exits should be at least 2 metres behind the plane of the engine air intakes.
Lateral separation should be as great as possible.

4.2 BLAST CONTROL

4.2.1 The maximum possible amount of gun gas expansion should be achieved before its entry
into the atmosphere. Blast deflectors can be beneficial in controlling the shape of the pressure
wave to reduce the effect felt at the engine air intake.

4.3 ENGINE DESIGN

4.3.1 When the guns and air intakes are unavoidably close, the characteristics of the engine,
its intakes, and its fuel system should be such as to minimise the blast effects.

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LEAFLET 8

INSTALLATION OF EXPLOSIVE DEVICES

GENERAL RECOMMENDATIONS

1 INTRODUCTION

1.1 Explosive devices are inevitably a potential hazard to both air and ground crews. They
are a hindrance during servicing, and may also necessitate additional tradesmen and
installations at Service units to deal with them. They should not therefore be used if the object
can be achieved without them. However, the increasing difficulty in accomplishing many
operations, particularly those of an emergency character (removal of canopies and the like)
seems likely to lead to an increase in their use. Their potential danger and their use in
emergency functions demands a very high standard of reliability. This Leaflet therefore reviews
their characteristics and makes recommendations concerning their use.

2 CHARACTERISTICS OF EXPLOSIVES

2.1 DETERIORATION

2.1.1 All explosives suffer from chemical deterioration, which, in general, increases rapidly with
rising temperatures. On a purely temperature/chemical change basis alone the Ordnance Board
report that, for a typical explosive device one month at 60°C (140°F) corresponds to 1½ years at
32°C (90°F)

2.1.2 Temperature changes, jolts and vibrations may affect mechanical properties, cause
alterations in dimensions, break seals or fragile parts and cause crumbling of explosives.

2.1.3 A combination of these physical and chemical effects is likely to lead to the entry of
moisture as the store breathes, thus providing a further potential source of deterioration.

2.2 REDUCTION IN SAFETY

2.2.1 Explosives may deteriorate in such a way that they become less safe than when new.
Materials for explosive devices for use in aeroplanes are selected as far as possible so that
degradation does not make them unsafe, consistent with having a material of the high degree of
reliability necessary if the device is to function correctly.

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2.2.2 Explosives used in aeroplanes are frequently detonated electrically. One way in which
deterioration, particularly that caused by excessive vibration or rough handling, reduces safety
margins, occurs because dust from the explosive can be ignited by much less energy than that
normally required. For instance, a case is recorded in which ignition was deemed to have
occurred through a static charge unknowingly acquired by a man engaged on servicing
operations. Action has been taken with the Services to deal with the static charge risk. Other
accidental firings have occurred through induced currents from nearby circuits or from radio
frequency radiations, and from the application through defective insulation of a potential
difference from the metallic structure of the aeroplane which, except in the case of some
explosive circuits, is used as a common negative lead.

2.3 LIFING

2.3.1 Explosives deteriorate with time, and to ensure that they remain serviceable while they
are installed, a life must be allotted. In the case of new types, or of existing types used under
substantially different conditions, this may involve carrying the devices as passengers in the
actual aeroplane environment in which they are to be used so that they may be examined and
tested periodically until a realistic life has been determined. The Services naturally wish for the
longest possible life consistent with fitting into the normal aeroplane servicing pattern.

3 RECOMMENDATIONS

3.1 DUPLICATION

3.1.1 In view of the high degree of reliability required, consideration should be given to the
possibility of duplication of the system or of individual explosive devices.

3.1.2 The decision on the extent of duplication to be provided will need to be considered for
each particular application and will depend on such factors as:

(a) the reliability of the explosive devices,

(b) the seriousness of a failure,

(c) the complexity of the circuits,

(d) the power demands,

(e) the feasibility of designing and accommodating duplicated installations, neither of

which will interfere with the correct functioning of the other, and

(f) the environment of the device, which may affect its reliability.

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In considering the above factors, the advice of DSA-MAA-Cert-ES4-ArmSys@mod.uk should be

sought, particularly on items (a) and (d)

3.2 ACCESSIBILITY

3.2.1 Clause 3.4 calls for explosives to be easily accessible. It has been found, particularly with
explosive bolts used to operate jettison devices, that the degree of handling necessary, which is
largely determined by accessibility, is a major factor in the prevalence of damage and in the
production of dust from the explosive charge. This as noted above, materially increases its
sensitivity. Cases have also been found of the insulation of detonator leads being damaged, and
in some cases the leads themselves being broken during the fitting of difficult bolts. Wherever
possible, leads should be further protected at vulnerable points by the use of insulating sleeves.

3.3 PROXIMITY OF RADIO FREQUENCY AND OTHER CIRCUITS

3.3.1 The precautions necessary in connection with the proximity of other circuits are detailed
in Part 1, Section 6, Clause 6.6

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LEAFLET 9

IN-FLIGHT REFUELLING SYSTEMS

GENERAL RECOMMENDATIONS

INTRODUCTION

NOMENCLATURE

1.0.1 Nomenclature and legends for use on controls, panels, and displays in aircrew stations:

Air to Air Refuelling AAR

A/R

SCOPE

1.1.1 This Leaflet sets out the design requirements for the installation of air to air refuelling
equipment in both tanker and receiver aircraft which are additional to those given in Part 1,
Section 5, clause 5.2

PROBE AND DROGUE SYSTEM

1.2.1 Unless otherwise specified in the aircraft specification the probe and drogue system shall
be used whenever the aircraft requires (AAR) capability.

GENERAL REQUIREMENTS FOR TANKER INSTALLATION

1.3.1 The tanker (AAR) system may take the form of one or a combination of the following
types:

 Fuselage centreline station single or twin power source derived from tanker
hydraulic electrical or pneumatic system.
 Wing pylon mounted dry pod power source derived from tanker system or ram air
turbine containing integral fuel storage.
 Fuselage pylon mounted wet pod power source derived from tanker system or ram
air turbine containing integral fuel storage (buddy buddy system)
 Fuselage pylon mounted dry pod power source derived from tanker system or ram
air turbine.

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1.3.2 The aircraft specification for a tanker aircraft will indicate whether the AAR equipment is
to be in package form for use as role equipment. In such cases the installation and removal time
of the equipment shall be minimised (actual time to be as agreed with the Project Team Leader).
Loose equipment associated with the role change shall be kept to a minimum.

1.3.3 The aircraft specification for a tanker aircraft will also:

 Define the number of refuelling stations to be installed the range of fuels to be used
and the rates of flow required from each station.
 define the speed and altitude envelope in which the dispensing of fuel and the hose
trail and rewind operation must be possible
 Indicate whether the tanker is required to be equipped as a receiver and whether
the tanker crew should be provided with a means of monitoring the receiver aircraft
L (e.g. periscope or close circuit television).
 Indicate if the tanker requires to be provided with sufficient illumination to enable
the receiver pilot to carry out AAR at night with safety.
 Indicate whether or not the signal and tunnel lights need to be compatible with a
night vision goggles equipped receiver aircraft.
 Indicate whether there is a requirement for covert air to air refuelling and provide
details as appropriate.

GENERAL REQUIREMENTS FOR RECEIVER INSTALLATION

1.4.1 The aircraft specification for the receiver aircraft will:

 Define the form of probe to be used - fixed retractable or removable. The interface
with the aircraft systems will be defined in the systems structural specification.
 Indicate whether emergency probe extension is required.
 Indicate whether the receiver aircraft will be required to receive fuel from a tanker
aircraft fitted with a boom to drogue adaptor.
 Define the tanker type and the speed altitude envelope from which the aircraft is to
receive fuel.
 Indicate if there is a requirement for probe lighting ant requirement for probe
lighting any requirement for brightness central being stated.

OPERATIONAL REQUIREMENTS

2.1 Flight envelopes appropriate to carriage hose trail fuel transfer and hose rewind shall be
established in accordance with the aircraft specification.

2.2 Carriage of the AAR system or any aspect of the refuelling operation shall not degrade
the safety of the aircraft. Consideration shall be given to the effect of dumping fuel or jettison of
any part of the AAR system with regard to fire hazard or impact damage to the tanker.

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2.3 The system shall permit repeated dry contacts (e.g. fir training) without degradation of
any system capability.

AERODYNAMIC FLYING QUALITIES

3.1 The aerodynamic and flying qualities of tanker and or receiver aircraft shall be examined
in accordance with Part 1, Section 2, Clause 2.18

SYSTEM PERFORMANCE

AERODYNAMIC PERFORMANCE

4.1.1 In non turbulent conditions the fully trailed hose and drogue shall present a sufficiently
stable target to the receiver to maximise the probability of successful receiver contacts.

4.1.2 The hose and drogue shall recover from any instability induced by non damaging
external influences once those external influences have ceased.

HOSE DRUM UNIT PERFORMANCE

4.2.1 Trail and rewind operation. The hose trail and rewind speed shall be controlled to
minimise instability of the hose and drogue at any point during trail or rewind throughout the
AAR envelope. In particular there shall be no contact between the hose and drogue and the
tanker airframe during the trail and rewind operation such as to be a flight safety hazard or
cause damage to either the airframe or hose drogue.

4.2.2 Hose response:

 Engagement - The response system shall be capable of accelerating the

hose in the rewind direction to prevent hose whip, looping, oscillation or
excessive slackness at receiver closing speeds of up to 3 m/sec. The
response system shall also allow engagement at closing speeds as low as
0.6m/sec.
 Disengagement - The maximum relative speed at which the receiver can
withdraw without inadvertent disengagement occurring shall be 1.2 m/sec.
In addition the trail response action shall preclude hose oscillation whilst
the receiver is in contact.
 Response - During engagement contact and disengagement any hose
oscillation or heave whether created be aerodynamic or, mechanical effect
shall be minimal shall not overload any part of the tanker or receiver
aircraft and shall naturally damp out.

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4.2.3 Arrest of motion - The motion of the hose drum and hose assembly shall be safety
arrested at both full trail and full rewind.

FUEL SYSTEM PERFORMANCE

4.3.1 Pressure at coupling - The fuel transfer system shall regulate the static pressure at the
coupling to 345 kPa ± 35 kPa throughout the range of flow. The system shall not subject the
reception coupling to pressure other than that due to static head during normal engagement and
disengagement. Note: The control pressure of 345 ± 35 kPa shall not be exceeded except for
short duration surge peak pressures in accordance with 4.3.2 In cases where the AAR system is
at its maximum flow rate capability and the receiver back pressure is not sufficient to generate
the control pressure them lower pressures are acceptable.

4.3.2 Surge Pressures - The system shall be designed so that surge pressures do not exceed
the proof pressure of the systems (receiver, tanker, pod or hose drum unit as defined in Part 1,
Section 5, clause 5.2. Possible sources of pressure surge are:

 Valve closure in the tanker refuelling system.

 Valve closure in the receiver refuelling system.
 Receiver disengagement at any flow rate up to and including the system
maximum.

CONSIDERATION SHALL BE GIVEN TO THE RATE OF CLOSURE OF REFUELLING

VALVES TO MINIMISE THE SURGE PRESSURE EFFECT.

4.4.1 NOTE: For access (hose drum performance) and (fuel system performance) above,
instantaneous peak surge pressures in the hose and reception coupling may exceed the steady
state proof pressure but shall not cause permanent deformation, nor limit hose or coupling life.

4.4.2 With tanker single failure conditions the tanker refuelling system shall not generate
drogue probe interface pressures greater than the proof pressure or 828 kPa.

4.4.3 Stall pressures - The fuel transfer system shall be designed so that any stall pressure
experienced by the receiver aircraft is no greater than 414 kPa.

ENVIRONMENTAL CONDITIONS

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GENERAL

5.1.1 Unless allowed by the aircraft specification, the installation of AAR capability shall not
restrict the environmental conditions in which the basic aircraft can operate when the AAR
equipment is in the stowed condition.

ELECTROMAGNETIC COMPATIBILITY
5.2.1 Installation of AAR equipment shall not compromise the electromagnetic compatibility of
the aircraft, nor shall the operation of the AAR equipment be adversely affected by the existing
airframe environment.

5.2.2 The system shall be designed so that it can withstand the effects of lightning strike, such
that with or without the hose trailed fuel vapour ignition will be prevented and no physical
damage will occur such as to cause a hazardous situation. Consideration shall be given to
maintaining the AAR unit in an operational condition following a lightning strike.

SYSTEM DESIGN

INSTALLATION REQUIREMENTS

6.1.1 When installed, AAR equipment shall not interfere with satisfactory operation of parent
aircraft equipment (e.g. slats, flaps, undercarriage doors etc). Further, the AAR equipment shall
retain adequate clearance with the ground during take off and landing including emergency
cases (e.g. with oleo and tyre collapsed).

6.1.2 The location of each AAR system and the trailed position of the hose and drogue shall be
such that:

 Adequate clearance is maintained between the tanker and receiver and between each
receiver on the approach to and during contact over 30º included angle cone centred
on the hose normal trail position.
 The hose and drogue are clear of any significant destabilising effects due to
aerodynamic wake, jet efflux or propeller slipstream.
 Any destabilising effects influencing the receiver handling or positioning on the
approach to and during contact are minimised.
 The mating dimensions of the reception coupling shall conform to STANAG 3447.

6.1.3 The design dimensions of the probe installation should be such that:

 The mating dimensions of the nozzle probe mast shall conform to STANAG 3447.
 A clearance space shall be provided around the nozzle probe mast installation in
accordance with STANAG 3447.

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 The probe itself shall be located so that its nozzle is adequately within the receiver
pilot’s vision when the pilot views the tanker aircraft and the trailed drogue during
closure to contact. The probe location shall be such that the effect of airflow around
the receiver in drogue stability just prior to contact is minimised. Interaction of tanker
wake and receiver bow wave shall also bi taken into account. In addition the probe
shall be located so as to minimise any adverse aircraft handling effect or any effect in
engine air intakes.

6.1.4 Ground Handling and Ground Support Equipment:

 It shall be possible to gain easy access to the AAR equipment to allow pre and post
flight checks and maintenance without removal of the equipment from the aircraft.
 Ground test procedures shall be adequate, simple and brief.
 The requirement for use of special to type ground equipment shall be minimised.
 The equipment shall have provision for hoisting, loading and transportation.

STRUCTURAL DESIGN

6.2.1 The strength of the AAR installation and the associated aircraft structure shall b designed
in accordance with Part 1, Section 3, Clauses 3.1 and 3.2 and be capable of withstanding all
loads generated by the system and applied to the system throughout the defined flight envelope
for carriage and refuelling operations. See Para 1.3.1 above.

6.2.2 Probe.

 Refuelling probes may be fixes, removable or retractable. The probe shall be able
to be locked in the extended position. Telescoping proves shall not permit fuel to
enter between the inner and outer tubes. The probe installation shall not degrade
the performance of the aircraft outside that required by the aircraft specification.
 The probe shall be provided with a weak link so that the nozzle will break away to
prevent any abnormal condition resulting in loads in excess of the maximum design
loading being applied to the robe. Loads which should be considered during the
design of the weak link should include but need not e limited to; axial, radial, and
moment breakaway loads for the disconnection loads between the nozzle and the
reception coupling, and the effects of fuel pressure on such loads.
 In all cases the design of probe fairings, doors and mechanisms shall be such that
the drogue cannot be caught upon them. It is therefore desirable that doors be
closed after the probe is deployed.

6.2.3 The design of the drogue and coupling shall be such that:

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 Those parts likely to contact the receiver shall be resistant to damage, or

detachment, and shall not cause significant damage to the receiver.
 The failure of any single part of the drogue which may result from the drogue
striking the receiver , or the receiver ‘spoking’ the drogue ribs shall not cause the
total detachment of any part of the drogue.

FUEL SYSTEM DESIGN

6.3.1 Fuel tanks and vents - All aircraft fuel tanks, including those added as part of the tanker
conversion shall be provided with sufficient inward venting so as to prevent the possibility of
inward collapse. Similarly the receiver vent system shall be capable if venting fuel, should a
failure of a fuel cut off valve build up in the fuel tanks. Where fuel tanks are installed within an
aircraft pressure cabin, they shall be vented to atmosphere and shall be contained within
structural enclosures. Means if detecting leaks into the enclosure shall be provided and the
enclosure shall incorporate overboard drains. Care shall be taken to provide vents of sufficient
size to cope with the worst conditions of operation and to eliminate the risk if the vents being
blocked due to icing or foreign matter. (Part 1, Section 5, clause 5.2 refers).

6.3.2 Receiver valves - Unless otherwise specified, a valve shall be incorporated into the probe
such that if the nozzle breaks off at the weak link the valve shall be retained in the nozzle and
shall close to seal off the reception coupling so that fuel be given to providing protection of the
receiver aircraft system in the event of nozzle breakage.

6.3.3 Fuel pipes.

 Fuel pipes shall not run through passenger, crew, cargo or baggage compartments nor in
hazardous proximity to hot air ducts, electrical wiring and electrically operated equipment
contained in bays unless they are without couplings and adequately protected against
potential sources of ignition and damage. Any space between a pipe and its protection
shall be adequately vented and drained. See also Part 13. Section 3.5, Clause 3.5.15
 Consideration shall be given to the need for purging fuel pipes associated with the AAR
equipment following completion of the operation, if the residual fuel could constitute a fire
hazard.

6.3.4 Fuel System Integration.

 Fuel storage included as part of the AAR equipment shall be available for use by
the tanker when required, except where the fuel type carried by the tanker for
transfer to a receiver is unsuitable for use by the tanker itself. In this case the 2
types of fuel shall be segregated so that it is impossible for the fuels to be mixed or
for the fuel transfer system or the tanker engine(s) to be fed with the wrong type of
fuel.
 The tanker system shall be capable of supplying the AAR equipment at sufficient
rate to meet the tanker refuelling requirement specification without compromising
the tanker engine fuel feed.

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 The effect of a single failure in the AAR equipment supply system on maximum
transfer rate shall be minimised. Failure of the main transfer pump shall not prevent
fuel transfer. However, reduced rates of flow will be acceptable.
 Devices for relief of surge pressure shall not require spillage of fuel outside of the
fuel system.
 There shall be no leakage from any part of the system prior to contact, during
contact or post contact. Leakage during the act of contact or breakaway, whether
normal or emergency, shall be minimised even at the most adverse condition of
probe coupling engagement.
 In addition to the requirements of Para 6.2.1 above the probe location shall take
into account the risks associated with fuel spillage on to the windscreen and with
fuel entry into engine air intakes or any other intake.
 There shall be no requirement for fuel to be supplied by the aircraft to the AAR
package once refuelling is complete (e.g. to lubricate bearings etc) which
significantly increases the aircraft’s minimum landing fuel.

ELECTRICAL/ ELECTRONIC SYSTEM DESIGN

6.4.1 Electrical equipment - Installation of associated electrical equipment shall comply with
the requirements of Part 1, Section 4, Clauses 4.26, 4.27 and Part 1 Section 6 Clause 6.6

6.4.2 Static electricity - Electrical connection (to discharge static) shall be established between
the tanker and receiver before fuel is transferred.

6.4.3 Bonding.

 Full electrical bonding of AAR equipment shall be provided.

 Following contact, equipotential of tanker and receiver aircraft shall be maintained
via a conductive hose and drogue.

6.4.4 Radio communications - It shall be possible to conduct the refuelling operation safely
without use of radio communications between tanker and receiver. However, when radio
communication is used the equipment and aerial installation shall be safe and shall enable
effective communication with the receiver in formation with, in close proximity to, and on contact
with the tanker. It may be acceptable to prohibit specific air to air or air to ground
communications such as HF during AAR operations. The Project Team Leader shall be notified
of any such restrictions.

COCKPIT CONTROLS AND INDICATORS

6.5.1 In addition to the requirements of Part 1, Section 4, clause 4.19, the following minimum
controls and indicators shall be provided in the tanker and receiver:

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Tanker Controls.

 Signal light brightness

 Hose jettison control
 Fuel valve control
 Emergency signal control
 Hose trail and rewind control
 ON/OFF master switch

Tanker indicators:

 Fuel flow indicator

 Fuel transferred
 Hose stowed indicator
 Hose position or movement indicator
 Fuel valve position
 Drogue at full trail and ready for engagement
 Signal repeater indicators

Receiver controls:

 Refuel valve control

 Probe position (for retractable probes)
 Probe lighting controls (if fitted)

Receiver indicators:

 Probe position (for retractable probes)

 It shall be possible to assess the fuel contents available for transfer either
by reference to existing indicators or by provision of additional indicators.
Unless otherwise specified, fuel amount transferred shall be measured in
kg and the fuel flow indicator gauged in kg/ min.

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1 IN-FLIGHT REFUELLING ASCC Air Std 10/49

CONTROLS
(a) Selection/Reset Controls In single or tandem seat
Aircraft operated by the left
hand. In multi-seat Aircraft
near to and operated by the
second crew member or flight
engineer.
(b) Disconnect Control In single or tandem seat
Aircraft on the control column
and operated with the right
hand. In multi-seat Aircraft on
the control column and
operated by the pilot.

FUEL SYSTEM CONTROLS

6.6 INDICATORS AND MARKINGS

6.6.1 Signal lights.

 Signal lights shall be included to provide the receiver with an indication of the
status of the air to air refuelling equipment. These lights shall be mounted adjacent
to each other in a position that is clearly visible to the pilot of the receiver aircraft
when astern of the tanker in pre contact positions.
 Where the tanker is fitted with multiple independent refuelling stations, separate
light systems shall be fitted for each station.
 These lights shall be duplicated to allow redundancy and shall be capable of being
dimmed for night operation.
 The following lights shall be used and operated in the given order.

6.6.2 RED - When the master switch is on and when the hose is stowed, trailing, rewinding or
otherwise unsafe for receiver contact. It shall be possible to operate this light manually when
required.

6.6.3 AMBER - When the AAR system is ready for receiver contact.

6.6.4 GREEN - When the hose has been pushed into the refuelling range (nominally 1.5 m
from the full trail position) and the AAR system fuel valve has opened.

 Refer to Section 3 clause 3.5 for signal sequence and response.

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6.6.5 Hose lights - Lights shall be provided to illuminate the refuelling hose markings (see Para
6.6.3) can be seen at night. The colour of the light shall be white. The light shall be installed or
shielded in such a manner as to prevent the light from being a source of direct or reflected glare
to the tanker crew or receiver pilot. Consideration shall be given to varying the intensity of this
light to cater for ambient light conditions and to minimise the tanker signature.

6.6.6 Hose markings - The hose shall be marked to provide the receiver pilot with indications
of the inner and outer limits of the hose refuelling range, the optimum position within that
refuelling range and the position when disconnection at full trail is imminent. Refer to Fig 3 for
recommended pattern of hose markings.

6.6.7 Drogue lighting - The drogue shall be illuminated for night operation. The lighting shall be
self contained within the drogue and shall not require power from the tanker for operation.

6.6.8 Tanker markings - Markings shall be applied to the tanker which are clearly visible to the
receiver pilot and give guidance as to the correct positioning and movement of the receiver for
the refuelling operation. The markings shall be effective by both day and night. Refer to Fig 1
and Fig 2 for recommended patterns of tanker markings.

SAFETY CONSIDERATIONS

SAFETY REQUIREMENTS

7.1.1 in addition to the requirements defined elsewhere in this Clause, the equipment shall be
designed so that no single failure on the tanker or the receiver shall cause fuel or fuel vapour to
be released into the cockpit or cabin of either aircraft it in any other way endanger their safety.

7.1.2 A safety assessment and a zonal analysis shall be made of the AAR system and of its
interface with the tanker aircraft own fuel system. Reference shall also be made to Part 1,
Section 5, Clause 5.2

7.1.3 The AAR installation shall be so designed that a single failure subjects the receiver
aircraft to minimal foreign object damage. Internally mounted AAR units shall be adequately
isolated from crew, passenger and freight compartments so that the operation of the AAR units
does not endanger personnel nor compromise the usage of such compartments.

7.1.4 Where the AAR installation uses any of the tanker aircrafts vital systems i.e. fuel or
power supplies, safeguards shall be taken to protect the tanker from the consequences of
malfunction of the AAR units. It shall be possible to isolate the AAR unit from the aircrafts
systems.

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EXPLOSION PROOFING (also see Section 3 clause 3.5)

7.2.1 The system shall be designed to minimise the fuel vapour within the init as well as
preclude ignition of any vapour that does exist. This requirement may be satisfied by a
combination of:

 Design (e.g. incorporation of flame traps in electrical boxes)

 Test
 Air purging if pod interiors HDU compartments. This operation shall not require the
tanker to de pressurise.

JETTISON AND DUMPING SYSTEM

7.3.1 Where the aircraft specification requires the capability for package or hose jettison, or
package fuel dumping, the jettison dumping systems shall operate satisfactorily throughout a
separately defined jettison dumping envelope.

7.3.2 The jettison dumping systems shall be powered independently of the normal package
systems.

7.3.3 In the case of package jettison or hose jettison, pyrotechnic devices shall not be used,
unless permitted be either the tanker aircrafts or equipment specification.

7.3.4 The jettison dumping operation e.g. guarded or locked toggle. Double pole switches are
preferred.

7.3.5 The hose drum outlet shall be sealed at the point if hose separation. The seal shall
withstand system pressure.

OTHER EMERGENCY SYSTEMS

7.4.1 Where emergency systems such as emergency hose trail and or rewind or emergency
probe extension are required by the aircraft specification these systems shall not compromise
the reliability of the basic system. Where they are of the “one shot” type, they shall not degrade
the safety of the aircraft once they have been operated. These systems shall be powered
independently of the normal package systems.

DE ICING AND ANTI ICING

7.5.1 De icing arrangements for the nozzle and coupling are not required.

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7.5.2 Consideration shall be given to the possible need for providing anti icing techniques for
the AAR unit. In such an event provision of AAR unit temperature monitoring, heating control
and attendant cockpit indicators will be required.

TESTS REQUIRED

GROUND TESTS

8.1.1 Ground tests shall be conducted to demonstrate compliance with the requirements if the
aircraft specification for the following aspects:

 role change
 Fuel capacity of tanker AAR package (usable and total)
 Proof pressure test
 Fuel transfer tests
 Surge pressure tests (normal and emergency breaks and receiver cut off)
 Mechanical and electrical function
 structural requirements
 environmental requirements
 emergency system operation and failure cases
 EMC requirements
 drainage

FLIGHT TEST

8.2.1 Flight testing shall be conducted in accordance with Part 1, Section 2 – Flight.

AIR-TO-AIR REFUELLING
TEST EQUIPMENT

9 INSTRUMENTATION

9.1 Details of the parameter ranges, accuracies and resolutions are given in Part 1 Section 2
Leaflet 10, Table 1.

9.2 The following parameters should be recorded for the tests detailed in Section 3, Clauses
3.5.36 - 78, for both tanker and receiver Aircraft, or a lesser selection determined as appropriate
and agreed by the Project Team Leader.

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Item Parameter

1 Time base
2 Manual event marker
3 Crew speech
4 Indicated airspeed
5 Altitude (pressure)
7 Total temperature
8 Angle of attack
9 Pitch attitude
10 Bank angle
11 Sideslip angle
12 Heading
13 Pitch rate
14 Roll rate
15 Yaw rate
16 Longitudinal acceleration
17 Lateral acceleration
18 Normal acceleration
19 Flap/slat setting
20A Wing sweep position
21 Airbrake position
22 Failure state
24 Fuel contents
25 Pitch inceptor position
26 Roll inceptor position
27 Yaw inceptor position
28 Pitch inceptor force
29 Roll inceptor force
30 Yaw inceptor force
31 Pitch trim position
32 Roll trim position
33 Yaw trim position
34 Pitch motivator position
35 Roll motivator position
36 Yaw motivator position
44 Throttle position(s)
45 Rotational speed(s)

9.3 For the tests described in Clauses 3.5.50 to 58 and 3.5.71(b), chase or receiver Aircraft
equipped with video or cine camera should be provided.

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Black and white stripes 75 mm typically wide

Dimensions in mm
NOT TO SCALE
Suggested Under Wing Air to Air refuelling Pod Alignment Markings
Fig 1
FORWARD

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Suggested Tanker Underfuselage Alignment Markings

Receiver Aircraft View
Fig 2

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Typical Hose Identification Markings

Fig 3

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LEAFLET 10

ARRESTING HOOKS FOR LAND-BASED AEROPLANES

1 INTRODUCTION

1.1 This leaflet gives background and supplementary information relating to the requirements of
Clause 3.6

1.2 Most runways will have two arresting gears but there is a tendency to provide three or more
arresting gears on some. In a particular emergency landing a pilot may opt to engage the first available
gear at high speed rather than a later one. There are also a number of cases, apart from those
connected with failure to stop, where the first arresting gear is engaged as a precautionary measure.
Such instances include failure of the aeroplane directional control system and landing with the
undercarriage in an unsafe condition. In take-off conditions it is rare for a failure to occur which prevents
some reduction of speed before arresting gear is engaged but nevertheless the engaging speed may be
high.

1.3 When the aeroplane enters the arresting gear it will not necessarily be moving along the runway
centreline. It is usual to specify that requirements shall be met at all off-centre distances up to 20% of the
total distance between the runway edge sheaves and for tests to be done up to this distance. For current
gears this is 12m (40 ft).

2 FACTORS AFFECTING DESIGN HOOKLOADS

2.1 In all arresting gears the load on the arresting hook will vary throughout the arrest. At any point
the load is determined by a number of factors. There are three basic phases of an arrest:

(a) initial impact,

(b) dynamic braking,

(c) steady braking.

2.2 These terms are used loosely and the three phases are far from distinct. In most arrests they will
merge one into the other. Only where the aeroplane is significantly lighter than that for which the
arresting gear is designed will the initial impact hookload be clearly distinguished from the several peaks
of the hookload in the dynamic phase and exceeds them in magnitude. There is always considerable
overlap between the later part of the dynamic phase and the early steady braking phase. A distinction
between the three phases must be made in relation to the performance of an arresting gear although the
performance data may be given in such a form that it is not clear from which phase the specific limitation
is derived. Figs. 1 and 2 clearly show the three phases. Fig. 3 is a fairly typical performance and
limitations diagram and has been annotated to show which of the phases may have provided the limiting
speed for a given hookload and mass. The constant speed limits below 9.027 kg mass could result from
either impact or dynamic loads. The curves between 9.072 and 22.680 kg could be dynamic or steady
braking or a combination of both. The curves from 22.680 kg to 27.216 kg are steady braking limits
caused, generally, by a peak late in the pull-out as a result of the aeroplane mass being higher than that
at which the maximum efficiency of energy absorption is obtained. Fig. 4 illustrates these different
characteristics.

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2.3 For any combination of arresting gear and aeroplane the hookload on impact is determined by
the trail angle, the initial impact velocity, and the mass and material of the hook cable. In the dynamic
phase it is determined by the velocity of the aeroplane in relation to the mass and inertia of the moving
parts of the energy absorber. In the steady braking phase the hookload depends on the kinetic energy of
the aeroplane and the energy absorption efficiency of the arresting gear. When differences between
arresting gears are assessed the differences between each of these three phases of the arrest must be
considered independently. Further, as the hookload in the dynamic phase can be increased by
reflections of the stress wave caused by initial impact both in the arresting gear and in the aeroplane, it is
important to consider all combinations of aeroplane mass and speed in relation to the performance data
for each arresting gear to be used. It is also essential to enquire whether the published data for each
arresting gear provide the maximum hookload expected in each phase and also to identify which of the
phases provides the overriding value given in the data.

2.4 IMPACT

2.4.1 Where available data for an arresting gear are based on the steady or dynamic braking loads it is
necessary to formulate an estimate of the initial impact load for the aeroplane/hook-cable combination
being considered because this may provide a greater hookload, and because the direction of the
resultant forces is different. Simple theory predicts the horizontal impact retarding force at:

5 1
 
K
2 m(V) 3   3 newtons
 2

Note: The symbols are defined at the end of this leaflet.

2.4.2 This formula makes no allowance for pretension loads or for increases caused by the reflection of
the stress wave from runway edge couplings. A further 30% should be added to allow for this (Ref 3).
The vertical component of this force is not a maximum when the hook is fully down and the variation of
this component should be studied throughout the upswing. Reference 2 gives some data on this aspect.
For example, for a stiff arm at a trail angle of 60° to the ground on impact the maximum vertical
component occurs during the upswing at an angle of 37°.

2.5 DYNAMIC BRAKING

2.5.1 This is the most difficult phase to estimate. In it the moving parts of the arresting gear are
accelerated violently. The forces required can be calculated (Ref 4) but relevant trials data will be more
accurate if available. Where none are available, bearing in mind the read-across criteria of Para 9.2, a
rough estimate can be obtained by applying the following dynamic magnification factors to the above
formula 2.7 at 100 kts (185 km/h) 3.3 at 150 kts (277 km/h) 3.7 at 180 kts (333 km/h). An alternative is to
base estimates on calculated mean retardations for various aeroplane masses and speeds and apply
appropriate maximum/mean ratios based on experience with similar arresting gears. The following table
gives values for rotary hydraulic arresting gears taken from Refs 4 and 5. Ref 4 also gives some data for
water spray gears.

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WA/WO 0.2 0.4 0.6 0.8 1.0 1.1 1.2

P/M 3.5 2.1 1.7 1.5 1.35 1.5 1.9

2.6 STEADY BRAKING

2.6.1 If trial data for steady braking loads are not available for aeroplanes of similar mass and speed
an error factor of 1.5 should be applied to the value obtained by any theoretical calculations based on
formulae derived from simple Newtonian mechanics or on data obtained from trials in any other arresting
gear that does not fully meet the read-across criteria of Para 9.2.1

3 DESIGN PHILOSOPHY

3.1 The design of the complete installation of an arresting hook may be considered in four parts:

(a) the hook and beak shapes,

(b) the suspension arm,

(c) the aeroplane structure,

(d) the vertical and lateral damping systems.

3.2 Clearly the structure of the whole aeroplane must have adequate static strength for the worst
operational case and fatigue strength to provide adequate life for operational use to the expected
spectrum of loads. The hook point or beak (depending on the design) will have a life of a limited number
of arrests before it is worn to the extent that it must be replaced. Depending on cost and ease of
replacement it is a matter of policy to be determined between the Chief Designer and the Project Team
Leader as to whether the suspension arm and the damping and centralising gear should be treated as
part of the aeroplane structure or as a replaceable part. If they are removable, but not easily so, they
may be put into a third category in that they are statically designed to the same requirements as the
aeroplane structure but have a fatigue life which is less than that of the aeroplane. The strength of the
hook, suspension, and airframe should be adequate to arrest the aeroplane in all specified combinations
of mass and engaging speed. If any specified combination of mass and speed is beyond the energy
capacity of one of the arresting gears considered, the installation should be designed to match the
characteristics of the arresting gear having the greatest energy capacity including any projected gears.

3.3 The length of the hook suspension arm, its trail angle and the shape of the hook must ensure as
far as possible a satisfactory engagement for all aeroplane attitudes and configurations which could arise
as it traverses the hook cable. The trail angle limit of 80° stated is regarded as being beyond the angle
for which satisfactory damping can be provided with a stiff suspension arm. The trail angle should
therefore not exceed 70° and 60° is a preferred value for optimum damping (Ref 2).

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3.4 The location of the hook suspension pivot point on the aeroplane must be considered in relation
to several factors. Clearly it must not be forward of the c of g of the aeroplane. To provide the best
chance of successful engagement after the arresting cable has been trampled by the mainwheels, and to
provide maximum yawing stability during the arrest, it should be as far aft as possible. However if it is at
the tail of an up-swept rear fuselage having a bumper forward of the pivot point then this may, in a
maximum tail-down attitude, cause deflection of the arresting cable and a hook-skip.

3.5 The location of the hook pivot point and the length of the hook suspension arm are also
interdependent in preventing 'cable-slap'. After initial impact the hook cable is subjected to complex
vertical and lateral motion arising from the propagation of the immediate post-impact waves. Care must
be taken to ensure that, wherever possible, contact between the cable and the aeroplane structure,
particularly the tailplane, is avoided. Where this is clearly impossible special protection may need to be
provided. Under certain circumstances there may be considerable lateral movement of the hook in its
fully up position and this may cause scraping of the edges of the stowage tunnel. This also should be
avoided if possible.

3.6 If it is accepted that the flailing arresting hook cable may strike the airframe then consideration
should be given to the effect of the impact and subsequent cable shedding on any equipment located in
the area. Similarly if the hook in its stowed position protrudes beyond the rear fuselage lower skin line
and this is close to the ground it may cause a hazard during trampling.

3.7 All questions of ground clearance must be considered against the requirements of Part 1, Section
4, Clauses 4.13.7 and 4.13.8 which call for trampling tests to be done when any doubts exist about the
clearance being adequate.

4 VARIATION OF AEROPLANE ATTITUDE

4.1 Aborted take-off - There will be two broad areas to be considered, braked and unbraked, but in
both it must be presumed that the main undercarriage and tyres may be fully extended and the nose
wheel fully compressed unless it can be shown by calculation that this combination is impossible.

4.2 Landing - There are two extremes. The first is the airborne engagement at maximum incidence. It
is possible for the arresting hook to engage the hook cable at the same instant as it touches the runway
and before the wheels. The requirement for the maximum angle of 80° between the suspension arm and
the runway (the trail angle) is intended to guard against the hook jamming in this type of touchdown. This
could occur even with an 80° trail angle if the closing angle between the hook pivot point and the runway
(including pitch rate effects) exceeds 10°. However, this is not a normal occurrence, even in an
emergency, and if it is to be considered the Aeroplane Specification will state a requirement. There
would also be a requirement to consider undercarriage strength in this case. The second extreme is the
case where, without significant decrease of speed after a heavy touchdown at high velocity, the main
oleos are extending and lifting the tail of the aeroplane. The nosewheel may be fully compressed. It is
this case which will normally determine the length of the suspension arm.

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4.3 Roll-back - At the end of an arrest residual energy in the arresting gear tape may pull the
aeroplane backwards a short distance. If this motion continues after the hook cable has become slack
because the pilot does not maintain adequate engine power the hook will drop to the ground and, if the
trail angle is greater than the differences between a right angle and the friction angle, it may jam. In this
case it is preferable for the arm to fail as a strut than for damage to be done to the trunnion and back-up
structure and consideration must be given to ensuring that these components are not overloaded. In an
extreme case where the total energy is considerably greater than that for which the arresting gear tapes
are designed to be used normally, maximum tape-stretch will occur and excessive roll-back will follow.
This presents the pilot with a difficult control problem and there may be a peak hookload which is greater
than the maximum occurring during the earlier part of the arrest. This is known as two-blocking (see Fig.
4) and should be prevented in service by appropriate operating limitations.

5 EFFECT OF AEROPLANE ATTITUDE ON HOOK SHAPE

5.1 While the range of arresting hook movement and the design of the beak must be adequate for all
these cases it is important that the profile of the nose and throat of the beak should be optimised for the
median case. This will generally be close to the normal static taxying attitude.

5.2 The range of ground lines which have to be considered will generally be smaller for an airfield
landing than it would be for a carrier-based deck-landing Naval aeroplane. Nevertheless the shape of the
arresting hook will be correctly determined by the same basic principles. These are given in detail in Ref
1 but their application to airfield arresting may be summarised as follows:

(a) All forward faces of the hook, hook beak, and lower end of the suspension arm must be
provided with the correct impact radius. (When an arresting cable is impacted by the hook a kink is
developed which makes an angle of approximately.

 2V  1
1.1 x tan  1   with its original line. The impact radius is the largest radius which can be
 K 3
fitted within this kink to make a tangent with the arms of the kink at each side of the hook).

(b) The vertical throat radius must be greater by 2 mm (0.079 in) than that of the largest hook
cable to be engaged.

(c) The wrap-round radius must suit the characteristics of the hook cable of the arresting gear
which maximises it. (The wrap-round radius is the plan radius of the throat of the hook). For steel
cable the radius should be not less than three times the cable diameter.

(d) The take-off angle must suit the geometry of the arresting gear which minimises it. (The
take-off angle is the angle between the arresting cable and the direction of motion at the end of the
pull-out).

(e) Beak face angles must be determined to give the best possible entry for the cable into the
throat whether the cable is flat on the runway or has bounced to its maximum height.

(f) Beak face angles must also provide adequate coverage of the extremes. Where this is not
possible with a fixed beak, a hinged beak becomes essential.

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(g) The sole of the hook must be flat from beak to heel unless it is hinged. In this case the
individual sections should be flat.

6 EFFECT OF HOOKLOAD ON THE DYNAMICS OF THE AEROPLANE

6.1 On engagement the hook arm is swung violently upwards by the impact forces and will be
approximately stable when the aeroplane has moved 2½ hook arm lengths irrespective of velocity. The
load line as viewed from the side of the runway will be sensibly straight from the hook arm pivot through
the hook throat to the runway edge sheave. There will of course be some variation caused by movement
of the hook cable and other elements of the arresting gear. The amount by which this load line initially
passes above or below the aeroplane centre of gravity will determine the subsequent behaviour of the
aeroplane in pitch. This load line is usually below the centre of gravity and therefore usually provides a
direct increase in nose-down pitching moment. If it acts above the centre of gravity a dynamic analysis of
the pitching motion of the aeroplane is necessary to show that it will not become stable at a tail-down
attitude and that the tail of the aeroplane will not hit the runway.

6.2 Nosewheel-slam is most pronounced in landings where there is an in-flight engagement at a high
angle of incidence and is therefore most likely to be greatest in emergency landings where the first
(approach end) arresting gear is engaged from a late approach. However the maximum slam effect does
not necessarily occur at maximum mainwheel vertical velocity and all practical combinations must
therefore be explored.

6.3 Whether the load line acts above or below the centre of gravity of the aeroplane in the impact and
dynamic phases of the arrest the effect of the design hookloads on nosewheel loads must be considered
both with and without the application of brakes. Another critical case for the nosewheel may arise during
an aborted take-off with heavy braking.

7 VERTICAL AND LATERAL LOADS

7.1 The axial design hookloads are the principal loads determining the component sizes of the
arresting hook installation. Depending on the relationship of aeroplane characteristics to the
characteristics of the arresting gear in each of the three phases of the arrest the axial load will be a
maximum in one of them. However the vertical and lateral loads at the hook pivot or trunnion will not
necessarily be a maximum at the same time. They will be affected by aeroplane geometry, mass, and
moments of inertia; and by the stiffness of the damping system in each plane. Data available from tests
of one aeroplane/arresting gear combination cannot therefore be directly equated to any other
aeroplane/arresting gear combination without a considerable amount of detailed correction. Where
typical vertical and lateral components are required for the formulation of a fatigue spectrum they should
be obtained during prototype tests in the relevant arresting gear.

7.2 MAXIMUM VERTICAL LOAD

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7.2.1 The vertical component of the hookload will be determined by aeroplane pitch attitude and hook
trail angle, in relation to the instantaneous axial hookload. The greatest value may occur during impact
following an in-flight engagement. It could also be a maximum during aborted take-off engagement at
high speed. At whatever point the maximum deceleration occurs in the pull-out the attitude of the
aeroplane should also be considered and the vertical load determined. It will not usually be a maximum
at maximum trail angle. For a stiff arm at a trail angle of 90° the maximum theoretically occurs when the
arm reaches 45°. At a trail angle of 60° it is theoretically at 37° (Ref 2). However this will be considerably
modified by flexibility of the arm and a dynamic analysis is necessary to establish the forces.

7.3 MAXIMUM LATERAL LOAD

7.3.1 The greatest lateral load which might arise is one which by its nature cannot be easily determined
and which cannot be deliberately measured experimentally. It would arise in the case where an
aeroplane for some reason, such as a mainwheel tyre burst at a critical time before engagement, enters
the arresting gear at a large angle of yaw. This would be further increased if the aeroplane was off-
centre but running towards the centre and if the throat radius of the hook was too small for the arresting
hook cable or was caught on a snag in the cable caused by a previous engagement. In the extreme,
welding of the hook and the cable can take place as a result of the heat generated by friction between
them. Under more normal conditions some side load components will be generated by normal off-centre
engagements but in these the natural tendency of the aeroplane is to yaw further away from the centre
line and, as this tends to equalise the hook cable tensions on either side of the hook, the side loads will
be lower.

8 DAMPING, HOOK BOUNCE AND RUNWAY ROUGHNESS

8.1 To ensure a successful arrest it is necessary to prevent the hook from bouncing (or skipping)
over the arresting cable during taxying following impact on any normal runway surface excrescence and
it is necessary to prevent a similar bounce after touchdown close to the arresting cable. Lateral instability
of the suspension arm must also be prevented.

8.2 TAXYING CASES

8.2.1 Runways are normally initially laid to high standards but deteriorate with continued use producing
roughness and other surface changes which may cause the hook to bounce. The hook will fail to engage
the arresting cable if it bounces high enough for the nose radius of the beak to strike the cable above its
centre line. As the cable will have been disturbed by the passage of the aeroplane wheels it is presumed
that the hook cable is flat on the ground when the hook strikes it. To ensure successful engagements the
operators will normally maintain the runway surface free of excrescences up to 15 m (16.4 yds) from the
arresting cable and it will be adequate to show that the nose of the hook beak will return to the runway in
less than this distance; that is, that the hook-skip-distance does not exceed 15 m (16.4 yds) in the cases
specified.

8.2.2 If the damping system is required to prevent hook skip on runways of more than normal
roughness it will be necessary for criteria to be determined in consultation with the operators who control
the condition of the surface. It should be remembered however that tarmac runways can be rolled,
concrete runways can be ground, and snow and ice can be cleared from the essential area.

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8.2.3 Where the maintenance requirement of Clauses 3.6.16 to 3.6.18 conflicts with the damping
requirements of Clause 3.6.10 the damping requirements must be met and special provisions will need
to be made for stowage of the hook after use.

8.3 LANDING CASES

8.3.1 Not only must the vertical damping force be adequate to prevent hook-skip in an extreme
combination of longitudinal parameters but the lateral damping force must prevent lateral instability.
With any practical amount of bank at touchdown a side force at the hook will be generated by the forces
created by the closing velocities. If the component of the vertical forces acting on the arm is greater
than the lateral friction force at the hook the latter will move sideways. In an extreme case the lateral
stop could be broken or the suspension arm could be bent. The lateral damping system must be
designed to prevent this instability. Note that, at a given angle of bank and rate of tail-down pitch, the
lateral force will be proportional to the vertical damping moment. While a high damping moment is
indicated as a solution to a hook bounce problem, a low moment helps to reduce lateral instability. It is
therefore a matter of design to determine a level which provides the best balance between the two
conflicting requirements. A hinged beak will also help to reduce hook bounce and thus may alleviate the
problem by demanding a reduced damper effort for this purpose. The above presupposes a fully
articulated joint at the attachment of the arm to the aeroplane. If a V-frame or combination of V-frame
and articulation are proposed the same requirement applies. The lateral instability may be easier to
prevent but the forces generated may be greater and the mass of the back-up structure increased.

9 TEST PROGRAMME

9.1 TESTS IN THE DESIGNATED ARRESTING GEAR

9.1.1 A full programme would explore all relevant combinations of configuration, speed, mass, and off-
centre distance, taking measurements of hookloads in three axes (or axial hookloads and angular
deflections) in each of the phases of the arrest (impact, dynamic, and steady braking) and other
parameters of interest and would repeat these tests in each arresting gear expected to be used in
service.

9.1.2 In practice this will not normally be possible. The programme adopted should aim therefore to
cover the following objectives in the minimum number of trials:

(a) To establish maximum safe speeds, at landing and take-off masses, for maximum normal
hookloads and any other limitation which may arise:

(1) on-centre,

(2) at the required maximum off-centre distance.

(b) To provide adequate data to allow estimates by extrapolation of measured hookloads and
component forces to the design cases with adequate reliability to cover them for service use.

(c) To prove the functioning of the complete installation in all conditions required.

(d) To determine whether any modifications are necessary to enable the installation to meet
the requirements.

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(e) To confirm necessary inspection procedures, following an arrest, appropriate to both the
aeroplane and the arresting gear.

9.2 READ-ACROSS TO OTHER ARRESTING GEARS

9.2.1 Tests done in one arresting gear may be accepted for clearance purposes to allow operational
use of alternative arresting gear only when all of the following requirements are met:

(a) The hook cable is of the same material and the same or smaller diameter.

(b) The peak dynamic hookloads in the alternative gear have been measured on another
aeroplane at similar masses and have been found to be lower.

(c) The maximum steady braking hookloads in the alternative gears have been measured by
another aeroplane at similar masses and have been found to be lower.

(d) The maximum energy to be absorbed is less than the design limit for the alternative
arresting gear.

(e) The energy absorbing devices are of similar type.

9.2.2 If any one of these criteria cannot be met and it is necessary to use estimated hookloads for
either the dynamic or the steady braking phases of the arrest then it must be assumed that the
estimated loads may lie in error either way, and the maximum allowable speeds should be adjusted
accordingly. Errors in estimating the loads are discussed in Para 2 of this leaflet.

9.2.3 If the distance between runway edge sheaves and the energy absorbers, or the pull-out of the
alternative arresting gear, is significantly different from that of the arresting gear in which the aeroplane
has been tested then the forces caused by off-centre effects may be different and check tests are
recommended. The width of the runway (distance between sheaves) will also affect the steady braking
performance and the split distance (from sheave to energy absorber) will also affect the dynamic
performance of the arresting gear in a straight pull.

9.2.4 In some arresting gears it is possible for the dynamic and/or steady braking loads to be varied
considerably by minor modifications to the arresting gear energy absorbing devices. Where there is any
doubt about the forces caused by an engagement a check of the performance is recommended.

9.2.5 Where any of the criteria for read-across of Para 9.2.1 are not met brief check trials are strongly
recommended. If there are any doubts about the criteria of Para’s 9.2.2 and 9.2.3 check tests are
essential if full use is to be made of the facilities available with maximum safety.

9.3 PARACHUTE STREAMING

9.3.1 If a brake parachute is used as a means of providing deceleration on landing some tests to
assess its compatibility with the arresting gear may be necessary unless it can be shown by other means
that streaming before and after the engagement will not adversely affect the arrest.

9.4 THRUST REVERSERS

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9.4.1 Thrust reversers may also be used to provide landing deceleration and in this case, in addition to
the compatibility tests of Para 9.3, further consideration should be given to the effect on the temperature
environment to which the arresting system is exposed.

SYMBOLS

K Velocity of sound in the arresting cable - 3300 m/sec for steel wire rope
m Line density of arresting hook cable (kg/m)
V Engaging speed (m/sec)
Vr Take-off reject speed (kts)
WT Maximum design take-off mass (kg)
P/M Ratio of Peak to Mean arresting force
WA Aeroplane Mass (kg)
WO Aeroplane Mass (kg) for which an arresting gear is designed or
optimised.

REFERENCES

1 RAE - TR7027 Suggett G W -- The shape of arresting hooks for deck landing aircraft.

2 R&M2980 Thomlinson J - A study of the arresting hook bounce problem.

3 RAE - Tech Note Willis, Chisman, and Bullen - Measurement and suppression of
NA 204 tension waves in arresting gear systems.

4 RAE - TR70098 Randall T G - Performance analysis of the RAF Mk.1 Rotary

Hydraulic Arresting Gear.

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FIG.1 - TYPICAL HOOKLOAD ON A TIME BASE

FIG.2 - TYPICAL HOOKLOAD ON A DISTANCE BASE

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FIG.3 - GENERALISED ARRESTING GEAR PERFORMANCE AND LIMITATIONS DIAGRAM

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LEAFLET 11

INSTALLATIONS FOR EMERGENCY RECOVERY FROM STALL AND SPIN

GENERAL INFORMATION AND RECOMMENDATIONS

1 INTRODUCTION

1.1 This Leaflet gives information related to Section 3, Clauses 3.7.4 to 3.7.10

1.2 EMERGENCY RECOVERY INSTALLATIONS

1.2.1 It is necessary to provide an emergency stall and spin recovery device on aeroplanes undergoing
stalling and spinning trials for the following reasons:

(a) Although recovery from the stall is often achieved by the adoption of well established
techniques and the use of the normal flying controls, in some types of aeroplane configuration,
(e.g., those with T - tails) penetration of the stall can lead to stable flight at very high angles of
attack. In these circumstances recovery can be difficult, if not impossible, using only the normal
aeroplane controls.

(b) On aeroplanes which are not required to be recoverable from a spin, stalling
investigations may nevertheless lead to a spin condition.

(c) Although an aeroplane required to be recoverable from a spin may have acceptable
model spin recovery characteristics, scale effects throw some doubt on the full-scale interpretation
of the model tests. It is not always practicable to reproduce, in both model and full-scale tests, the
same range of spinning motions and the behaviour of the aeroplane within a given type of spin
may be quite different from that expected.

1.2.2 Unless the Project Team Leader is satisfied that the risk of entering a low airspeed flight
condition, from which recovery is unlikely, is sufficiently low to be acceptable, an emergency recovery
installation is required for stalling and spinning trials. This requirement may however be waived if special
circumstances (e.g., limitations on pressures for blown flap systems) require that the trials be made at
low altitudes where operation of a recovery installation would be unlikely to save the aeroplane.

1.2.3 No recommendations are made in respect of emergency recovery installations for use in
uncontrolled motions at high airspeeds where the loads imposed on the airframe are already severe; if
such an installation is considered necessary, each case will have to be individually assessed.

1.3 AIDS TO PILOTAGE

1.3.1 In circumstances in which the aeroplane could probably have recovered, accidents have
happened because the pilot failed to apply correct flight control techniques, usually for one or more of
the following reasons:

(a) The instruments fitted did not provide sufficient information for the diagnosis of the flight
condition and hence of the recovery technique required.

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(b) The pilot was unable to interpret the information correctly.

(c) The information supplied was misleading or unreliable.

(d) The motion was so violent that the pilot could not read the instruments or monitor the
recovery technique to assess whether it was being implemented successfully.

1.3.2 The requirements of Section 3, Clauses 3.7.5 and 3.7.6 cover the safeguards required to meet
such circumstances. Monitoring of flight conditions by ground observers, including a pilot with relevant
experience, is a particularly valuable safety measure. It has become common to fit telemetry on
aeroplanes undergoing spinning trials; greater use in trials exploring approaches to limiting angles of
attack is recommended.

1.4 SYSTEMS FUNCTIONING

1.4.1 Due to the flight conditions experienced, systems malfunctioning may occur. In particular, engine
surge or flame-out may be experienced, leading to degradations in the capabilities of electrical and
hydraulic systems. Augmentation of the normal emergency systems may be necessary to ensure
availability of the power supplies required but implications for production aeroplanes should be
considered where such augmentation has been applied.

1.4.2 To relieve the pilot's workload, information relevant to systems functioning should, wherever
practicable, be transmitted automatically to a ground station.

1.5 CIRCUMSTANCES JUSTIFYING RELAXATIONS

1.5.1 The requirement to fit any or all of the installations discussed above may occasionally be waived
when analogies with aeroplanes of proven satisfactory behaviour, whether of a different type or an
earlier variant of the same type, provide sufficient evidence that recovery will be achievable. Relaxations
may only be permitted when there is sufficient confidence that the scope of the investigations can be
confined to safe flight conditions and that the pilot can be provided with sufficiently reliable information
for the flight limitations to be observed. In assessing relaxations of safety provisions, adequate
considerations should be given to the consequences of stores carriage and to inertial effects.

1.5.2 As flight experience increases, the necessity for full implementation of all the safety provisions for
all trials on each individual aeroplane may diminish particularly where stall warning or stall prevention
devices have been fitted and proven; installation of an emergency recovery system may not be justified
for stalling trials within conditions already safely explored. Nevertheless, accidents have happened to
aeroplanes in investigations of what were believed to be minor and justifiable extensions of flight
envelopes or piloting techniques. Under these circumstances the advantages of telemetry are
considerable and its provision should be considered whether or not an emergency recovery installation is
fitted.

1.5.3 As development of an aeroplane proceeds, changes to its profile, mass distribution and control
system design may necessitate repeat stalling and spinning, trials. The standard of fit of safety
provisions for such trials should be agreed with the Project Team Leader.

2 ADVISORY INFORMATION

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2.1 DESIGN DATA

Note: Advice should be obtained from AD AS DMPS on the most suitable means of affecting stall or
spin recovery for a particular aeroplane.

2.1.1 Extensive model testing is necessary to obtain reliable data for the design of an emergency
recovery installation which will provide sufficient forces and moments to ensure recovery from the most
adverse flight conditions.

2.1.2 For aeroplanes for which the most adverse condition is the stable stall, the information required
should be obtainable from wind tunnel tests such as those reported in RAE TR 67197, but the case for
supplementary free-flight model tests should be considered.

2.1.3 For aeroplanes which may spin, wind tunnel investigations of aerodynamic characteristics at high
angles of attack are required but are not sufficient, and must be supplemented by tests on spinning
models. For aeroplanes for which spin recovery is a requirement, model spinning tests are specified in
Clause 3.7, more comprehensive investigation may be needed of the different behaviour patterns that
may be experienced in post-stall gyrations and spins, depending upon the entry conditions and the use
made of the flying controls before recovery is initiated. The flatter the spin the more difficult recovery may
be. Model tests should also include parachutes where possible.

2.1.4 It will normally be sufficient for the installation to be effective with flying controls free, or
alternatively with the controls nominally centralised, but with some allowance for departure from this
condition depending upon how well the pilot can be expected to maintain the neutral condition.

2.1.5 For parachute installations, determination of the wake characteristics behind the stalling or
spinning aeroplane will also be necessary (see Leaflet 12, Para 2.2.1).

2.2 INSTRUMENTATION

2.2.1 For spinning trials, additional instrumentation should be provided on a separate panel to enable
the pilot readily and reliably to determine the angle of attack, the altitude, the flying control positions, the
direction of spin, and whether or not the spin is inverted. Warnings of the attainment of critical altitudes in
relation to parachute streaming and abandonment of the aeroplane should be provided.

2.2.2 For stalling trials, reliable information on angle of attack and sideslip is necessary.

2.2.3 If it is necessary for the pilot to monitor engine or other systems conditions during the trials,
sufficiently prominent presentation of the critical parameters should be provided.

2.2.4 The above information, and any other data which will permit a team of ground observers to assist
the pilot, should be telemetered to a ground station for all trials in which problems may arise.

2.3 PROFILE AND MASS DISTRIBUTION

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2.3.1 Small changes in the profile of the aeroplane can have a significant effect on the behaviour of the
aeroplane in stall and spin. It is therefore important that any additional structure required to house the
recovery device should be contained within the existing profile of the aeroplane as far as possible. As
inertial effects can also be significant, the mass distribution of the aeroplane should be changed as little
as possible.

2.4 CHOICE OF INSTALLATION

2.4.1 In certain circumstances, a rocket installation may prove to be a better recovery device than a
parachute for the following reasons:

(a) In the deep stall, and in some spinning motions, the restoring moments required are not
provided very efficiently by a streamed parachute, which may have to be so large that the rear
fuselage structure has to be extensively modified to accommodate the installation.

(b) A rocket installation can be designed to provide the recovery moments needed more
effectively and without the drag and deceleration penalties of a parachute installation.

(c) There are potential dangers of the parachute damaging the airframe, fouling the tailplane
or impeding the escape of aircrew if its recovery action fails.

2.4.2 The disadvantages of a rocket installation include:

(a) Greater electrical complexity and the necessity for additional safety precautions.

(b) The need to identify unambiguously the direction of the spin and to provide alternative
firing directions for spins of different forms.

(c) The danger particularly with solid fuel rockets, of providing too violent a reaction in some
instances, e.g., in passage into an inverted steep stall or reversed spin.

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LEAFLET 12

INSTALLATIONS FOR EMERGENCY RECOVERY FROM STALL AND SPIN

PARACHUTE INSTALLATIONS

1 INTRODUCTION

1.1 This Leaflet gives information related to Section 3, Clauses 3.7.11 to 3.7.31 It deals with the use
of a parachute as an emergency recovery device. Reference should also be made to Leaflet 25.

2 ADVISORY INFORMATION

2.1 INSTALLATION

2.1.1 The most satisfactory type of installation is one in which the parachute is attached to the rear of
the fuselage, as far aft as possible, and on the plane of symmetry of the aeroplane. Installations using
two parachutes, one on each wing, should be avoided for the following reasons:

(a) The difficulty of co-ordinating the streaming and jettisoning devices.

(b) The asymmetric problems if one parachute should fail to deploy.

(c) The possibility of the two parachutes becoming entangled.

2.2 DEPLOYMENT AND INFLATION

2.2.1 The turbulent wake behind a stalling or spinning aeroplane tends to slow down or even prevent
parachute inflation and to reduce the drag forces which would otherwise be exerted. The extent of the
turbulent wake behind the aeroplane in stall or spin should therefore be estimated from model tests, and
the rigging lines and cable attaching the parachute to the aeroplane should be long enough to ensure
inflation in sufficiently undisturbed air.

2.2.2 To ensure that the parachute will inflate within the required 3 seconds, it may be necessary
forcible to eject the parachute, or an auxiliary parachute, by some means such as an ejection gun.

2.2.3 When an auxiliary parachute is used to deploy the main parachute, the cable between them
should be long enough to allow the auxiliary parachute to be ejected clear of the wake before the main
parachute leaves its container.

2.2.4 Model tests may be necessary to check that any parts of the installation which are jettisoned on
streaming the parachute will not interfere with the operation of the parachute or damage the aeroplane.

2.2.5 Satisfactory streaming of a parachute cannot be guaranteed by model tests; the type of
parachute to be used will be established by design calculations and comparison with other types.

Note: Advice should be obtained from a Design Approved Parachute Manufacturer.

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2.2.6 It is desirable that neither the streamed parachute nor the cable should foul any part of the
aeroplane. If fouling cannot be avoided, protection from adverse consequences should be provided e.g.,
by strengthening vulnerable structure and by adding guards to protect the flying controls.

2.3 JETTISON

2.3.1 Inadvertent streaming in flight could in certain circumstances lead to catastrophe; for instance, on
the approach or during take-off the unexpected sudden increase in drag could lead to loss of control. It is
therefore necessary to incorporate a design feature which ensures that if inadvertent release of the
parachute from its container occurs, no unacceptable load is applied to the aeroplane. It is
recommended that the feature should be such that the parachute and cable are not attached to the
aeroplane until the appropriate control is operated.

2.3.2 To guard against failure of the aeroplane structure in the event of the parachute being streamed
at excessive speeds it is necessary to incorporate an automatic jettison device, which is normally a weak
link, between the parachute and the structure. It is not possible to depend on the parachute itself failing
in such circumstances, as the scatter in parachute strength is so wide. If the ultimate strength of the
weak link is designed to be consistent with the proof strength of the aeroplane structure to which the
parachute is attached, the maximum drag possible without seriously damaging the aeroplane structure
will be available for use.

2.3.3 The scatter on the strength of the weak link should be as low as possible; the aim should be to
design the device so that its maximum failing load will correspond to the proof strength of the fuselage.

2.3.4 In some circumstances it may be possible to meet the requirement of Section 3, Clause 3.7.21
only by ensuring that the parachute is jettisoned during the escape sequence. It may therefore be
necessary to provide an auxiliary jettison system, which acts faster and more reliably than normal but
perhaps at the expense of damage to the airframe.

2.4 DESIGN OF THE PARACHUTE

2.4.1 Strength and drag aspects of brake parachute design are discussed in Leaflet 25 and should be
applied where relevant to the design of an emergency recovery parachute.

Note: Advice should be obtained from a Design Approved Parachute Manufacturer.

2.5 GROUND TESTS

2.5.1 A test should be devised to check the complete sequence of operations involved in the use of the
system. In particular, it should be demonstrated that the method of ejection adopted ensures that any
auxiliary parachute is thrown clear of the turbulent wake. The snatch load, the opening load and the
steady load should be simulated during the test.

2.5.2 The jettison device should be tested to ensure that it does not release the parachute under
snatch loads. A method of making this test is to use a section of the actual fuselage and the parachute
lines in order to reproduce the correct stiffnesses. The snatch load should be applied by dropping
weights acting through pulleys to give the correct directions to the loads to represent both flat and steep
spins, erect and inverted, and the stall.

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2.5.3 Normal functioning of the jettison device and satisfactory operation under low loads (e.g.,
parachute failed or incorrectly inflated) should be similarly tested to cover the various loads and angles
up to the maximum values which can occur in the spin or drive.

2.5.4 Tests should be made on sufficient numbers of the automatic jettison device to ensure that the
strength and scatter on strength meet the requirements of Section 3, Clause 3.7.24

2.5.5 The type of test rig and test methods recommended for brake parachutes in Leaflet 25 should be
used as a guide in the testing of emergency recovery parachutes.

2.6 FLIGHT TESTS

2.6.1 Various practical difficulties may arise in demonstrating that the requirement of Section 3, Clause
3.7.26 is met. The drag forces created by emergency recovery parachutes have increased so much in
recent years that it is doubtful whether an air towing facility could be provided for a large parachute.
Furthermore, it is not normally possible to carry out the towing test on the aeroplane to which the
installation is fitted because the strength of the weak link may be insufficient to withstand the drag force
at 1.3 Vp.

2.6.2 The use of a parachute of a well proven type could be sanctioned on the basis of design
calculations, by agreement with the Project Team Leader. In any event, if the towing flight test is not
done, it should be shown by calculation or other means that the parachute will meet the requirement.

2.6.3 The need for, and the timing of, the tests required by Section 3, Clause 3.7.31 will be decided
according to the circumstances of particular cases. As far as possible, before flight trials are undertaken
by an Experimental Establishment, the validity of design assumptions should be checked by analysis of
the full scale behaviour of the aeroplane during the contractor's flight trials, combined with measurement
of the drag (and any other relevant features) of the parachute obtained during the functioning trials of the
installation.

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LEAFLET 13

INSTALLATIONS FOR EMERGENCY RECOVERY FROM STALL AND SPIN

ROCKET INSTALLATIONS

1 INTRODUCTION

1.1 This leaflet gives information related to Section 3, Clauses 3.7.32 to 3.7.47 It deals with the use
of a rocket, or rockets, as an emergency recovery device. Subsequent references to rocket should be
interpreted as applying also to more than one rocket.

2 ADVISORY

2.1 DEPLOYMENT OF RECOVERY ROCKETS

2.1.1 A single axis rocket installation will not in all cases provide protection against all stall, super stall
and spin conditions. For aeroplanes whose layout is such that the main potential hazard is a stable super
stall it might be sufficient to provide a rocket installation giving only a nose down pitching moment, but for
an aeroplane which may need to be recoverable from a spin it will be necessary to provide rocket
installations giving yawing moments in either direction, e.g., by the use of a motor thrust vector control,
which would also enable the line of thrust to be corrected after recovery from the spin.

2.1.2 It is possible that the pilot may not be sure of the attitude or spin direction of the aeroplane.
Either, instrumentation to show the pilot which rocket motor to fire, or preferably automatic rocket motor
selection, should be provided.

2.2 CHOICE OF ROCKET MOTOR

2.2.1 Either solid or liquid fuel rockets may be used for a recovery rocket installation. In general, a solid
fuel rocket will result in the simpler installation, but a liquid fuel rocket has the advantage of a controllable
thrust characteristic, enabling more control to be exercised over the progress of the recovery and so
avoid the danger of entry into an adverse situation in the reverse sense (see also Leaflet 14, Para 2.4). A
liquid fuel rocket may also permit the fuel supply to be cut off in case of fire. The adverse effects of
rockets continuation to burn during the later stages of recovery from the stall should be considered.

2.2.2 It is preferable to use an existing rocket motor (e.g., missile boost motor), but its suitability should
be fully discussed with AD AS DMPS, DOSG for Safety and Suitability for Service and other groups such
as AWAC, ASD etc. which would look at the Aircraft Self Damage and Integration. In particular the
strength of an existing rocket motor should be checked to see that it is satisfactory for aeroplane use,
particularly after all burnt time.

2.3 THRUST CHARACTERISTICS

2.3.1 The type of rocket installation and the magnitude and duration of rocket thrust required for
recovery should also be determined by using the results of the tests discussed in Leaflet 11. If solid fuel
rockets are used, it may be necessary to use a bank of them to give a variable thrust characteristic. The
firing drill should be developed with the aid of an aeroplane simulator. Both the thrust characteristics and
the firing drill should be agreed with the Project Team Leader.

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2.4 SAFETY CONSIDERATIONS

2.4.1 As live rockets should be carried only on those flights on which stalling or spinning trials are to be
carried out, the design of the rocket installation should be such that the rockets and fuel can be easily
and quickly installed in the aeroplane.

2.4.2 The firing circuits should be such that they can be checked on the ground and in the air, and the
rockets subsequently made live immediately before the stalling or spinning trials are started, and made
safe after the completion of the trials and before descent for landing.

2.4.3 Fire risks which should be considered include those due directly to the hazardous nature of the
propellant and the initiation system and those due to hot gases and debris ejected during burning. The
effects on the installation of any local heating (e.g., from a jet pipe) should also be taken into account.

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LEAFLET 14

TARGET TOWING INSTALLATIONS

DEFINITIONS & GLOSSARY

1 INTRODUCTION

1.1 The simplest form of target towing installation may comprise only a single point
attachment/release unit on the aeroplane with the towed assembly arrayed on the ground alongside the
aeroplane. With increasing complexity the installation may comprise a ram air turbine powered winch
capable of streaming and recovering a target on several kilometres of tow line.

2 DEFINITIONS

2.1 Towing System

2.1.1 Discardable - Where the entire towed assembly (target plus tow line) is discarded from the towing
aeroplane on completion of the towing sortie prior to landing.

2.1.2 Recoverable - Where the entire towed assembly can be wound back on to the towing aeroplane
and re-stowed on completion of the sortie prior to landing.

2.2 Tow Length

2.2.1 Fixed - Where the streamed tow length is predetermined and cannot be varied or altered by any
command from the towing aeroplane.

2.2.2 Variable - Where the streamed tow length is under the control of the towing aeroplane and can
be varied or altered on command.

2.3 Deployment Period

2.3.1 Short - Where the deployment and/or recovery phase is of sufficiently short a period that it may
be allowed to run to completion without compromising the aircrew's options for alternative action if
necessary.

2.3.2 Long - Where the deployment and/or recovery phase lasts for a protracted period such that the
aircrew's options for alternative actions is compromised and hence may require halting of the
deployment/recovery phase or even an immediate shedding of the towed assembly.

2.4 Deployment Arrest

2.4.1 Unrestrained - Where little or no attempt is made to moderate the rate of streaming of the tow line
and target assembly and in consequence high shock loads are imposed on the towing installation during
the deployment phase.

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2.4.2 Restrained - Where the rate of streaming of the towline and target is moderated, or controlled, by
some energy dissipating/absorbing device such that the deployment phase does not impose significant
shock loads on the towing installation.

2.5 Fixed Installation - That part(s) of the installation which remains fitted to the aeroplane during the
towing sortie. May be fitted internally or may be an external store.

2.6 Towed Assembly - That part of the installation which is streamed from the aeroplane during the
towing sortie and including any discardable item thereof. Initial disposition may be as an internal store,
external store, or picked up from the ground.

2.7 Launch Mode

2.7.1 Ground Launch - Where the towed assembly is laid out on the ground adjacent to the aeroplane
and attached to the aeroplane just prior to take-off. The aeroplane takes off pulling the already streamed
towed assembly behind it. (A variation of this technique is where a low flying aeroplane snatches the laid
out towed assembly from the ground).

2.7.2 Air Launch - Where the towed assembly is carried by the aeroplane in a stowed condition and is
streamed from the aeroplane on command whilst airborne.

3 GLOSSARY OF TERMS USED IN TARGET TOWING OPERATIONS

STREAMING ) The dispensing of the tow line from the stowed

TRAILING ) condition.

WINDING IN ) Gathering of the tow line and/or target back on to the

RECOVER(Y) ) parent aeroplane.

RETRIEVAL ) Gathering of a tow line and/or target from DZ after dropping

from the aeroplane.

LAUNCH Initial movement of tow line/target from the aeroplane with

intention of taking in tow.

DROP ) Final release of towed assembly from the aeroplane, in

RELEASE ) normal mode on to a DZ.

JETTISON ) The emergency release of the towed assembly, underwing

store etc., from the aeroplane.

STOWED Initial state of installation on the aeroplane.

Final state of recoverable equipment.

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TOW ) The state of the towed assembly when fully streamed and
IN TOW ) following the aeroplane's manoeuvres.
STEADY TOW )
STATIC TOW )
SHEDDING An action of relieving the aeroplane of a load.
DISCARD The action of shedding a store, tow line or target.

CARRIAGE Flight of the aeroplane with the towed assembly in the

stowed condition.

DZ Dropping Zone - Prepared area on to which a discardable

towed assembly is dropped for subsequent retrieval.

CARRIER An intermediate assembly between the main store and the

aeroplane structure.

SNATCH The condition of accelerating a stationary (or substantially

lower velocity) towed assembly up to the airspeed of an
already moving towing aeroplane.

DEPLOYMENT The entire action from initial launch until final state of target
in tow.

DEVELOPMENT The action of a fabric target unfurling into its full

shape.

STOP Halting the streaming of a variable length towline at

designated length or an intermediary position on command.

ARREST The automatic halting of an unrestrained towline at full

length.

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LEAFLET 15

TARGET TOWING INSTALLATIONS

GENERAL AND OPERATIONAL REQUIREMENTS

1 PROXIMITY OF ENGINES

1.1 Impingement of Jet Efflux or Propeller Slipstream.

1.1.1 Impingement of jet efflux or propeller slipstream on the fixed assembly should be avoided. Where
impingement is not entirely avoidable the adverse effects should be minimised. Problems may arise in
trying to launch a target into a turbulent airstream.

1.1.2 Similarly impingement of jet efflux or propeller slipstream on the towed assembly should be
avoided where possible. However, the very nature of towing an object behind an aeroplane will often
place the towline, and occasionally the target itself, in the engine wake. Suitable design, or protection, for
operation in a hot and/or turbulent environment will need to be considered.

1.1.3 Ground launched systems require attention to be paid to the combined effect of aeroplane
attitude and power during take-off.

1.1.4 With air launched systems the proximity of the trajectory of the streaming towed assembly
together with the final towed attitude should be examined to reduce any adverse effects. Heat protection
of the portion of the towed assembly closest to the aeroplane may be required.

1.1.5 The use of re-heat during any towing phase should be examined and its effect on the towed
assembly considered.

2 LOSS OF CONTROL

2.1 Uncontrolled situations which need consideration include, but are not limited to:-

(a) Failure of the target towing system to decelerate, or stop, at the appropriate streaming or
recovery points.

(b) Acceleration beyond normal running velocities in either streaming or recovery phases.
Likely to be experienced in combination with (a).

(c) Effects of excessively high running velocities on mechanisms.

(d) Disintegration of rotating components in overspeed conditions and consequent

centrifuging of debris.

(e) Whiplashing by severed tow line in:-

(1) Close proximity to aeroplane exterior.

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(2) Winch.

(3) Interior of aeroplane.

(f) Unpredicted behaviour of damaged target.

3 COLOUR SCHEMES

3.1 Training - Where target towing aeroplanes are dedicated to that role and, in particular, are used
for the training of other units or services it is advisable that the aeroplane be painted with a
distinguishable pattern.

3.2 Operational - Where the target towing is undertaken by an operational aeroplane as a secondary
role then the aeroplane should retain its operational colour scheme.

4 COMPATIBILITY OF ENVIRONMENTAL SPECIFICATIONS

4.1 It should be remembered that just as the fixed installation may be released to a lesser
specification than the aeroplane likewise the target may be released to a still lesser condition. Under
such circumstances the carriage condition needs particular attention.

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LEAFLET 16

TARGET TOWING INSTALLATIONS

AERODYNAMIC AND FLYING QUALITIES

1 INTRODUCTION

1.1 This leaflet gives information related to Section 3, Clauses 3.8.11 to 3.8.16

2 STATE OF AEROPLANE

2.1 Certain parameters, such as mass, moments of inertia, centre of gravity position, wing swing or
thrust setting, may vary over a range of values during a flight phase as a consequence of target trailing,
recovery, release etc. The Contractor should define, for consideration as discrete states, a limited
number of values of these (or similar) parameters including the most critical values and the extremes
encountered during the flight phase concerned and consider the case where such changes occur in
accordance with the correct operation of the system.

2.2 In deciding which aeroplane states must be examined, it will in some cases be necessary to
consider effect due to mass distribution which are not symmetrical caused by or associated with external
store complements but asymmetrical mass distribution in other circumstances should not be overlooked.
Where relevant, the values of such moments should be included in the definition of the aeroplane state.

3 PERFORMANCE

3.1 Variations in performance as the consequence of the operation of target towing equipment fitted
to an existing aeroplane should be considered, quantified and quoted in the relevant aeroplane aircrew
documentation.

4 HANDLING

4.1 Aeroplane handling characteristics should not inhibit flying to the qualities of Levels 1 and 2 -
limiting parameters being assessed during flight testing (Part 1, Section 2, Clause 2.1).

4.2 Consideration should be given to the variation in drag during the various phases of towed target
flights - this variation being in addition to the increase in drag created by any fitted external store.
Variation will be apparent during the following flight phases (For category and phase definitions, see Part
1, Section 2, Clauses 2.1.17 and 2.1.18 and Section 2 leaflet 1).

(a) Category C. TO phase - increasing drag during ground snatch of a towed assembly.

(b) Category B. CL, CR and D phases - varying drag (with airspeed) during the towing of a
target.

(c) Category A. WD phase - increasing drag caused by the streaming of a towed assembly.
This drag value will vary with the tow length and will be instantaneously increased if a packed
target deploys as the end of the towline.

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(d) Category A. WD phase - decreasing drag caused by the recovery of a towed assembly.
This drag will vary with the tow length and be instantaneously decreased if the towed assembly is
released.

4.3 Consideration should also be given to the variation in handling characteristics caused by Lateral
and longitudinal stability, flutter, assembly, asymmetric roll etc., etc., as a result of the towed target
installation and operation.

5 FAILURE CASES

5.1 The effects of a failure of components/systems on the aeroplane flying qualities should be
considered. The critical case can be grouped into two categories i.e.:-

(a) Target system failure.

(b) Aircraft system failure.

5.2 Target system failure - considerations should include the inadvertent streaming of a stowed
target assembly during take-off or landing; the failure to release a target when commanded; the flying
qualities of a damaged target and the failure to recover a streamed target.

5.3 Aircraft system failure - considerations should include the sudden aircraft power loss (asymmetric
and/or symmetric); the failure of rudder effect; the failure of cockpit indicators.

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LEAFLET 17

TARGET TOWING INSTALLATIONS

LOADING AND SHEDDING

1 INTRODUCTION

1.1 In general installations will be of such bulk and/or weight to justify handling the major constituents
as discrete items, with the towline and towed body being loaded/attached separately onto the already
fitted fixed installation.

1.2 With small, simple installations it may be possible to treat the entire installation, including towline
and towed body, as a readily attachable/detachable store i.e., expendable decoy devices.

1.3 Shedding of towlines is an airborne operation to rid the aeroplane of the towed assembly or
towline remnant under circumstances which may range from routine to emergency.

2 LOADING OF TOWLINES AND TARGETS

2.1 For preference it should be possible to fit and remove towlines without removal of the fixed
installation from the aeroplane.

2.2 Towlines may be wound directly to and from the fixed installation.

2.3 Towlines may be carried on reels or trays etc., and in such cases it is permissible to fit and
remove the towline complete with its carrier.

2.4 Where the towline is stowed in the fixed installation it should be possible to attach/detach the
target from the towline at the launching position.

3 EXAMPLES OF TOWLINE SHEDDING DEVICES AND TECHNIQUES

3.1 Operator Commanded

(a) Electro/Mechanical Release Unit.

(b) Explosive Powered Cutter.

(c) Mechanical Guillotine.

(d) Controlled Streaming of Entire Towline.

3.2 Automatic Response

(a) Shear Pin.

(b) Specific Tensile Link.

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(c) Overload Triggered Release Unit.

(d) Release Unit or Cutter Actuation by Overspeed or Overtension Detection Circuits.

4 SHEDDING CIRCUMSTANCES

4.1 The circumstances under which towed assemblies may require to be shed are categorised
below:-

(a) As a routine operation or completion of the towing task where the towed assembly is
expected to be re-usable after retrieval from the DZ.

(b) As a routine operation or completion of the towing task where the towed assembly is
expendable and not retrieved.

(c) As an alternative to full recovery where the state of the towed assembly or other
considerations make recovery inadvisable.

(d) As an emergency operation (i.e., JETTISON) when it becomes imperative to relieve the
aeroplane of the towing load as quickly as possible, i.e. loss of control or abandonment of the
towing aeroplane.

5 SHEDDING CONDITIONS

5.1 Where the shedding of the towed assembly is also a normal function of the target towing system
the jettison system may be a similar, even identical, device provided that independent supplies are used.

5.2 Duplex release units may be used providing that their design allows that no single failure of a
constituent component prevents the operation of the release.

5.3 Where duplex release units are used one 'half' may be used as a means of jettison.

5.4 Where necessary weak links may be incorporated into the attachment of towed assemblies to
protect the aeroplane from abnormal structural or aerodynamic loadings.

5.5 Weak links may take the form of towing links incorporating shear pins or short lengths of specific
strength cable at the aeroplane end of the towed assembly.

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LEAFLET 18

TARGET TOWING INSTALLATIONS

COCKPIT CONTROLS AND INDICATORS

1 INTRODUCTION

1.1 The minimum controls required by target towing installations are defined in Section 3, Clause
3.8.40. Where space permits specific control panels (complete with their indicators) should be provided.

1.2 It is appreciated that some target towing aeroplanes will be adaptations of aeroplanes designed
for other purposes. The siting of cockpit controls may then be limited to choosing which of the existing
controls are most suitable for use.

1.3 Controls required for the operation of the streamed target are not listed here but their inclusion
should be borne in mind during the design stage.

1.4 Although it is argued that with automatic or pre-set systems it may not be essential to inform the
aircrew of the state of the tow the aircrew's confidence and trust is greatly increased by feedback of
data. It is recommended that explicit information be displayed whenever possible, particularly to allow
discrimination between normal and critical conditions.

2 RECOMMENDED INDICATORS

2.1 No Indication - Where operational restraints dictate discardable, short, fixed tow length systems
(e.g., combat decoys) can be operated without cockpit indicators although confirmation that the device
has been activated will be of value to the aircrew. When any indicator is omitted then further
consideration should be given to the reliability of the target towing system to justify the assumption that it
will react predictably.

2.2 Tow Length - Recoverable and/or variable tow length systems require progressive streamed tow
length indication. This will be reversible on recoverable systems.

2.3 Tow Line Tension - Progressively indicating the tension in the tow line. When compared to
predicted values can confirm state of the tow i.e., intact, target lost, or icing. Complementary to "tow
length" and can be used to monitor the streamed state following the loss of length indication.

2.4 Velocity - Indicating rate of streaming or recovery of tow line. Essential on manually controlled
systems. Desirable on automatic systems where it confirms normal velocity state, acceleration and
deceleration points and warns of over running and over speeding situations.

2.5 Direction - Indicating in which direction the tow line is moving. On long tow systems may indicate
response of system before it can be perceived by "tow length" indication.

2.6 Stowed - Confirming that the towed assembly is correctly loaded and available for use. In a fault
mode indicating that the towed assembly has not left the aeroplane as expected.

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2.7 Secured - Confirming that the towed assembly is correctly attached and capable of being taken in
tow. May function as the inverse of "discarded".

2.8 Launched - Confirming that the target has been launched from the aeroplane and implying that
the tow line is streaming. May function as the inverse of "stowed".

2.9 Discarded - Indicating that the towed assembly is not attached to the towing point. Normal
indication that a discardable towed assembly has been dropped. On air launch systems warns against
launching an unattached tow. May function as the inverse of "secured".

2.10 Recovered - Confirming that towed assembly has been recovered back on to the aeroplane. May
function in combination with "stowed".

Fixed Tow Variable Tow

Discardable Recoverable
GL AL AL
Normal
1. Launch and Stream - / /
3 Stop - - /
3 Recover - - /
4 Discard / / /
(Routine for discardable systems,
alternative for recoverable systems)
Emergency
5 Jettison Tow (Pilot's selection) / / /
6 Jettison Tow (Ejection synchronised) / / /
7 Jettison External Inst. - / /
(where permitted - Pilot's selection)

TABLE 1 - REQUIRED CONTROLS

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Fixed Tow Variable Tow

Discardable Recoverable
GL AL AL
1 Stowed - / /
2 Secured / / -
3 Launched - / /
4 Recovered - - /
5 Discarded / / /
6 Tow Length - - /
7 Tow Tension - - /
8 Velocity - - /
9 Direction - - /

TABLE 2 - RECOMMENDED INDICATORS

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LEAFLET 19

REDUCTION OF VULNERABILITY TO BATTLE DAMAGE

GENERAL REQUIREMENTS

1 GENERAL (See also Ref 1)

1.1 The survivability of an aircraft in battle conditions depends on its Susceptibility (Detection,
Acquisition, Tracking, and Threat Avoidance), Vulnerability and Recoverability.

1.2 The vulnerability of an aircraft is a measure of how it can withstand an attack from a warhead
burst or a projectile impact. This is usually expressed as a probability of kill given a hit (PkjH) within a
defined time period. This time period starts when the weapon / projectile first interacts with the aircraft
and ends at the time associated with the particular kill category under consideration. Kill category time
frames can vary between a few seconds to several days or weeks and may therefore include
considerations of recoverability. An example of a short duration kill category is if a bomber aircraft is on
its approach run and bomb release is about to occur. An attack at this stage would lead to a mission kill
within a few seconds. A long duration example is where an aircraft has been damaged in combat and
has returned to base, but requires extensive repairs before it can be made operational. In this situation,
the aircraft might be out of operation for several days or even weeks. One of the most relevant kill
categories is where the attacking weapon / projectile causes the aircraft to loose the ability to maintain
controlled flight within a specified time of weapon / projectile interaction. A definition of the three common
UK kill categories is shown in Para 5 below. Identifying the correct kill category time-frame is essential
because it has a significant influence on achieving an aircraft design with minimal vulnerability. For
example, fire damage and fuel tank leakage effects will not contribute to a flight control kill category
within 15 seconds of weapon / projectile interaction (denoted F15 sec) because there is insufficient time
for any significant fire damage or fuel loss to occur. However, if the kill category time-frame is increased
to 20 minutes (denoted F20 min), then both fire damage and fuel leakage affects could now make a
significant vulnerability contribution. This could lead to a kill of the aircraft unless vulnerability reduction
techniques are included in the aircraft design to reduce the effects, in this case, by including fire
suppression and fuel tank self-sealing.

2 DESIGN AIMS

2.1 The objective of the design aim is to set a standard for flight control, structural integrity and
operation of all systems. The primary concern is with flight-critical systems but consideration may also be
given to the vulnerability of mission-critical systems. The general strategy for the reduction of
vulnerability to battle damage is based on:

(a) Design for tolerance of battle damage at component level.

(b) Aircraft layout to minimise weapon effects (e.g. component separation).

(c) System/Component redundancy.

(d) Damage control equipment (e.g. fire & explosion suppression and fuel tank self-sealing).

(e) Shielding by non-flight-essential components, armour or other structure.

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(f) Reduction of airframe vulnerability.

2.2 The Design Organisation should consult with the PT leader and establish whether the
aircraft has specific vulnerability issues and, if so, in what way(s). Regarding design compliance,
vulnerability modelling and analysis should be used to manage vulnerability within the context of
the overall survivability requirement. Vulnerability assessment and analysis shall be a primary
design requirement at the outset. Work should be done to define and specify the threats and their
effects. This would provide a consistent measure for assessing the compliance of the aircraft.

2.3 The vulnerability analysis is an iterative process, the results of which will show how far the
design meets the requirements of Section 3, Clause 3.9 However, the results of the analysis will
be influenced by trade-off effects, and the impact of these results on the design will be influenced
by the time and money available for re-design after completion of the analysis. To minimise
redesign issues, the vulnerability analysis process should be commenced as early as possible in
the design process.

2.4 The principal design aims shall be to:

(a) Maximise the probability that no single threat effect will degrade the flying qualities
of the aircraft below Level 3 of Def Stan 00-970 Part 1, Section 2, Clause 2.1.19

(b) Minimise the overall aircraft PkjH, i.e. to minimise the probability of the aircraft
being killed in relation to the particular kill category (see Para 5)

3 PROTECTION MEASURES

The following is a list of primary measures which should be considered, but there may be others. For any
particular project some will be more important than others and some may be omitted. The order of
priority will be determined by the Vulnerability Analysis. The primary concern is with flight-critical
systems, but consideration may also be given to mission-critical systems. There is a need to identify
flight-critical equipment and components, while mission-critical ones may also be identified. This can be
done by considering the aircraft’s appropriate flight-critical or mission-critical component failure logic, and
identifying which particular equipment and components are required for survival.

3.1 FUEL SYSTEM

(a) Design of components to tolerate battle damage.

(b) Continuance of supply.

(c) Protection against leakage, explosion and fire (e.g. self-sealing, ESF, fire
suppression)

(d) Low volatility fuels. (Note some RPAVs have petrol engines).

3.2 POWER SYSTEMS (electrical, mechanical, hydraulic, etc.)

This should include localised electro-mechanical and electro-hydraulic control actuators.

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(a) Design of components to tolerate battle damage.

(b) Redundancy of components (prime movers, generators and wiring)

(c) Use of fire-resistant materials. (For example, low flammability hydraulic fluid. Such
fluid should be used in the whole hydraulic system, i.e. in the hydraulic
components of the undercarriage equipment, as well as the hydraulics in the
aircraft control system).

(d) Prevention of mechanical jamming from weapon / projectile damage, i.e. jam
proofing.

(e) Over-load sensing systems to redirect electrical power.

(f) Leak sensing hydraulic systems.

3.3 FLYING CONTROLS SYSTEMS

This should include flight control surfaces, as well as flight control computers and their ability to deal with
aircraft configuration changes due to battle damage.

(a) Design of components to tolerate battle damage.

(b) Redundancy of components (e.g. control cables, control rods and end-fittings,
control surface hinges, flight control computers, electrical / optical cable runs.)

(c) Use of fire resistant materials (e.g. titanium or steel control rods rather than
composite ones)

(d) Prevention of mechanical jamming from weapon / projectile damage, i.e. jam
proofing.

(e) Leak sensing hydraulic systems.

3.4 PROPULSION SYSTEM

This should include VSTOL drive-trains (including swivelling nozzles) and propeller systems, as
well as FADECS.

(a) Design of components to tolerate battle damage.

(b) Redundancy of propulsion system components.

(c) Separation and protection of engines and/or vital components. (This includes
designs for the containment of engine debris (considered to be an onboard
energetic source (see Section 3, Clause 3.11)) as well as the containment of
engine fire. On a twin-engine aircraft, an armoured and thermally resistant firewall
in between the engines could be a suitable solution.

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(d) Protection of ancillaries (e.g. armour or non flight-critical systems)

3.5 LIFE-SUPPORT SYSTEMS

These systems should be included for fixed and rotary-wing aircraft. Consideration needs to be
given to the oxygen supply to the pilot, the temperature control of the cockpit and cockpit air
conditioning, particularly in terms of the system performance when parts of the system are
damaged by weapon / projectile attacks.

(a) Design of components to tolerate battle damage.

(b) Redundancy of vital equipment / components.

(c) Separation of vital equipment / components.

(d) Protection of ancillaries, e.g. armour or non flight-critical systems.

3.6 AVIONICS & COCKPIT SYSTEMS

This includes flight-critical avionics and cockpit systems.

(a) Design of components to tolerate battle damage.

(b) Redundancy of systems, equipment and components.

(c) Separation of vital equipment / components.

(d) Protection of ancillaries, e.g. armour or non flight-critical systems.

(e) Ability to operate in reversionary modes.

3.7 CREW (See also Leaflet 20)

(a) Redundancy. (Refer to Section 3, Clause 3.10)

(b) Close armour protection.

3.8 ROTOR AND TRANSMISSION SYSTEMS

This includes rotorcraft rotor and transmission systems and transmission systems for driving
aircraft lifting fans.

(a) Design of components to tolerate battle damage.

(b) Redundancy of components.

(c) Separation of vital components.

(d) Protection of ancillaries, (e.g. armour or non flight-critical systems)

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(e) Ability to operate in reversionary modes. For example, a pitch-bias unit could
maintain stability of a helicopter’s tail rotor when controls are severed.

3.9 STRUCTURE

(a) Design to tolerate battle damage.

(b) Design for repair to battle damage.

(c) Modular Construction.

(d) Maximisation of accessibility.

(e) Reduction of airframe vulnerability by consideration of all design possibilities:

(1) Fail safe/multi load path structure.

(2) Material selection.

3.10 ON-BOARD MONITORING SYSTEMS

These systems identify and display system performance information to inform the pilot that the
aircraft has a reduced functional capability and/or flight envelope.

(a) Identification of what type of damage can be sensed.

(b) Automatic interpretation of sensor data to identify the effect on the aircraft’s flight
envelope.

(c) A straight-forward and unambiguous display of information is required for the pilot.

3.11 ONBOARD ENERGETIC SOURCES

These include rotating engines, pressurised bottles, accumulators, tyres, etc. These debris
sources should be identified along with all vulnerable flight-critical systems around them.

The design should consider the containment of the debris generated from these sources, or, the
protection to prevent damage to singularly-vulnerably flight-critical systems in the debris path.

4 AIRCRAFT BATTLE DAMAGE REPAIR (ABDR)

The Design Organisation should provide a damage assessment methodology as well as

instructions for repair and proving.

Design to tolerate battle damage

4.1 When designing for battle damage repair the designer should consider not only how to restore full
airworthiness when the aircraft is back at base but also how to get the aircraft airborne and back to base
following a forced landing.

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4.2 The designer should also consider the information to be put into the (ABDR) Repair Manual(s).
After combat damage, a decision will have to be made about each damaged area in the structure and
systems, as to whether or not it must be repaired immediately. The manuals should therefore contain the
information necessary for such decisions which will normally, but not always, be made by aircraft
maintenance and repair staff.

4.3 The designer should incorporate the ability for self-diagnostics testing to identify equipment faults
and malfunctions that could help to pin-point battle damage, thereby aiding repair. Design features such
as removable panels, line replaceable units, and cable and pipe line identification, should be included to
assist with the battle damage repair task.

4.4 Flight system designs with respect to degraded capability should ensure the ability to power-up
and fly aircraft with degraded flight and mission systems functionality which may result following rapid
partial ABDR repairs (e.g. ability to over-ride Build In Test no-go warnings)

5 UK KILL CATEGORIES

F(t)- Within time (t) following the damaging strike the aircraft will become permanently
incapable of controlled flight (periods of (t) are normally 0, 15 secs, 5 mins, 20 mins, 30
mins)

C(t)- Within time (t) following the damaging strike the aircraft will become unable to perform the
stated mission (periods of time (t) are normally 2, 5 and 30 secs).

E(t)- The aircraft receives damage which will keep it grounded for repairs for time (t). The
preferred periods for assessment purposes are 8, 24, 48 hours and infinity (i.e. write off).

Note: The times quoted are typical of those used in comparative studies. Times used for
a particular analysis will depend on the type of project and the mission profile
under consideration.

6 US KILL CATEGORIES

These have been taken from Ref 2 and are given in Table 1 in this leaflet with the UK Kill Categories for
comparison.

7 FURTHER INFORMATION

7.1 Where the information contained in this Leaflet is not sufficient for the designer's purpose,
particularly with regard to the Vulnerability Analysis, and with regard to reports or further advice on
Biological and Chemical effects, reference should be made to the Weapon System PT Leader in the first
instance.

7.2 American information on Design for Survivability and Vulnerability Reduction is contained in the
MIL-HDBKS and MIL-STDS listed below. The Standards contain requirements and definitions. The
Handbooks contain background information and acceptable methods of compliance.

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REFERENCES

1 MIL-STD-2069: (S/S by Mil-HDBK-2069) Aircraft Survivability

2 MIL-STD-2089: (S/S by Mil-HDBK-2089) Aircraft Survivability Terms.

3 RAE TR 79123 - September 1979 - by R.G.E. Mallin.

4 MIL-HDBK-268(AS) 5 Aug 82: Survivability Enhancement, Aircraft, Conventional

Weapon Threats, Design and Evaluation Guidelines.

5 MIL-HDBK-336-1, 2 and 3. 25 October 1982; 26 August 1983; 31 January 1983;

Survivability, Aircraft, Non-Nuclear, VOL 1 - General Criteria, VOL 2 - Airframe,
VOL 3 - Engine.

6 AGARD - CP 212 - P 218396 - Design for reduction of aircraft vulnerability - by Lt

Col R.T. Remers USAF.

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Definition American British

(MIL-HDBK-2089 Para (RAE)
5.3.3.2)
Aeroplane is lost from the Attrition kill F(t)
inventory by:
- disintegrating immediately KK - kill t = 0 sec
on being hit (Catastrophic kill)
- falling out of control

- in 15 seconds - t = 15 sec
- in 30 seconds K - kill -
- in 5 minutes A - kill t = 5 min
- in 20 minutes - t = 20 min
- in 30 minutes B - kill t = 30 min
- before completing C - kill -
mission

Unable to complete stated Mission abort kill C(t)

mission within time (t) - t = 2 sec
following a strike - t = 5 sec
- t = 30 sec
Unable to complete part of Mission limiting -
mission condition
Mission completed but repairs Mission available kill -
required before next mission
Special Rotorcraft Category Forced landing kill -
in which damage or a warning
required a forced landing

Damage causes aeroplane to Repair time kill

miss its next scheduled - E(t)
mission - t = 8 hrs
- t = 24 hrs
- t = 48 hrs
t = Infinity
(write-off)
Grounded for Repair. Damage E - kill
which makes structural damage -
on landing probable (e.g. a burst
tyre)
V - kill
VSTOL aeroplane only. -
Damage which causes the loss of
vertical operational ability.
NUCLEAR
No appreciable damage -
normal turnaround Sure safe -
Mission cannot be completed
but continued controlled
flight is possible Mission kill -

Damage which causes aeroplane

to fall immediately out of
control Sure kill -

TABLE 1 - BRITISH AND AMERICAN KILL LEVELS

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LEAFLET 20

PROTECTION OF AIRCREW AGAINST CONVENTIONAL WEAPONS

GENERAL REQUIREMENTS

1 GENERAL

1.1 This Leaflet contains non-mandatory requirements, background information, and advice
on the provision of protection systems for aircrew. It covers those parts of the requirements that
are not already incorporated in Section 3, Clauses 3.9 and 3.10

2 DESIGN

2.1 The protection of aircrew stations and personnel against the designated threats for the
aircraft will be achieved by suitable design of appropriate protection systems.

2.2 The following protection systems should be considered:

(a) Increased separation of crew members.

(b) Shielding by structure or non-critical components.

(c) Armour protection or fragment suppression materials.

2.3 The systems selected and the amount of protection provided, will be determined by the
Vulnerability Analysis of Section 3, Clause 3.9

3 ARMOUR

3.1 The effectiveness of armour protection for a particular crew member will be improved, or
the mass penalty reduced, by the following:

(a) Location of the armour as close as possible to the crew member.

(b) Location and shaping of the armour to provide protection for more than one crew
member and/or for vital equipment at the same time.

(c) Use of armour as structure.

(d) Use of removable armour as special role equipment.

3.2 The effectiveness of material for armour protection will be influenced by the following:

(a) Mass and bulk of the material.

(b) Deflection characteristics.

(c) Spall characteristics.

(d) Multi-hit capability.

(e) Durability.

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(f) Ease of replacement.

(g) Cost and availability.

(h) Armour shape options (e.g. flat, 1-D or 2-D curved panels) or flexible armours.

(i) Ease of mounting the armour (within or on the aircraft)

3.3 Armour may be provided as:

(a) body armour worn by the crew member,

(b) armoured seats with integral or bolt-on armour,

(c) armour added to or integrated into floors, sidewalls bulkheads and instrument
panels,

(d) transparent armour where visibility is required, and

(e) external armour attached to the airframe in the vicinity of aircrew.

3.4 Where body armour and structural armour are both provided, the designs should be
integrated to eliminate gaps and overlaps.

REFERENCES

Reference 1: ASCC Air Standard 61/102/14A “Armour Protection for Aircrew”

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LEAFLET 21

PROTECTION FROM THE EFFECTS OF NUCLEAR EXPLOSIONS, LASER WEAPONS, CHEMICAL

AND BIOLOGICAL WARFARE AGENTS

DEFINITIONS

1 INTRODUCTION

1.1 The following definitions apply to terms used in Section 3, Clause 3.11 and in the associated
Leaflets concerning the Nuclear, Laser and Chemical and Biological threats respectively.

2 DEFINITIONS – GENERAL

2.1 DEFINED THREAT EFFECTS - A threat list will be generated for each air weapon system and
concept of operation.

2.2 MISSION - The task to be performed during a sortie.

2.3 MISSION - ESSENTIAL SYSTEM - A system that is essential to the successful completion of an
air mission.

2.4 PROBABILITY OF OCCURRENCE - A function of three factors:

(a) Probability of encountering a particular threat.

(b) Threat Lethality.

(c) The susceptibility of the aircraft to the threat.

2.5 SORTIE - An operational flight by one aircraft.

2.6 SPECIFIED THREAT EFFECTS - Those Threat Effects specified in the aircraft Specification or
by the Project Team Leader.

2.7 SYSTEM - An aggregate of hardware, software, and man that satisfies a specific end-use
function.

2.8 THREAT EFFECT - The definition of a threat in terms of those physical characteristics which
affect aircraft design.

2.9 THREATS - Those hostile elements of an environment which could reduce the ability to an
aircraft, its systems, and crew to perform its mission.

3 DEFINITIONS - CHEMICAL AND BIOLOGICAL THREATS

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3.1 BALANCED CHEMICAL HARDENING - A concept of system chemical hardening in which all
sub-systems have been brought to a commensurate hardness level in the sense that no one link (sub-
system) in the system chain is overhardened, nor is any link in the chain of significantly greater
vulnerability than the rest of the sub-systems.

3.2 BIOLOGICAL AGENT - A micro-organism which causes disease in man, plants, or animals or
causes the deterioration of material.

3.3 BLISTER AGENT (VESICANT AGENT) - A chemical agent which injures the eyes and lungs, and
burns or blisters to the skin (for unprotected personnel). BLISTER AGENTS cause a variety of damage
effects to materials, surfaces and finishes unless these are appropriately hardened.

3.4 BLOOD AGENT - A chemical compound including the cyanide group, which affects bodily
functions (which may lead to death) by preventing the normal transfer of oxygen from the blood to body
tissues (for unprotected personnel). BLOOD AGENTS do not cause any material damage; they are a
hazard to unprotected personnel only.

3.5 NERVE AGENT - A chemical agent which injures / incapacitates the eyes and lungs, and can
cause incapacitation or death via absorption through unprotected skin. Nerve agents cause a number of
effects, including affecting the transmission of nerve impulses, leading to paralysis. At very low dosages,
nerve agents will cause miosis (eye effects) which will impact the ability to maintain control of the aircraft
or to complete the mission (for unprotected personnel). NERVE AGENTS cause a variety of damage
effects to materials, surfaces and finishes unless these are appropriately hardened.

There are other classes of CW agent.

3.6 CHEMICAL AGENT - A chemical substance which is intended for use in military operations to kill,
seriously injure, incapacitate man through its physiological effects. Excluded from consideration are riot
control agents, herbicides, smoke and flame.

3.7 CHEMICAL HARDENING - Measures taken during the design and construction of military
equipments to avoid damage to materials caused by CW agents and to reduce or eliminate the hazard to
personnel arising from the presence of chemically contaminated surfaces. Associated with chemical
hardening is DECONTAMINABILITY, and good hardening practices will also facilitate decontamination.

3.8 CHEMICAL OPERATIONS - Employment of chemical agents to kill, injure, or incapacitate for a
significant period of time, man or animals, and deny or hinder the use of areas, facilities or material; or
actions to defend against such employment.

3.9 CHEMICAL SURVEY - The directed effort to determine the nature and degree of chemical
hazard in an area and to delineate the perimeter of the hazard area.

3.10 CHEMICAL SURVIVABILITY - The capability of a system to withstand hostile chemical warfare
environments(s) without suffering loss of its ability to accomplish the designated mission(s)

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3.11 CHEMICAL VULNERABILITY - The characteristics of a system which cause it to suffer

degradation (i.e. impact the capability to perform its designated mission) as a result of having been
contaminated in a hostile chemical warfare environment.

3.12 CONTAMINATION - The deposition and/or absorption of radioactive material, biological, or

chemical agents (in liquid or vapour form) on, into and by structures, areas, personnel, or objects. See
also induced radiation; residual radiation.

3.13 CONTAMINATION CONTROL - Procedures to avoid, reduce, remove or render harmless,

temporarily or permanently nuclear, biological and chemical contamination for the purpose of maintaining
or enhancing the efficient conduct of military operations.

3.14 CONTAMINATION DENSITY - The amount of liquid chemical agent contamination per unit area
of surface usually expressed in g.m -2.

3.15 DECONTAMINANTS - Materials or processes employed for the purpose of promoting the
removal, dissolution, dilution or destruction of chemical or biological warfare agents from contaminated
surfaces or assemblies.

3.16 DECONTAMINATION - The process of removing, destroying, neutralising, making harmless

chemical or biological agents, or by removing radioactive material from contaminated surfaces or
assemblies.

3.17 DESORPTION TIME - The time taken for the concentration of chemical agent vapour evolved
from a contaminated surface to decrease to a safe level.

3.18 FILTRATION - The process of actively reducing chemical agent liquid or vapour or biological or
radioactive particulate matter in the atmosphere by the passage of the contaminated air through a
suitable military specification filter.

3.19 HERBICIDE - A chemical compound which will kill or damage plants.

3.20 PENETRATION TIME - The time taken for a CW agent vapour or liquid to penetrate through a
specified thickness of material to produce a hazardous condition, usually expressed as a vapour dosage,
on the other side.

3.21 PERMEABILITY - A measure of the susceptibility of solid materials to penetration by CW agent

Liquids.

3.22 PERSISTENCY - An expression of the duration of effectiveness of a chemical agent which is

dependent upon the physical properties of the agent, the method of dissemination, the weather
conditions and the characteristics of the contaminated surfaces.

3.23 POLYMERIC MATERIAL - The general description of all types of polymers including paints,
thermoplastics, thermosets, elastomers, rubbers and transparent coatings.

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4 DEFINITIONS - NUCLEAR THREAT

4.1 AIR BURST - The explosion of a nuclear weapon in the air, at a height greater than the maximum
radius of the fireball.

4.2 BALANCED NUCLEAR HARDENING - A concept of a system nuclear hardening in which all
sub-systems have been brought to a commensurate hardness level in the sense that no one link (sub-
system) in the system chain is overhardened, nor is any link in the chain of significantly greater
vulnerability than the rest of the sub-system.

4.3 BALANCED SURVIVABILITY - Survivability is balanced when the following conditions are
satisfied:

(a) All sub-systems and components have approximately equal survivability for each specific
nuclear environment.

(b) The entire system is not vulnerable to one or more environmental effects, while having
adequate survivability for all other associated effects.

4.4 ELECTRICAL ISOLATION - Separation of electrical circuits, signals, or data to preclude

ambiguity, interference, or information perversion. This may be achieved through physical isolation or by
any property which distinguishes one electrical signal from all others (for example, time, phase,
amplitude, or frequency).

4.5 ELECTROEXPLOSIVE DEVICE (EED) - Any electrically initiated explosive device within an
electroexplosive sub-system having an explosive or pyrotechnic output, and actuated by an
electroexplosive initiator.

4.6 ELECTROMAGNETIC COMPATIBILITY (EMC) - The ability of electronic equipment sub-systems

and systems to operate in their intended operational environments without suffering or causing
unacceptable degradation because of unintentional electromagnetic radiation or response.

4.7 ELECTROMAGNETIC INTERFERENCE (EMI) - Any electromagnetic disturbance which

interrupts, obstructs, or otherwise degrades or limits the effective performance of electronic/electrical
equipment. It can be induced unintentionally, as a result of spurious emissions and responses,
intermodulation product, and the like.

4.8 ELECTROMAGNETIC PULSE (EMP) - A sharp pulse of radio frequency (EMR) produced when
an explosion occurs in an unsymmetrical environment especially at or near the earth’s surface or at high
altitude.

4.9 ELECTROMAGNETIC RADIATION (EMR) - Radiation made up of oscillating electric and

magnetic fields and propagated with the speed of light. Includes gamma radiation, X-Rays, ultra violet,
visible and infra red radiation, and radar and radio waves.

4.10 ENDO ATMOSPHERIC - Within the atmosphere, altitude less than 35 km.

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4.11 EXO ATMOSPHERIC - Outside the atmosphere, altitude greater than 35 km.

4.12 FAIL-SAFE - A design feature of a nuclear weapon system/component which ensures that, under
failure, no critical functions, damage to equipment, or injury to personnel will occur.

4.13 FIREBALL - The luminous sphere of hot gasses which forms a few millionths of a second after a
nuclear explosion. The fireball is a result of the absorption by the surrounding medium of thermal X-Rays
emitted by the extremely hot weapon residues which are at several tens of million degrees.

4.14 FISSION - The process whereby the nucleus of a particular heavy element splits into (generally)
two nuclei of lighter elements, with the release of substantial amounts of energy. The most important
fissionable materials are uranium-235 and plutonium 239; fission is caused by the absorption of
neutrons.

4.15 FLUENCE - Fluence is the time integrated flux i.e. the number of particles per unit area.

4.16 FLUX - Flux is the number of particles (photons) per unit area per second.

4.17 FUSION - The process whereby the nuclei of light elements, especially those of the isotopes of
hydrogen, namely deuterium and tritium, combine to form the nucleus of a heavier element, with the
release of substantial amounts of energy.

4.18 GAMMA RADIATION - High frequency electromagnetic radiation with a very short wavelength
(10-11 to 10-14m) emitted from atomic nuclei, and accompanying many nuclear reactions. Physically
gamma rays are identical with X-Rays of high energy, the essential difference being that X-Rays do not
originate from atomic nuclei, but are produced in other ways, like slowing down fast electrons of light
energy.

4.19 GROUND BURST - The explosion of a nuclear weapon at the surface of the earth or at a height
above the surface less than the radius of the fireball at maximum luminosity (in the second thermal
pulse).

4.20 INITIAL RADIATION - Radiation produced by a nuclear explosion within 1 minute following the
burst. It includes neutrons and gamma rays given off at the instant of the explosion, gamma rays
produced by the interaction of neutrons with weapon components and the surrounding medium, and the
alpha, beta, and gamma rays emitted in the fission products and other weapon debris during the first
minute following the burst.

4.21 MISSION COMPLETION - Mission Completion kill category is the level of response of an aircraft
corresponding to a damage level between incipient and catastrophic damage. The aircraft should be just
able to accomplish its assignment satisfactorily.

4.22 NEUTRON (symbol n) - A neutral particle with no electrical charge. Its mass is 1.00867 atomic
mass units or 1.67492 x 10-27 kg. It is present in all atomic nuclei except those of ordinary (light)
hydrogen. Neutrons are required to initiate the fission process, and large numbers of neutrons are
produced by both fission and fusion reactions in nuclear (or atomic) explosions.

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4.23 NUCLEAR FALL-OUT - The settling of the radioactive debris resulting from anatomic explosion.

4.24 NUCLEAR HARDENING - Nuclear hardening is the employment of any technique which
circumvents or mitigates the effects of an adverse nuclear environment, that is, which improves nuclear
survivability.

4.25 NUCLEAR RADIATION - Particulate and electromagnetic radiation from atomic nuclei in various
nuclear processes. The important nuclear radiations, from the weapons standpoint, are alpha and beta
particles, gamma rays and neutrons. All nuclear radiations are ionizing radiations, but the reverse is not
true; X-rays for example, are included among ionizing radiations, but they are not nuclear radiations
since they do not originate from atomic nuclei.

4.26 NUCLEAR SURVIVABILITY - The capability of a system to withstand a nuclear environment

without suffering loss of its ability to accomplish the designated mission.

4.27 NUCLEAR SURVIVABILITY CRITERIA - The specific nuclear environmental level used to define
the nuclear survivability required of a given system.

4.28 NUCLEAR VULNERABILITY - The characteristics of systems which cause it to suffer

degradation (i.e. in the capability to perform successfully its designated mission) as a result of having
been subjected to a given level of a hostile nuclear environment.

4.29 RADIATION DOSE - The total amount of ionising radiation absorbed by material or tissue,
commonly expressed in Grays (or sub-multiples thereof) and indicating the absorbing material, e.g.,
cGy(Si).

4.30 RESIDUAL RADIATION - Nuclear radiation, chiefly beta particles and gamma rays, which
persists for some time following a nuclear explosion.

4.31 SLANT RANGE - The direct distance between an explosion and a target, as opposed to the
horizontal distance between ground zero and a target which is the ground or surface range.

4.32 SURE-KILL - Sure-kill is the level of response corresponding to a catastrophic damage condition
which results in essentially immediate loss of the aircraft.

4.33 SURE-SAFE - Sure-safe is that level of response which in no way affects mission completion and
allows the aircraft to return home or to an alternate base in an essentially undamaged condition.

5 DEFINITIONS - LASER THREAT

5.1 BANDGAP ENERGY - A material which has an energy of absorbsion from a ground to an excited
state.

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5.2 DAZZLE - Dazzle is the temporary degradation of an optical or electro-optical system due to the
scattering of radiation beyond the geometrical image. The affect ceases either immediately the laser
radiation is removed or once the eye/system recovers from any after image.

5.3 FLUENCE - Fluence is defined as the amount of energy (J) applied at the input aperture of an
optical system or at the surface of a structure per unit area (cm2). Fluence is a measure of Energy
Density (J/cm2).

5.4 IN-BAND - A laser system is said to be in-band when it operates at a wavelength which is
transmitted by the optics and is within the sensitivity band of the target sensor.

5.5 MEAN POWER DENSITY - This is the product of laser pulse energy per unit area and the pulse
repetition rate.

5.6 OUT-OF-BAND - A laser is out-of-band if it operates at a wavelength which the optics of the
target absorbs rather than transmits the energy.

5.7 PLASMA - The generation of a very high temperature gas or mixture of gas and vapour which
consists of ionized atoms and free electrons due to high levels of laser radiation vaporising materials or
structure of the target.

5.8 POWER DENSITY - Power per unit area applied to a target by a CW laser (W/cm2)

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LEAFLET 22

PROTECTION FROM THE EFFECTS OF NUCLEAR EXPLOSIONS, LASER WEAPONS, CHEMICAL

AND BIOLOGICAL WARFARE AGENTS

NUCLEAR WEAPON EFFECTS ON AIRCRAFT

1 INTRODUCTION

1.1 This leaflet gives information on the effects of a nuclear explosion on an aircraft and provides
methods which may be used to evaluate the hardness of the aircraft, and avionic systems and guidance
on how this hardness may be improved. For a detailed description of the characteristics of a nuclear
explosion see any number of published sources such as 'Handbook for Analysis of Nuclear Weapons
Effects on Aircraft'. DNA 20841 Volume 1. Definitions of the terms used are given in Leaflet 21.

1.2 The limitations on the capabilities of aircraft to deliver conventional stores are imposed by such
factors as range and payload and enemy defensive actions. With the emergence of nuclear weapons the
enhanced nuclear yield has placed the aircraft in danger of being damaged by the weapon it has
delivered. Additionally consideration of the effects of the nuclear explosion on parked Aircraft and also
the conditions under which Aircraft may be shot down or killed by the effects of nuclear weapons needs
to be given. It is necessary therefore to construct sure-safe, mission completion, and sure-kill envelopes
for the aircraft under consideration.

1.3 The effects of the weapon characteristics on the aircraft or its crew include velocity (gust) effects,
overpressure effects, thermal radiation effects, initial radiation effects, residual radiation effects and
combinations of these features.

2 NUCLEAR WEAPON EFFECTS ON AIRCRAFT

2.1 GUST EFFECTS

2.1.1 As an aircraft is engulfed by a blast wave it encounters changes in material velocity, pressure,
temperature and density. Hence angle of attack, sideslip, dynamic pressure and Mach number change,
causing changes in the aerodynamic loads acting on the aircraft.

2.1.2 If the aircraft is intercepted by a blast wave from below, then eventually a translational velocity
will ensue as the aircraft responds to the gust. Ultimately the increment in loading will reduce below the
initial value caused by the encounter. The highest loadings are likely to occur soon after the interception
in a very dynamic manner that depends strongly on the characteristics of the aircraft and flight
conditions; "riding with the gust" may limit the duration of the loading.

2.1.3 In general the change in aerodynamic conditions during an encounter with a blast wave will
cause moments on the aircraft which will result in angular accelerations. A stable aircraft will rotate into a
gust once it is fully immersed, but the initial response may be different and will depend, for example, on
the direction of the blast and whether the aircraft has tail or canard surfaces. This initial response may be
particularly marked where an aircraft is overtaken by a blast wave, where there may be a significant time
difference between, for example, the change of lift on the wing and on the pitch control surfaces.

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2.1.4 Severe loadings are required in order to produce damage corresponding to a sure-kill condition
and these loadings can generally occur only if the gust induced angle of attack or sideslip is large and
often well beyond the angle for which linearity can reasonably be assessed. However, if the aircraft is
manoeuvring when it encounters a blast wave, the additional load needed to produce damage will be
reduced.

2.1.5 The change of aerodynamic loading will in general cause dynamic structural deformations of the
fuselage and aerodynamic surfaces, which will modify the loading and response of the aircraft because
of the associated local velocities and displacements.

2.1.6 As the deformation of a structure increases beyond the onset of buckling its load carrying
capability decreases and may become sufficiently low that a sure-kill condition exists.

2.1.7 Although it is unlikely that the autopilot would be in engagement under 'battlefield' conditions, it
should be noted that an autopilot which is maintaining constant barometric altitude could react violently
to the change in pressure accompanying the blast wave. An autopilot would normally disengage
automatically under such circumstances but the response of active control systems and automatic terrain
following systems should be considered.

2.1.8 Parked Aircraft may also be damaged by bending of the fuselage or vertical tail due to
aerodynamic loading of the vertical tail. For sure-safe conditions, this bending will be elastic- for sure-kill
conditions, inelastic response may also be important.

2.2 OVERPRESSURE EFFECTS

2.2.1 Overpressure influences smaller elements of the structure such as a skin, the stringers, and the
frames, particularly pressure on the fuselage. When an aircraft is struck by a blast wave, the pressure on
the side of the fuselage facing the burst point is increased above the incident value by reflection, and a
local loading of short duration is generated. As the blast wave continues to engulf the aircraft, the
pressure on the side of the fuselage facing the burst point decays to the pressure behind the blast wave.
The characteristic loading, then, is a high reflected pressure (from two to eight times the overpressure
associated with the blast wave) which decays very rapidly, in a few milliseconds, to the value of the
pressure behind the blast wave. This high pressure short-duration pulse is then followed essentially by
the much longer duration, but lower pressure, pulse characteristics of the blast wave.

2.2.2 It is primarily the high reflected pressure, short-duration pulse which is responsible for damage to
skin panels, stringers and frames. These structural elements are vulnerable to such short duration
loadings because of their high frequencies. For the converse reason, the much lower frequency major
components are influenced very little by the short-duration loading.

2.2.3 The short duration pulse produces dishing-in of skin panels and buckling of stringers and frames
or portions of stringers and frames. As in the case of analysis of gust effects, analysis of overpressure
response need consider only elastic response for the sure-safe case, but should properly include
inelastic response for sure-kill conditions.

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2.2.4 Overpressure damage is generally the predominant effect for parked aircraft. For in flight aircraft,
however, overpressure damage is usually of minor importance in comparison with gust effects.
Overpressure becomes important for in-flight aircraft chiefly in those regions for which gust effects are
small; namely, for bursts almost directly in front or directly behind the aircraft.

2.3 THERMAL RADIATION EFFECTS

2.3.1 The radiant exposure of an aircraft in-flight varies widely with atmospheric conditions, orientation
of the aircraft with respect to the burst, aircraft velocity, the ground reflecting surfaces, and clouds.
Reflection adds and scatter may add to the direct radiation, and under some circumstances the thermal
energy incident on an aircraft in space may be two or three times the direct thermal energy computed at
a given slant range. Conversely, when a heavy cloud layer is between the burst and the aircraft, the
radiant exposure may be only a fraction of the predicted value for a given range. In other situations,
reflected radiation from clouds may contribute significant thermal energy to areas of the aircraft shaded
from direct radiation. During weapon effects tests of an aircraft flying in a cloud above the burst, the
radiant exposure at the top of the aircraft and its cockpit area was observed to be as much as one-fourth
of the direct radiation on the lower surfaces. This experiment demonstrated the need for protection of
weapons delivery aircraft from radiant exposure from any direction.

2.3.2 The motion of the aircraft during the time in which significant thermal radiation is being emitted by
the fireball can exert a very important influence on the thermal radiation incident upon the aircraft.

2.3.3 The absorptivity of the aircraft metal skin and the angle of incidence of the thermal radiation
affect the amount of energy absorbed by the structure: the boundary layer in the air flow adjacent to the
structure leads to convective cooling. Very thin skins are rapidly heated in damaging temperatures,
because the energy is absorbed by the skin much more rapidly than it can be dissipated by conduction
and convective cooling. The reduction of aircraft vulnerability to thermal radiation may be achieved by
coating materials with low absorptivity paints, by eliminating ignitable materials from exposed surfaces,
and by substitution of thicker skins for very thin skins.

2.3.4 An irradiated aircraft thin metal skin panel, supported by internal structure which is usually much
cooler, may be heated to a point where it may be badly buckled by thermal stresses or melted. For a thin
skin in either case, there will be essentially no temperature variation through the thickness. A step higher
in complexity is the thick skin case which involves a temperature distribution across the thickness of the
skin. A still more complex temperature distribution occurs in built-in structures, with air gaps acting as
insulators between spars, stringers, and skin.

2.3.5 Thin skins of CFC structure may be subject to surface damage and delamination because of the
high thermal gradient produced through the skin due to the low thermal conductivity of CFC. Thermal
stressing may not be a problem due to the low thermal expansion of CFC.

2.3.6 Biological injury of the crew from intense thermal radiation and damage to non-structural
elements which would adversely affect mission performance should also be considered when dealing
with thermal criteria. In many cases, these problems can be minimized by adequate protective measures
such as reflective coating applied to transparencies.

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2.3.7 Aircraft inadequately hardened may sustain permanent damage at very low thermal levels as a
result of ignition of rubber and plastic items, fabric seals, fixed landing gear aircraft tyres, cushion and
headrest covers.

2.3.8 Aircraft painted with dark paint are especially vulnerable to thermal radiation damage because
the dark painted surfaces absorb three to four times the thermal energy that is absorbed by polished
aluminium surfaces or surfaces protected with reflective paint.

2.3.9 Consideration should be given to the smoke/toxic emission characteristics of materials used in
crew compartments to ensure that the application of specified thermals levels will not result in any
significant degradation in crew performance.

2.3.10 For small yields, thermal radiation is generally of secondary importance for both parked and in
flight aircraft; thermal radiation may be dominant for both for high-yield weapons.

2.3.11 Sections 4, 5, and 6 of Ref 3 covers Material Properties relevant to the Thermal Radiation
Environment, Temperature Rise and Distribution, and Thermal Radiation Effects respectively and
constitute a 'designers guide' relative to these aspects.

2.4 NUCLEAR RADIATION EFFECTS

2.4.1 The vulnerability of aircraft to nuclear radiation effects depend upon the effect of nuclear radiation
on the crew, electronic gear, special weapons, or instruments which may be carried by the aircraft. The
important radiation consists of gamma rays and neutrons emitted during a brief period after the nuclear
explosion; both forms of radiation travelling significant distances through air capable of producing
harmful effects in living organisms and electronic components.

2.4.2 EMP can damage or cause malfunction in aircraft electric circuits, cables and electronic
components. EMP fields couple to aircraft by direct penetration through electrically poor joints in the
aircraft skin, diffusion through CFC, aerials, and through ports such as transparencies and hatches. The
impact of EMP upon an aircraft is a system dependent phenomenon and each aircraft electrical and
electronic system must be assessed independently.

2.5 RESIDUAL RADIATION EFFECTS

2.5.1 Residual radiation can be a problem to both equipment and personnel. Aircraft which will traverse
a nuclear dust cloud will receive an immersion dose due to the decay of the radioactive material in the
cloud. In addition, if air outside the aircraft is taken into the aircraft to cool electronics (or for other
purposes) then the equipment inside the aircraft may receive a cockpit/cabin dose as well.

2.5.2 Fallout which has been deposited on an airbase will contribute to the radiation dose received by
the aircraft while they are parked. The response of the equipment to the residual gamma radiation will be
the same as for the initial radiation and dependent upon the total dose received.

2.5.3 As with equipment, residual radiation can be a concern to aircrew members during flight or
ground operations, however it is unlikely that dose to the aircrew will be sufficient to have an impact on
the success of the mission. The crew will require an air filtration unit to remove particulate radioactive
matter when flying through nuclear debris clouds or ground manoeuvring in a fallout region.

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3 NUCLEAR HARDNESS EVALUATION OF AIRFRAME AND EQUIPMENT

3.1 The nuclear weapon effects considered in this paragraph are airblast, thermal pulse, initial
nuclear radiation, electromagnetic pulse (EMP). The various weapon effects are those immediately
external to the aircraft, equipment, or personnel. The degree of hardening to reduce nuclear vulnerability
with respect to specific weapon yields will be a compromise between possible cost and performance
penalties on the one hand and improved serviceability on the other. In calculating the hardness of an
aircraft structure it is necessary to determine and compare the hardness for each nuclear weapon effect
i.e. airblast, initial radiation, thermal radiation and EMP effects.

3.2 EVALUATION OF AIRBLAST EFFECTS

3.2.1 When an aircraft is parked or flying in the vicinity of a nuclear blast it may be subject to both
dynamic and static overpressure effects. The dynamic effects result in gusts which in turn apply loads to
wings, tailplane, and vertical aerodynamic surfaces.

3.2.2 GUST ANALYSIS - AERODYNAMIC COEFFICIENTS - To evaluate the gust loading it is

necessary to calculate the relevant aerodynamic coefficients before and during the nuclear explosion.
Ref 1 Appendix B Section B 1.1 Parts A and B, provides a method of calculating the aerodynamic
coefficients for Wing, Horizontal Tail, and Vertical Tail surfaces.

3.2.3 GUST ANALYSIS-IN-FLIGHT Ref 1 Appendix B, Sections B2.1 and B2.2 illustrate two methods
of analysis:

(a) The analysis given in B2.1 is based upon determining the load factor produced on the
aircraft during the blast encounter, accounting roughly for the fact that this load factor is
dynamically applied, and comparing the resultant effective load factor with the critical load factor.
The parameters associated with this analysis are flight attitude, weapon yield, AUW at time of
interest, pre-blast aircraft velocity, pre-blast load factor for straight and level flight (n1=1), upload
and download limit load factor (N), and wing plan forms.

(b) An alternative method which is more involved is given in B2.2. This analysis is based
upon determining the bending moment produced by the blast encounter at each of the relevant
positions on the aircraft and comparing them with the corresponding critical value. The static
bending moment is calculated using the inertia factors and airloads and the result multiplied by an
approximate dynamic factor in order to estimate the maximum bending moment which will be
experienced. The dynamic factor is a composite factor which represents the alleviation of the loads
due to rigid body translation ("riding" with the gust) and the overstress due to structural dynamic
effects.

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(c) Critical bending moments are required for sure-safe and/or sure-kill conditions. For
suresafe, the critical bending moments are based upon design limit conditions but for sure-kill, the
critical moments are based on design ultimate conditions and a lethal ratio. The lethal ratio is
determined from a simple representation of post failure response of a single degree of freedom
system. Apart from the specific geometry of the aircraft under consideration, the following
parameters are required for the analysis; flight attitude, weapon yield, gross AUW, true airspeed,
pre-blast load factor for straight and level flight, fundamental wing bending frequency (rad/sec),
fundamental fuselage vertical and horizontal bending frequencies, fundamental tail vertical and
horizontal bending frequencies, and aircraft mass distribution.

3.2.4 GUST ANALYSIS - PARKED AIRCRAFT

(a) Aircraft with low wing loadings are particularly susceptible to this type of loading which in
its parked position may have the aerodynamic surfaces set at a large angle of attack relative to the
ground. Head on encounter presents the most severe loading condition for this case. Skidding is
possible whenever the drag, coupled with lift overcomes the frictional forces between ground and
tyres. As sudden skidding damage criteria is dependent on the distance of the aircraft from other
objects it is recommended that no analysis be undertaken to quantify this case. Blast approaching
from the rear of the aircraft will induce negative lift and produce downward forces on the wheels
which may damage landing gear and main supporting structure. Vulnerability studies of parked
aircraft have found that tie-downs are not very effective in reducing motion induced damage for
encounters with high strength blastways. In the method of analysis referenced below, the forces
exerted by any tie-downs which might be present have been neglected.

(b) A method to calculate the Gust-Effects on Vertical Tail and Rear Fuselage for parked
aircraft is given in Ref 1, Para 4.2, Method 2, Part A. The analysis given is based upon determining
the bending moment produced by the blast encounter at two positions on the aircraft and
comparing each of these bending moments with the corresponding critical value. The static
bending moments derived from the application of airloads is multiplied by an approximate dynamic
factor which accounts for the fact that the loading is dynamically applied, in order to estimate the
maximum bending moment that will actually occur. Critical bending moments are required for sure-
safe and/or sure-kill conditions.

(c) Ref 1, Para B4.2, Method 2, Part B provides a method of analysis to determine the
dynamic pressure required from a head-on gust to lift the aircraft to such a height that upon impact
with the ground, the landing gear will deflect a critical amount. This critical deflection is related to
the design sinking speed for normal landing. The dynamic pressure thus obtained is taken to
define the sure-safe condition. In the sure-kill condition, a similar procedure is followed but the
critical deflection is replaced by a new value which is related to a higher sinking speed. The basic
parameters required for the analysis are; weapon yield, aircraft gross AUW, angle of attack of
aircraft in parked position relative to a horizontal head wind, design limit sinking for landing, vertical
difference between aircraft C of G at lift-off and aircraft C of G position in parked position and
aircraft geometry.

(d) Ref 1, Para B4.2, Method 2, Part C indicates a method to obtain sure-kill gust envelopes
for aircraft in the parked position when the mode of damage is crushing of the landing gear.

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3.2.5 OVERPRESSURE ANALYSIS - Overpressure effects primarily influence secondary structural

elements such as, panels, stiffness, frames, canopies, radomes, etc, in contrast to gust effects which
influence deformations of primary structural components such as wings, fuselage, and horizontal/vertical
tail units. Skin panels may yield or rupture; longerons, stringers and frames may fail by compressive
yielding or local buckling. The fuselage is generally the most susceptible to these types of damage and
hence the overpressure analysis is more relevant to this area of the aircraft. Ref 1, Chapter C1 gives
methods of analysis for assessing the overpressure effects on fuselages and a more sophisticated
method is also referenced with no restriction to fuselage structures.

3.3 EVALUATION OF THERMAL RADIATION EFFECTS

3.3.1 The response of the aircraft to the incident thermal energy exhibits itself as a temperature rise in
the aircraft skin. Several parameters influence the magnitude of this temperature rise; the most important
being skin thickness, material, surface condition, cooling air flow over skin surface etc; radiation of
thermal energy to the atmosphere and conduction of nuclear energy to the atmosphere.

3.3.2 Sure-safe conditions are based on an allowable temperature rise of the aircraft skin. Melting of
the aircraft skin is a requirement for a sure-kill situation. In this case, the temperature rise of the skin is
followed to the melt temperature, and further temperature and further heat input is necessary to produce
melting. Typical calculations of the envelope that defines the sure-safe and sure-kill regions with respect
to thermal radiation on aircraft in-flight or parked are given in Ref 2 (Capabilities of Nuclear Weapons
Part II, Chapter 13, Page 60)

3.3.3 For sure-safe the criteria is assumed to be the skin panel temperature value which produces a
20% reduction in the Modulus of Elasticity when applied to the thinnest structural skin the fuselage. For
each burst orientation, skin panels located in the following regions should be considered:

(a) For a burst directly below the aircraft, the lower surface of the fuselage within 45° of the
normal to the bottom of the fuselage.

(b) In a burst directly above the aircraft, the upper surface of the fuselage within 45° of the
normal to the top of the fuselage.

(c) In a burst directly to the side of the aircraft, the side surface of the fuselage not covered
by (a) and (b) above.

Aircraft improperly prepared may sustain serious damage at very low thermal levels as a result of ignition
of items such as rubber and/or fabric seals, fixed landing gear tyres, cushions and headrest covers etc.
Aircraft painted with dark paint are especially vulnerable to thermal radiation damage, because the dark
painted surfaces absorb three to four times the thermal energy that is absorbed by polished aluminium
surfaces protected with reflective paint.

3.3.4 Ref 3 Section 7.1. (Thermal Data Book D/DP(N)21/5/17) describes and quantifies the thermal
radiation effect for various classes of materials based on the results of both atmospheric weapon tests
and simulation trials.

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3.4 INITIAL RADIATION

3.4.1 These forms of initial radiation can have deleterious effects on the electrical and electronic
systems of the aircraft and if the dose is large enough can incapacitate the aircraft. In addition, the initial
radiation can have injurious effects on aircraft components such as plastics, rubber, fuels, lubricants, and
can reduce considerably the effectiveness of a nuclear weapon being carried by the aircraft.

3.4.2 Ref 1 (Handbook for Analysis of Nuclear Weapon Effects on Aircraft Vol 2), Appendix E,
Paragraph E1.1 provides a method of analysis for the determination of sure-safe and/or sure-kill nuclear
radiation envelopes based on specific gamma dose rate criteria (applicable to avionics in general).

3.5 EMP

3.5.1 Degradation of system performance may occur as a result of functional damage or operational
upset in which its performance is only impaired temporarily. Electronic components that are sensitive to
functional damage or burnout are given in Ref 2, (Capabilities of Nuclear Weapons Part II), Part II,
Chapter 9, Para 58.

3.5.2 The nature of a circuit has a strong bearing on the transients that cause damage; however, in
general, pulse lengths of microsecond and sub microsecond duration are required to cause problems.
Table 9-27 of Ref 2 Part II, Chapter 9 Para 59 shows a list typical of common active devices and the
approximate energy required to cause functional damage. The minimum energy required to damage
meters or ignite fuel vapours is about the same as that required to damage semiconductors. The energy
level associated with operational upset is typically 10 to 100 times less than that which is required to
damage sensitive semiconductor components.

3.5.3 A general approach to the analysis of a system with regard to its EMP vulnerability should include
the following steps;

(a) Identify susceptible subsystems and components.

(b) Determine the level of energy which will cause damage.

(c) Estimate exposure to coupled EMP energy.

(d) Estimate protection which may be provided by such devices as filters and clamps.

(e) Estimate level of damage and the effect on the ability of the system to carry out its
function.

3.5.4 Information on the Coupling Mechanisms into Equipment, Electrically Short Cables in Incident
EMP Fields, Coupling to Long Cables and Lines, and Aerials and Pseudo-aerials is given in Para’s 5.2,
5.3, 5.4, and 5.5 of Ref 4)

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REFERENCES
1 Handbook for Analysis of Nuclear Weapons Effects on Aircraft, Vol I and II, DRIC No P 243939
and 243940.

2 Capabilities of Nuclear Weapons (U) DNA EM-1(N) Part II, DRIC Acc No P205160.

3 Thermal Data Book D/DP(D)21/5/17.

4 Def Stan 08-4 (Part 4)

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LEAFLET 23

PROTECTION FROM THE EFFECTS OF NUCLEAR EXPLOSIONS, LASERWEAPONS, CHEMICAL

AND BIOLOGICAL WARFARE AGENTS

GENERAL RECOMMENDATIONS - CHEMICAL AND BIOLOGICAL WARFARE AGENTS.

1 INTRODUCTION

1.1 This leaflet gives recommended or acceptable methods of meeting certain of the basic
requirements stated in Section 3, Clause 3.11

1.2 Chemical and biological weapons are primarily anti-personnel / area denial weapons. They are
intended to kill or incapacitate or to enforce the adoption of protective measures which degrade military
efficiency. Chemical hardening is the use of designs and materials which resist the damaging effects of
chemical agents and additionally, facilitates the reduction of the hazard to personnel from any residual
chemical contamination which may be found on equipment so that protective measures may be relaxed.
The adoption of good chemical hardening practices (designs and materials) also facilitates
decontamination using decontaminants that are acceptable for use on aircraft (currently limited to hot
water and approved detergent solutions).

1.3 The use of chemical and biological weapons against airfields poses a threat to aircraft, aircrew
and maintenance crews during ground operations such as replenishment, start-up, taxy, takeoff, landing
and parking operations. Any CB hazard ingested into the aircraft, for example via the engine intakes /
APU and the environmental control system where one is fitted, presents a continuing hazard (in flight, on
the ground) to unprotected personnel. This hazard can be present for a lengthy period of time,
depending on the quantity and type of CB agent ingested, the specific aircraft design and the materials
used in the environmental control system (where one is fitted) and in the interior / crew spaces of the
aircraft. This interior hazard can persist long after the CB agents no longer represent a hazard at the
airfield.

2 AIRFRAME AND ENGINE CONTAMINATION BY CHEMICAL AGENTS

2.1 THE CHEMICAL HAZARD

2.1.1 From the viewpoint of chemical hardening, only those agents which remain liquid on a surface
long enough for solution or absorption into the substrate to take place or to allow spreading and
penetration into capillaries of various kinds represent a hazard. Agents of particular concern are
therefore the semi-persistent and persistent agents which can also damage materials, surfaces, finishes
and components.

2.1.2 CW agents may be encountered as pure, semi pure or thickened liquids depending on the
method of delivery. Pure or semi pure CW agents are generally disseminated as a fine spray whilst
thickened agents will be found in the size range 1-5 mm diameter. In each case, the liquid will generate
an associated vapour hazard, the nature of which depends on the CW agent, the surface exposed and
the meteorological conditions.

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2.1.3 Chemical agents may contact equipment and material directly if they are exposed to on-target
attack, or via secondary transfer from contaminated terrain e.g., on landing gear, by vegetation brushing
the undersides of aircraft or by transfer from contaminated personnel or loads.

2.1.4 Contamination of the aircraft and its equipment may be significantly increased if operations are
carried out on wet runways or taxiways, due to the combination of the low surface tension of the
chemical agents and the spray generated by the aircraft and its engines.

2.2 AIRFRAME

2.2.1 Chemical agents, when disseminated as liquids, may adhere to the surface of equipment or
spread over the surfaces and penetrate into capillary spaces such as screw threads, rivets joints,
flanges, aircraft panel gaps etc. The properties of these liquids are such that they are also able to
penetrate into materials such as rubber, plastics, paints, wood, foams, concrete etc. Hazards arise from
the inhalation of vapour off gassing from the free liquid and from vapour desorbing from within materials
into which the agent has dissolved or absorbed; skin contact with free or surface absorbed liquids also
presents a hazard. Although the magnitude of the residual hazard and impact on operations resulting
from aircraft exposure to persistent and semi-persistent CW agents in vapour form is much reduced
compared to the effects of liquid agent exposure, this mechanism of contamination must not be ignored.
In this case, the threat is to unprotected personnel, and not the aircraft (it will not be damaged, but it may
present an extended residual vapour hazard). Since the avoidance of contamination reduces the level of
the chemical hazard, the design of equipment should as far as possible permit its operation in such a
way as to minimise the degree of contamination to which it is exposed; the interaction of design with the
development of standard operating procedures (SOPs) should be recognised. However, reliance on
SOPs alone must be avoided. Contamination may therefore be dangerous both to the equipment user
and maintenance personnel but the nature of the hazard is different to each. In particular, the danger to
the maintenance personnel of contact with, and transfer of liquid chemical agent which may have been
trapped under coverplates and in screw threads etc., and which could be exposed when stripping
equipment down should not be overlooked, neither should the associated vapour hazard.

2.2.2 Many operational and training aircraft are equipped with either clamshell or sliding canopies
which are left partially or fully open during engine run-up and taxying in case emergency exit is required.
It should be assumed that under a CW or BW threat the canopy would be closed and where possible
covered (or the canopy seal inflated) during ground operations or between operations.

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2.2.3 Airflow rate to the crew compartments may vary considerably from taxi to take-off conditions. This
air is generally passed into the crew compartment at a high velocity to both condition avionics and to
provide crew comfort. If airborne CB contamination is present, this will be passed rapidly into the crew
compartment with the conditioning air. CB contaminants entering the crew space will contaminate all of
the surfaces with which it comes in contact, and CW agent vapour will be rapidly absorbed into surfaces
and finishes. BW agents will deposit on these surfaces. Once the contamination has entered the crew
space (and avionics compartments), this will present a residual hazard, potentially for long periods of
time, recognising that the environmental control system, where one is fitted, will also absorb and
subsequently release these contaminants into the conditioning air. Whilst adsorbed CW agents will
ultimately be dispersed to levels which permit flying personnel to operate without NBC protection, BW
contaminants may persist for much longer periods of time. Although decontamination methods for both
CW and BW hazards (compatible with aircraft interior and exterior decontamination) are expected to be
available in due course, the only methods currently available for removing the hazards is by operating
the aircraft (CW agents) or by wiping down all exposed surfaces (to remove BW agents deposited on
interior surfaces). Preventing the contaminants entering the crew spaces, using either airflow isolation or
filtration, are the only means to address the hazards at source.

2.2.4 Aircraft equipment subject to direct attack would include the following:

(a) Transparencies including canopy.

(b) Radomes.

(c) Aerial Arrays.

(d) External sensing equipment including counter-measure suites.

(e) All equipment in the undercarriage bays including the landing gear.

(f) Any other equipment exposed to atmosphere during the landing phase.

(g) Weapon hardpoints and their associated electrical interfaces.

2.2.5 Fans on air cooled equipment should be provided with particulate filters (see Section 3, Clause
3.11.22), and installed such that the equipment interior is under positive pressure. Consideration should
be given to the hermetic sealing of the equipment. Particulate filtration will only remove BW hazards and
CW in liquid (droplet) form but will not prevent the ingress of CW agent vapour. CW agent vapour
removal requires the use of an adsorbent containing filter (activated carbon)

2.3 ENGINE

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2.3.1 Use is generally made of main engine compressor bleed air as the source for cockpit / crew
compartment pressurisation, air conditioning, and other systems which use air (auxiliary power units may
also be used for this purpose). Pre-cooled bleed air further conditioned in an open-loop air cycle
environmental control system (ECS) is delivered to the cockpit / crew compartment and other systems
requiring conditioned air. In the event of an attack with Chemical or Biological Warfare (CBW) agents on
an airfield, aircraft operating in this environment will be subjected to contamination. Air from the
surroundings containing CW agent, either vapour or particulate (liquid droplets / aerosol) or BW agent
(particulate), will be drawn into the engine inlet, flow through the engine compressor to the compressor
outlet bleed air port and into the cockpit / crew compartment via the ECS, and into other using systems.

2.3.2 The quantity of chemical or biological agent that may be ingested by the engine air inlet is
dependent on engine power, aircraft speed and the concentration of the agent in the atmosphere or on
the ground. The CB hazard does not extend to a significant altitude above ground level, and so once the
aircraft has departed the airfield, the external hazard can be assumed to be negligible after the aircraft
reaches the appropriate altitude.

2.3.3 During idle or taxying, the air will be ingested from a large area forward of the engine inlet duct
but with power increasing up to maximum some of the inlet air will be in contact with the ground before
entering the engine inlet ducting. The change in aircraft attitude during take-off, however, will decrease
the quantity of air in contact with the ground until at lift-off, the effective conditions for agent
contamination ingestion are similar to that of taxying operations. Once airborne and at an appropriate
altitude, the hazard to the flying personnel is then that amount of CB agent which was ingested up to the
point the aircraft reaches this altitude.

3 AIRFRAME AND ENGINE DECONTAMINATION

3.1 AIRFRAME DECONTAMINATION

3.1.1 There are currently four decontamination methods which may be considered with special
reference to their compatibility with all relevant metallic or non-metallic materials and components used
in and on the airframe likely to be subjected to CW or BW agent contamination (there are no methods
currently available in service for decontaminating the interior spaces of aircraft). The methods are as
follows:

(a) Washing with a detergent and water. The action of this method is to dilute any liquid
chemical agent present on the surface and remove the surface contamination (liquid, or deposited
BW agent) from the aircraft. This technique does not remove CW agents that have absorbed or
dissolved into surfaces, materials or finishes. As a result, a residual off gas (vapour hazard) will
remain. BW agent decontamination by this means must not be presumed to be highly efficient or
effective. This method is ineffective against thickened CW agents. Thickened agent
decontamination is only currently achievable using weathering (see d)

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(b) Dusting with a solid powdered decontaminant, which leads to the absorption, but not
necessarily destruction of any liquid chemical agent that is present on the surface. This method will
not remove or decontaminate BW agents. In addition, this technique is labour intensive and does
not remove CW agents that have absorbed or dissolved into surfaces, materials or finishes. As a
result, a residual off gas (vapour hazard) will remain. The use of a powder on certain surfaces and
on sensitive areas of the aircraft must be avoided where there is a risk of abrasion damage or of
the powder entering sensitive equipment or assemblies.

(c) Chemical Agent Removal Using Sacrificial Coatings. Methods have been proposed
whereby decontamination is achieved by removing a sacrificial polymer film, using for example a
water-based stripping solution. Aircraft finished in an alkali removable temporary finish (ARTF) may
also be considered to be provided with a sacrificial coating. Of these two methods, only the ARTF
is in use. It should also be recognised that CW agent resistant paints are available for use on
aircraft, and these will not be damaged by liquid CW agents and their use will result in the residual
vapour hazard being substantially reduced because such paints absorb only small quantities of
CW agents. The use of such paints is encouraged, and reference to the relevant defence standard
should be made (Defence Standard 00-72 Chemical Agent Resistance Requirements for Coatings
Applied to Military Equipment)

(d) Weathering. This is the process whereby decontamination is achieved by exposing the
aircraft to the environment, including during flight. Over time, adsorbed CW agent will desorb
(evaporate), and some decontamination of surface BW contamination may also take place.
Weathering will also permit the evaporation of the liquid from a thickened CW agent challenge, but
it will not remove the thickener, which will remain adhered to the aircraft (this requires physical
removal). Weathering will not reduce any interior hazards; these are discussed above (2.2.3)

Optimum decontamination, using currently available methods, is through a combination of washing (a)
and weathering (d)

3.1.2 The general effects of some decontaminants on the airframe are given below but with the
exception of methods (a) and (d) above, the decontaminants present risks to the airframe.

(a) Chemical Agent Decontaminant (CAD) This is an aqueous solution of hydroxide and
sodium dichloroisocyanurate buffered at pH 10.3 with boric acid. The resulting alkaline chlorine
solution rapidly destroys chemical agents on materials and equipment. It is, however, harmful and
encourages the corrosion of metals, particularly light alloys. It must not be used on or near aircraft
which are to be returned to flight.

(b) STB or HTH (Super-Tropical Bleach or High-Test Hypochlorite). These chlorinated limes
contain 30% or 37% of available free chlorine by weight. The limes are made into a thin slurry with
water and are applied to surfaces with a brush. After about 30 minutes, the lime is hosed away with
water. Whilst on the surfaces, any camouflage protection is lost due to the "whitewash". The
mixture is corrosive to metals. The dry powder can cause spontaneous ignition of organic matter. It
must not be used on or near aircraft which are to be returned to flight.

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(c) Detergent and Water. A detergent/water wash is not strictly a decontamination procedure
in that the chemical and biological agents are not destroyed, they are only moved to another
location. This treatment is recommended where the use of a more aggressive decontaminant is
inadmissible. See 3.1.1 above.

(d) Fuller's Earth (FE). Fuller's Earth is mainly issued as a personal decontaminant but is also
reasonably effective for small equipments, except in the case of thickened agents. Agent is
physically absorbed and is effectively held by the powder provided that the chemical loading of the
absorbent is not exceeded. The dry material is abrasive to optical surfaces. The decontamination
powder is brushed off or mechanically removed from the airframe surface by personnel wearing
NBC equipment. This decontaminant must only be used on certain areas of an aircraft, and then
with caution. See 3.1.1 above.

(e) Sacrificial Coatings. See 3.1.1 (c) above.

3.2 ENGINE DECONTAMINATION Although CW agents are unlikely to degrade the high technology
alloys used in the construction of engine cores; it is recommended that subsequent to the ingress or
possible ingress of these agents that a liquid compressor washing procedure is carried out as a
precautionary measure. However, the high operating temperature and airflows within the core of the
engine will be such that CW agent absorption will not take place to any significant extent. Any BW
agents entering the core engine will be rapidly killed (after the first few compressor stages). The
compressor wash should remove residual contamination on the fan, on the intakes and from within the
engine bypass duct, which generally operates at a cold temperature. Most preferably, the compressor
wash should take place after flight, as flying will substantially diminish the hazard, for the reasons noted
in this paragraph. Notwithstanding the above, the engine fan area, intakes and bypass duct should not
be presumed to be thoroughly decontaminated by these means (i.e. it should be assumed that a residual
hazard will be present, requiring personnel to wear appropriate NBC protective equipment when working
on, or are near, the aircraft / engine)

4 DESIGN CONSIDERATIONS

4.1 To achieve chemical hardening the design should minimise the penetration of liquid chemical
agents into capillary spaces to facilitate the decontamination process. Capillary spaces are very difficult
to decontaminate completely. Tests have shown that a drop of agent falling on a screw head will
penetrate along the thread; after decontamination the residual vapour hazard is about 20 times greater
than from a similar drop on a plain surface. The design should also avoid other trapping features and
should employ CW agent resistant paint or other finishes. Because such finishes are non absorptive or
only absorb low amounts of CW agent, they require less attention during decontamination and produce a
much reduced residual vapour hazard. Note that deposited BW agents will remain on the surface until
they are removed, by flight or by washing. Fully effective BW decontamination by these means must not
be assumed.

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4.2 In general the external design (and the finishes applied) is of greater importance in controlling the
residual vapour hazard levels (and managing any damage risks) than the choice of materials of
construction, principally because the majority of the external surface of an aircraft consists of painted
metallic structure. However, the impact of liquid CW agent on certain materials and finishes must not be
overlooked; for example, liquid agent contamination of aircraft canopies and transparencies forming part
of the pressure fuselage can weaken the material (by a process known as environmental stress
corrosion) which could lead to explosive decompression during subsequent use of the aircraft. Features
which permit agent penetration and militate against efficient decontamination should be avoided or
sealed / coated or overpainted where possible and include the following:

(a) Joints in casings and inspection plates.

(b) Screw threads rivets, fasteners.

(c) Transparency seals.

(d) Switch, controls, meters.

(e) Connectors.

(f) Rotary and sliding seals.

(g) Hinge joints.

(h) Metal braiding on cables.

(i) Bowden and spun wire cables.

(j) Chains.

(k) Surface texture of materials.

(l) Panel gaps.

4.3 Deep surface concavities, unsealed or partially sealed quick-release fastener pockets, trap liquid
CW agents and prevent access and run-off during decontamination (by washing). The deeper and
narrower the concavity, the greater is the possibility of CW agent trapping and the more likely it is that
chemical decontamination will be ineffective. The surface design should be smooth with radiused edges
and corners especially where internal corners are unavoidable. Crinkle or textured finishes should be
avoided.

4.4 Defence Standard 08-41 illustrates some general design guidelines to minimise penetration of
liquid agents in capillary spaces and to facilitate decontamination. It also provides further advice and
information with regard to design for hardening and information on CW compliant (and non-compliant)
materials. Defence Standard 08-41 should also be reviewed. Defence Standard 00-72 defines the
requirements for CW resistant coatings and finishes, that when adopted in conjunction with good design
practices, will serve to minimise the hazards and any degradation to the aircraft.

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5 MATERIALS

5.1 DAMAGE TO MATERIALS BY CHEMICAL AGENTS - A number of liquid Chemical Warfare

agents are powerful polar organic solvents (solubility parameters in the range 8-10 Hilderbrands) and will
dissolve into / absorb into many polymeric materials. This can lead to damage effects including
blistering, softening and swelling, as well as other effects such as environmental stress corrosion. More
superficial effects can take place, such as discoloration. Absorbed / dissolved CW agent can also affect
the tensile properties of the material. Although some of these effects are reversible (by weathering),
some are not. Blistered or sticky surfaces may be difficult or impossible to decontaminate other than by
weathering, and ideally materials that behave in this way shall not be used on the exterior surfaces of
aircraft. If blistering or sticky spots are produced and are not removed, a contact hazard to service
personnel will exist. Post decontamination repairs to such areas are likely to be necessary. By design,
these risks must be avoided through the careful selection of appropriately resistant materials, or
alternatively by painting or coating such materials where their use is unavoidable. The likely agent
contamination density expected is the range 5-10 gm-2. The effect of such agents on the mechanical
properties of bulk structural plastics is likely to be insignificant except for a few polymers where
environmental stress corrosion of components under load (e.g. canopies) is possible. Degradation of
transparencies due to surface pitting and crazing is also possible, depending on the CW agent with
which contamination takes placed. Components such as acrylic or polycarbonate lenses, instrument
covers, windscreens and cockpit canopies are particularly vulnerable unless protection is provided (a
resistant coating or a cover). Some of the nerve agents will hydrolyse in air to produce dilute hydrofluoric
acid. This can lead to etching of glass and germanium surfaces. The choice of suitable blooming
coatings will give adequate protection to glass or glass faced vulnerable optical components. Coatings
do exist that can be applied to plastic based transparency materials, and these have been tested and
shown to resist damage by liquid CW agents.

5.2 DAMAGE TO MATERIALS BY DECONTAMINATION SOLUTIONS - Current UK decontaminants

may be regarded as alkaline (pH 10.5) aqueous solutions of hypochlorite. These decontaminants must
not be used on aircraft as a result of the corrosion / damage risks.

5.3 PERMEABILITY OF MATERIALS TO CHEMICAL AGENTS - Although present decontamination

techniques will permit the removal of free liquid agents from impermeable external surfaces, tactical
considerations may prevent immediate decontamination from being carried out. In the interval before
decontamination, agents will be absorbed into the body of many plastics, rubbers and porous materials.
Decontamination is not able to neutralize this absorbed agent and subsequent desorption (off gassing)
constitutes a continuing residual vapour hazard to personnel in the vicinity of the aircraft, in particular
when ventilation is restricted. It is important, therefore, when trying to impart a high level of chemical
hardening to select materials or finishes for the exposed parts of equipments which do not absorb CW
agents. The area and rate of desorption are the important parameters in determining the off gas hazard.
The absorption characteristics of some materials are outlined below: desorption characteristics are also
significant but are not as well defined in many cases.

(a) BARE METAL, GLASS, GLAZED CERAMICS - These surfaces are impermeable and can
be decontaminated readily to a level at which the residual hazard will be negligible.

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(b) FINISHES - Alkyd and acrylic paints absorb CW agents and subsequent vapour
desorption can continue for a lengthy period of time (depending on the paint, the CW agent, the
contamination levels and the prevailing environmental conditions). Some of these paints are
softened by liquid agent, and may be fully solubilised forming a liquid slurry. Catalytically hardened
aliphatic (two Part) polyurethane and epoxy paints are impermeable to CW agents and can confer
resistance to inferior permeable and porous substances. The use of such finishes is
recommended.

(c) FABRICS - Materials such as canvas, cottonwool, paper, leather etc will rapidly absorb
large quantities of CW agents. However, such materials are not generally damaged by liquid
agents. In most applications, reinforced impermeable plastics or polyurethane coated fabrics may
be substituted.

(d) WOOD - This material is absorbent unless protected by a CW agent resistant finish.
Wood is not damaged by CW agents.

(e) RUBBERS - Rubbers vary widely in their absorptive properties. Fluorinated rubbers
(viton) and bromo butyl rubber are the most agent resistant whilst silicone rubber is generally the
most absorptive and permeable. Some absorptive properties of rubbers are given in Ref 1. Few
rubbers are regarded as being chemically hard.

(f) PLASTICS - Plastics vary widely in their absorption of CW agents and individual plastics
vary in their properties from one manufacturer to another (due to variations in formulation –
molecular weight, branching, plasticizer and crystallinity). The moulding and mechanising
processes involved in the fabrication of components also has an important bearing on the surface
properties of polymers. Data from tables should therefore be treated as qualitative only and
confirmatory tests should be carried out on the particular candidate materials chosen unless
suitable and verifiable data exists. PTFE (Teflon, Fluon etc.,) is practically impermeable to CW
agent and is regarded as being chemically hard. Polyolefins (polyprolylene and polyethylene) are
relatively resistant to agent. Plasticizers tend to make materials more permeable so that plasticized
PVC is one of the most absorbent of the common plastics. The permeability of structurally
reinforced plastics is dependent not only upon the constituent resin and fibre, but also upon the
moulding process employed and the quality of surface finish achieved. Both polyesters and epoxy
resins are generally resistant to CW agent liquids with the catalytically hardened epoxies showing
the least permeability.

6 OPERATIONAL CONSIDERATIONS

6.1 Compliance with Section 3, Clause 3.11.5 requires that the aircraft and equipment be operated
by personnel wearing NBC clothing with the minimum loss of efficiency. In particular the following should
be given consideration:

(a) To prevent damage to NBC clothing, in particular the NBC gloves, sharp edges and
comers should be avoided.

(b) Deployment activities carried out by personnel wearing NBC gloves such as the operation
of controls, adjustments, maintenance functions, replenishment etc., will necessitate special
consideration being given to control spacings for accurate manipulation with a gloved hand in
which there is a loss of touch sense and dexterity.

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Table 1 below shows suggested minimum spacing for controls etc., to enable free manipulation with a
low risk of incorrect operation. The control spacings are more important than control size.

6.2 When consideration is being given to the tasks to be performed by personnel wearing the Service
respirator, due allowance should be made for the operator's limited field of vision, and to the fact that the
respirator face seal can easily be disrupted if the wearer is working in confined spaces where head
mobility is restricted. For focusing optical systems, eye relief of at least 30 mm should be allowed.

CONTROL TYPE GEOMETRY SPACING (mm)

Push Button Between button 15
Toggle Switch Between adjacent centres 20
Rotary Controls Clean annulus round periphery of control knob 25

TABLE 1

SUGGESTED CONTROL SPACINGS FOR ACCURATE MANIPULATION WITH A GLOVED HAND

REFERENCES

1. Annex C and D to CDE Guide to Chemical Hardening CDE Technical Memorandum No 79. (S/S
by Def Stan 08-11)

2. Defence Standard 08-41 Parts 1 and 2 Chemical and Biological Hardening of Military Equipment

3. Defence Standard 08-11 NBC Protection for Air Platforms.

4. Defence Standard 00-72 Chemical Agent Resistance Requirements for Coatings Applied to
Military Equipment

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LEAFLET 24

PROTECTION FROM THE EFFECTS OF NUCLEAR EXPLOSIONS, LASER WEAPONS, CHEMICAL

AND BIOLOGICAL WARFARE AGENTS

LASER WEAPON EFFECTS ON AIRCRAFT

1 INTRODUCTION

1.1 Laser weapons are those which cause dazzle or damage (see Para. 2).

1.2 This leaflet gives information on the effects of a coherent flux which may be produced by a laser
beam impinging on the aircraft; it outlines methods which may be used to evaluate the hardness of the
aircraft and guidance on how this hardness may be improved.

1.3 The characteristic effects of the laser weapon on the aircraft may include windscreen and canopy
damage, the melting or thermal degradation of metal or composite skins, damage to electrical wiring and
fluid lines, the penetration of electronic/equipment bays, the penetration of fuel tanks the ignition of
flammable fluids and the overloading or blinding of electromagnetic and optical sensors. The damaging
effects mentioned above are conditional on the value of the laser power density or energy density
impacting the target. Substantially lower power/pulse energy is required to incapacitate electro-optic or
infrared sensors, compared to the power needed to damage the structure of an aircraft. Dazzle of
sensors requires far less power or energy than damage to sensors. Countermeasure systems intended
to damage or dazzle sensors are far more likely to be encountered than structural damage laser
weapons.

2 LASER WEAPON EFFECTS ON AIRCRAFT

2.1 DAMAGE EFFECTS

2.1.1 Damage to the target can be inflicted when the radiation is transmitted either as one or more high
intensity pulses of short duration or a continuous beam. The electromagnetic energy produced by the
laser weapon is focused into an intense concentration or beam of coherent waves which is aimed at the
target air vehicles and held on the desired position until the absorbed energy causes damage to the
aircraft structure or systems. Eventual destruction of the aircraft may result.

2.1.2 Mechanisms of laser damage to a material depend greatly upon the type of laser employed and
its irradiance. Pulse lasers frequently operate at sufficiently high irradiance that a plasma is formed in
front of the target. Under plasma conditions, much of the infra-red laser beam is absorbed by the ionised
gas and ultraviolet and visible radiation is emitted by the hot plasma. The instantaneous surface
temperature during the laser pulse, whether above or below plasma threshold, is extremely high. During
the laser pulse, the chemistry of any organic coatings and adhesives is dominated by non-equilibrium
thermodynamic processes. Between pulses residual heat from the substrate leads to further degradation
of organic material by classical thermochemical mechanisms.

2.1.3 The damage mechanism associated with the absorbed energy from the laser beam can result in
the following damage processes which may affect structure and systems:

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(a) Burn-Through. Burn-through is a flux-induced damage process resulting from a interaction

between a high-energy laser beam and the target over a sufficiently long period of time. The beam
must be applied to a spot on the target until energy sufficient to melt or burn through the target is
absorbed. This requires an extremely precise pointing and tracking mechanism, possible with
appropriate feedback to identify and track the “spot” location on the target. The time required for
the beam to burn through a target material depends on the intensity of the beam and how
accurately the pointing and tracking mechanism can maintain the beam on the designated spot.

(b) Thermal Weakening. When the energy of the laser beam is insufficient to cause
burnthrough of the irradiated material, another damage process called thermal weakening can
become important. In this damage process, the material is rapidly heated by the incident radiation.
This heating causes the stiffness or elastic modulus of the material within the vicinity of the
irradiated zone to degrade and also creates a transient thermal stress field within the local
structure. If the laser beam is swept over a portion of a major load-carrying structure, such as a
wing, the combination of the material stiffness degradation, the thermal stresses, and the stresses
due to the flight loads could lead to a structural failure by buckling, by excessive plastic
deformation, or by fracture.

(c) Combustion. Heat generated by the incident flux can result in the ignition of combustible
material aboard the airborne target. Combustion of the fuel-air mixture in the ullage of fuel tanks
can occur directly, or damage to the fuel tanks due to the incident flux could allow fuel to leak into
areas where it could subsequently be ignited, creating a fire or explosion. In addition, the aircraft
structure itself may be prone to radiation-induced ignition. The use of various lightweight metals to
reduce aircraft weight can introduce a survivability problem in that many of these materials can be
ignited and will sustain burning at extremely high temperatures. Magnesium and titanium are two
such metals.

(d) Vaporisation. If the incident flux from a laser is delivered in a short duration, high-power
pulse, a dynamic loading damage process results. When energy sufficient to vaporise the surface
layer of the target is applied rapidly, the material is instantaneously converted into a gaseous state.
Inertia prevents the gaseous metal from expanding immediately, and tremendous pressures can
result. The resulting effect is similar to one that would be obtained if a thin layer of plastic explosive
was spread on the aircraft surface and detonated.

(e) Performance Degradation - Transparent Materials.

(1) The application of a sufficiently intense laser beam to a material which does not transmit
radiation at the operating wavelength of the laser results in the rapid heating of and/or damage
to that material. If the output of such a laser is directed at the cockpit transparencies of an
aircraft, the form of the resulting damage is such that obstruction of vision occurs in the effected
area irrespective of whether the transparency is of glass or polymeric construction.

(2) If the material used for the transparencies absorbs laser radiation then damage will occur
close to the surface, and is likely to take the form of flaking, pitting, crazing or ablation of the
surface. The effects do not necessarily get ‘worse’ with increased power; in some cases one
power density level may create surface crazing, but a higher level may ablate the surface and
leave the component reasonably transparent. The damage to any coatings applied to
transparencies is likely to take the form of burn spots, chipping or delamination of the coating.

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(3) The response of transparencies designed to provide ballistic protection is not primarily
penetration of laser radiation but degraded visibility. Because of the high temperature and heat
of vaporisation of glasses, almost any glass 1 cm or more in thickness will deny penetration to
10.6 um radiation if the glass can be held in the beam.

(f) Performance Degradation - Electro-Optical Systems.

(1) Laser radiation can effect electro-optical sensors in two ways, the first being dazzle which
is temporary and degrades the sensor system performance by producing a bright spot on an
image to a situation leading to complete saturation in which no image can be recognised.
Secondly, damage which is permanent can range from small pit marks on an image tube to the
complete destruction of the sensor, in which case the complete sensor system would become
unusable.

(2) The operating wavelength of the laser threat is considered to be either “in-band” or “out-of
band”. The wavelength of an in-band threat is within normal optical bandwidth of the sensor
response. Although sensors do not respond in a typical manner to out-of-band threats, damage
will still result at a given level of incident fluence or irradiance. For example, because the
outputs of Q-switched solid state lasers such as Ruby, Nd: YAG and doubled Nd: YAG are in-
band for most television cameras using low-light level charge-coupled-device (CCD) sensors,
damage and dazzle effects could be observed. However, CO2 laser radiation at 10.6 μm is out-
of-band and, in addition, because a glass camera lens does not transmit this wavelength, the
CCD sensor is not susceptible to damage. At very high irradiance levels the lens itself would be
susceptible.

(3) The susceptibility of optical components and systems to dazzle is seen as a systems
effect while damage can be both a component and systems effect and is related to the Laser
Induced Damage Threshold ( LIDT) of that component. There are several ways of defining
LIDTs which lead to different values for the same material. Thresholds can be defined in terms
of Surface Appearance, Optical Transmittance, Spatial Resolution, Scatter or Haze, and
Detector Performance. Damage thresholds of components and sensors should be measured on
an individual basis, as significant variability can be observed even between nominally similar
technologies from different manufacturers.

3 LASER HARDNESS EVALUATION AND THE PROTECTION OF AIRFRAME AND

EQUIPMENT

3.1 HARDNESS EVALUATION (HE)

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3.1.1 A large number of computer programmes or models have been generated and are available to
assess the vulnerability of aircraft to conventional threats involving penetration and single fragment
damage mechanisms. In order to damage the aircraft the laser beam must impact the aircraft albeit that
no HE charge is involved and hence the methodology for assessing the susceptibility of aircraft
constructed of metallic materials can comprise of essentially the same procedures as used in evaluating
the aircraft vulnerability to a single non-explosive penetrator. The hardness evaluation of aircraft
structures and components manufactured from Carbon Fibre Reinforced Plastics (CFRP) requires
consideration to be given to structural stiffness and residual strength loss due to the failure of any foil
hardening system and subsequent heat transfer to the substrate. Reference 4 gives data on the laser
damage mechanism and residual strength of CFRP unhardened structures. Critical structures and
system components which are considered vulnerable to laser-caused damage when illuminated by the
laser beam from specified oblique angles may be defined and the total aircraft vulnerability to the specific
laser threat determined based on the contribution of the individual critical components.

3.1.2 A susceptibility analysis of installed Electro-Optical sensor systems should be carried out using
estimated relative fluence and irradiance values at the various components within the sensor relative to
unity inputs, i.e. the radiation concentration factors or optical gains estimated for each component for the
sensor type. A database giving the “best-estimate” of LIDTs should then be used in conjunction with the
optical gains to estimate fluence and irradiance levels at the input aperture of the optical system which
are required to damage specific components. Reference 2 provides detailed descriptions of Electro-
Optical systems used by NATO air forces.

The data resulting from the analysis may be applied to Direct Vision Sights, Indirect Vision Sights
(Visible and near IR), FLIRs, and Missile Seeker Heads.

The reported LIDT of many visible and infrared materials and coatings vary significantly and this is in
part due to the different definitions of damage used by reporting agencies: some use the first signs of a
change in surface structure (the low end of the LIDT range) while others use gross melting or
catastrophic failure (the high end).

The recommended approach is to give a single value (for particular laser wavelength and pulse length)
which experience suggests will correspond to serious degradation in performance. For detectors, the aim
is to define the fluence or irradiance levels for which a change in electrical or opto-electronic
performance rather than a morphological change is observed. For optical components the aim is to
define fluence or irradiance levels which correspond to gross melting or pitting rather than the
appearance of a few microscopic blemishes.

3.1.3 With the development and availability of high power, high energy lasers in a compact form
suitable for a military use, fluences of hundreds or thousands of J/cm2 can be anticipated such that
airframe structural susceptibility will need to be evaluated. At fluence levels of this order deep
penetration of the beam inside the target is unlikely and hence the initial study should be limited only to
the interaction of laser radiation with the airframe skin and adjacent structural elements including fuel
tanks and system transmission lines.

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Reference 2 (Para’s 3.4.2 and 3.4.3) gives information on the interaction of CW Lasers and Pulsed
Lasers with aerospace materials including metals and alloys, honeycombs and metal composites,
surface finishes and paint schemes, transparent materials, together with the effects of applied stress,
airflow, and the effects of the coupling of pulse energy. Also Reference 2 Para (3.4.4 and Table 3.9)
provides a range of critical fluences in the major structural components of a generic fixed wing aircraft
based on computer simulations.

Various methodologies for assessing the Vulnerability of Aircraft to Laser Weapons including the
Analysis of High Energy Laser Irradiated Structure are listed in DRIC Subject Search No 95-001.

3.2 HARDENING AND PROTECTION OPTIONS

3.2.1 Electro-Optical Sensors and Equipments.

An extensive range of protection technologies are available including:

(a) Fixed line filters – which block specific threat wavelengths, but transmit the majority of the
spectrum appropriate to the sensor. This option is suitable only when the threat is confined to one
or a very few known wavelengths.

(b) Tunable filters – which can be externally tuned or self-activated to block the incoming
threat wavelength. Issues of response speed usually confine this approach to protection against
dazzle.

(c) Fast optical switches – which are activated automatically by incoming laser pulses by
high-intensity processes. Usually these technologies must be located at a focal plane.

(d) Spatial light modulators – which block the beam by virtue of its known angle of arrival at
the sensor. This is the only approach that may deal with very broad band (‘white light’) sources.

A comprehensive review of protection technologies is given in Ref 5.

3.2.2 Structures.

(a) Cockpit Transparencies. The laser hardening of windscreens and canopies manufactured
from plastic materials may be achieved by the usage of:

(1) Surface coatings which will reflect radiation at the CO2 wavelengths whilst being
transparent to the visible part of the spectrum.

(2) A material in which the incoming energy is absorbed within the bulk of the transparency
thus reducing the temperature rise experienced by the irradiated material.

(3) A material with high temperature stability which would be more resistant to laser damage.

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(4) Sacrificial surface ablation employed as a protective coating over the transparency or the
modification of polymers to exhibit ablation characteristics by the chemical alteration of
absorbing behaviour or the incorporation of a highly absorbing filler into the ablating polymer to
enable breakdown products associated with the energy absorbed at the surface to be removed
by the airflow.

(5) Liquid absorption technology which requires the external flow of the solution over the
canopy or windscreen during the laser attack such that the fluid absorbs the incident energy
thereby protecting the surface.

Thin metallic reflective coatings on as-cast polymethylmethacrylate transparencies are applied at

relatively low temperatures to avoid surface damage, but even with the addition of adhesion
promoting material, the in-service integrity of the design may require consideration to be given to
the applications of a suitable environmental protection outer coating. Alternatively, limited
protection may be achieved by the development of a ‘stick-on’ protection system suitable for the
application onto single or double curvature transparencies such that the reflective coating need
only be robust enough to withstand single or limited combat missions after which the ‘stick-on’
coating may be removed and replaced as required.

Generic glass composite ballistically tolerant windscreens manufactured from soda-lime glass are
not readily penetrated by irradiance, however at high radiation levels the degree of crazing,
cracking, and erosion of the material is sufficient to destroy visibility. The fabrication process of
ballistic windscreens generally involves the alternate layering of glass and adhesive interlayer
together with an inner antispall plastic back plate. Depending on the method of curing the
composite assembly, the induced internal stresses may have a marked effect on the extent and
type of laser damage and hence consideration should be given to adapting fabrication techniques
which minimises internal stress build-up by the use of low temperatures curing adhesive interlayer
material and glass with a low temperature coefficient of expansion.

(b) Airframe/Engine Structural Elements. The susceptibility of materials and structural/system

components to CW and Pulsed Laser-interaction is given in Reference 2. Table 3.9 of that
Reference provides some critical fluence levels associated with a wavelength of 10 μm, whilst Fig
3.16 indicates the susceptibility of an aircraft with special reference to specific targeted areas.

Two types of hardening concepts may be considered appropriate for possible application to aircraft
structures, equipments, and components; these techniques being the use of Reflectors or
Ablators/Insulators.

(1) The reflectivity of aluminium clad aircraft skin can be increased by electrochemical or
mechanical brightening processes. Electro-brightened high purity aluminium has a reflectance
of 97% at 10 μm but in an operational environment typical to military aircraft the electro-
brightened surface must be protected by applying an anodic film which may reduce the
reflectance to 30%. Contamination and oxidation will however decrease the reflectance
significantly and hence this countermeasure is only applicable to components in which 10 μm
reflectance can be maintained in the operational environment.

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Mechanical polishing and buffing will improve the hardness of aluminium to the laser threat and
this technique should give a reflectance of 85% under Service conditions. The reflectance of
polished aluminium can be maintained at a higher level by utilising protective schemes such as
acrylic wax or a suitable clear ablating polymer which prevents oxygen, water and various salts
attacking the substrate. Any protective overlay should not burn or char upon exposure to laser
energy as any residue may degrade surface reflectivity and also the overlay should burn off with
minimum transfer of heat to the substrate. The fitment of a second skin functioning as a
retroreflector and formed from thin brightened structural aluminium sheet secured to wing and
fuselage structure may be worthy of consideration thereby enhancing the exterior load carrying
structure of the aircraft and aircraft laser hardness. Reflective modified acrylic based paint
coatings may provide protection for surfaces such as steel which cannot be polished and
maintained without a protection coating.

(2) Ablation/Insulation materials which may be considered for structural hardening can
include the following: Teflon, Phenolic Carbon/Min-K insulation, PBI (polybenzimidazole)
composite, ESM (Silicone Elastomeric 1004), Polycrystalline Graphite, and Grafoil and Graphite
Felt.

Reference 3 (Fig 12) provides data on thickness required versus hardness, together with brief
descriptions of the above materials covering their potential use as a countermeasure.

A significant increase in the effective penetration time of the aircraft skin is likely to be achieved
by the application of ablative overcoating.

(3) Flammable Fluid Systems. Due to the large presented area and inherent vulnerability of
fuel tanks the HEL Vulnerability Analysis of an aircraft is greatly influenced by these factors. The
probabilities of fuel tank fires and explosions is conditional on such parameters as the fuel
vapour concentration in the ullage, fuel temperature, fuel tank geometry, fuel tank pressure, and
the materials used in the construction of the fuel tanks. When suitable fuel tank conditions exist
at the time of the encounter, tests have indicated that fires and explosions may be initiated by
relatively low HEL beam energy.

The prevention of aircraft fuel tank explosions may be achieved by either the installation of explosion
suppression foam; or by the introduction of an inert gas such as nitrogen which limits the oxygen
concentration in a fuel-air mixture by dilution (See Section 3, Clause 3.9 and Leaflet 19)

Laser induced fires within the fuel tank may occur if the laser beam impinges upon the liquid fuel surface
but on termination of irradiation any sustained fire will be conditional on the fuel vapour in the ullage,
ullage pressure, the temperature and type of fuel and any associated external airflow. Consideration
should be given to controlling in tank high energy laser fires by the installation of nitrogen inerting
systems or by conventional fire suppression techniques using extinguishants having a fire ‘knock-down’
performance at least equal to Halon 1301.

HEL penetration of fuel tank walls will lead to fuel leakage into adjacent dry bay areas such that with
continued irradiance may ignite the fuel and lead to a dry bay fire. Fire protection in these areas may be
achieved by the installation of either conventional active or passive fire prevention techniques (See
Section 3, Clause 3.9 and Leaflet 19)

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REFERENCES

NOTE: Documents marked * are classified and may be available only on a need-to-know basis.

1. * MOD(PE) Guidelines for the Protection of Optical and Electro-Optical Equipment against Laser
Attack, Issue 2.

2. * AGARD Advisory Report 273 Vol 2 Aerospace Applications Study 31 Defences against Directed
Energy (Laser) Weapons.

3. * Passive Laser Countermeasure Study (Applications) Vol 1, Systems Applications, Technical

Report No AFWL-TR-73-44.

4. * A Study of Laser Damage and Residual Strengths in Carbon Fibre Reinforced Plastics. RAE
Technical Report TR90017.

5. * EOPM Review Dstl/CR13625 V1.0, 2005, KJ McEwan.

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LEAFLET 25

BRAKE PARACHUTE INSTALLATIONS

SAFETY AND STRENGTH RECOMMENDATIONS

1 INTRODUCTION

1.1 A brake parachute installation must not only provide braking effort with great reliability when
needed, but it must be so designed that it cannot endanger the aeroplane in flight. The purpose of this
Leaflet is to amplify the requirements of Section 3, Clause 3.13, particularly with regard to safety and
strength.

2 SAFETY

2.1 Inadvertent streaming in flight could in certain circumstances lead to catastrophe; for instance, on
the approach or during take-off the unexpected sudden increase in drag could lead to complete loss of
control. It is therefore necessary to incorporate a device which, if inadvertent streaming takes place,
releases the parachute without applying a significant load to the aeroplane. It is recommended that the
device should be such that the parachute remains disconnected from the aeroplane until the pilot's lever
is operated.

2.2 To guard against failure of the aeroplane structure in the event of the release control being
operated at excessive speeds it is necessary to incorporate a weak link between the parachute and the
structure. It is not possible to depend on the parachute itself failing in such circumstances, as the scatter
in parachute strength is so wide. Without a weak link the average strength of the parachute would have
to be sufficiently low, so that the strongest specimen would fail before the aeroplane structure was
damaged. This would result in the weakest specimen failing at much lower speeds. With a weak link, if
its ultimate strength is equal to the actual proof strength of the aeroplane structure which carries the
parachute load, the maximum drag possible without seriously damaging the aeroplane structure will be
available for use in an emergency.

3 STRENGTH OF AEROPLANE STRUCTURE

3.1 The strength of the aeroplane structure is based on two landing cases, normal and emergency.
An ultimate factor not less than 2.0 is required in the normal case to cover the repeated applications of
the opening load. The emergency case is an arbitrary one and is based on the "re-land" case at take-off
weight but with the additional disadvantage of the flaps or other high-lift devices being out of action. An
ultimate factor not less than 1.2 only is required because the loads arise from the combination of the
emergency landing conditions and a shock load at the upper end of its scatter band.

4 STRENGTH OF PARACHUTE

4.1 The strength of the parachute presents a complex problem and the following considerations
should be taken into account:

(a) reduction in strength of material due to stitching,

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(b) reduction in strength of rigging lines after several stress cycles (high level fatigue),

(c) heat degradation of the material from work represented by hysteresis in the stress-strain
curve,

(d) surface damage (not always detectable by visual inspection) to lines, strops and canopy
from contact with the runway,

(e) unequal loading of rigging lines caused by asymmetric inflation of the canopy,

(f) reduction in strength of materials from heat in storage compartments, or when stressed in
vicinity of jet efflux, and

(g) reduction in strength of materials from the absorption of water.

4.2 Because of this general deterioration with use the parachute life as well as static strength is an
important factor in design. For this reason, the parachute strength requirements of Section 3, Clause
3.13.14 stipulate a life requirement for the normal landing case. Since it is not possible in design to
anticipate random mechanical damage, this life will be stated in terms of the minimum number of
streams required without failure under normal landing conditions, assuming no mechanical damage
occurs. This number will be such as to ensure adequate average life in service and will depend on the
aeroplane type and conditions of operation. When estimating this life it should be assured that the last
stream will be at the emergency landing speed. Those parachutes which are subjected to mechanical
damage during streaming will be rejected by the inspection made after every stream.

4.3 In the emergency case the factor of 1.4 on the maximum opening load is higher than the
corresponding factor for the aeroplane structure. This is to cover the wider scatter of strength of
parachute materials and also their deterioration with use. It is an arbitrary factor put forward until more
experience has been accumulated. It also helps to ensure that the parachute is stronger than the weak
link.

5 PARACHUTE CHARACTERISTICS

5.1 For the purpose of parachute design prior to the load measuring tests of Para 6 and for the
purpose of estimating braking performance, the parachute should be assumed to exert a steady drag
after opening of:

0.0012CD AV2 lb

where V (ft./sec.) = equivalent airspeed at the parachute taking into account any jet effects,
and

CDA (sq.ft.) = effective drag area of parachute or cluster of parachutes.

The drag area depends on the type and detail design of parachute, and if necessary, advice should be
sought from the parachute designer or the AD AS DMPS. As a rough guide, for many orthodox
parachutes CDA is approximately equal to the area of the circle of diameter equal to that of the canopy
when inflated.

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6 MEASUREMENT OF LOADS

6.1 It is difficult to estimate the values of drag area and shock factors with great accuracy, and for
design purposes it is necessary to rely on the results of load measurements made during streaming and
towing tests on a test aeroplane it which the parachute can be deployed at the required speed. Owing to
the variability in shock and steady drag loads a minimum number of tests is necessary to ensure
reasonably reliable results. In the normal landing case of Section 3, Clause 3.13.12, where the average
load is required the value taken should be based on the mean result of not less than 4 tests. In the
emergency landing case of Section 3, Clause 3.13.13, where the maximum load is required a test factor
based on the test results and on accumulated knowledge from results on similar parachutes should be
used. In the absence of more reliable information in any specific case a test factor of 1.3 should be
applied to the mean of 4 tests. If a sufficiently large number of tests can be made a test factor should be
estimated from the results obtained.

7 TESTS FOR RELIABILITY OF FUNCTIONING

(See also Section 3, Clause 3.13.16 to 3.13.19)

7.1 RIG TESTS

7.1.1 A test should be devised to check the complete sequence of operations starting with the
operation of the pilot's stream control followed by the ejection of the drogue or pilot parachute, the
simulation of the snatch load, the opening load and the steady load, and finally, on operation of the
jettison control, the opening of the jettison device.

7.1.2 The test rig should include all the operative components from the pilot's lever to the parachute
cable but need not include the parachute. The elasticity of the aeroplane local structure and of the
parachute lines, and the rates of load application should be reproduced.

7.1.3 The loads applied in the rig should be those occurring during deployment in normal landing
conditions (see Section 3, Clause 3.13.12). Jettisoning should however be demonstrated under the
steady load at landing speed, and at various angles up to that corresponding to a 30 knot crosswind, as
well as, in a few cases, under a low load corresponding to conditions at the end of the landing.

7.1.4 The sequence of operation should be repeated a sufficient number of times to demonstrate that
the required degree of reliability is attained, and that wear of the mechanism, which might interfere with
functioning during the life of the aeroplane, does not take place.

7.1.5 In cases where the environmental conditions are likely to affect the functioning, these should be
reproduced in the rig tests.

7.1.6 If the ideal of a fully representative rig test as in Para 7.1.1 is not practicable, because for
example a section of fuselage is not available, then separate tests may be made of different parts of the
installation. If this alternative is adopted, elasticity and rates of load application should still be correctly
represented (see Para 7.1.2) to check that the jettison device will not open inadvertently under the
snatch load, unless it can be shown that for the particular device used such opening is impossible.

7.1.7 As rig tests can never exactly reproduce the true operating conditions, the results of the tests
should be interpreted carefully and be correlated with the results of the flight tests discussed in Para 7.3

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7.2 OVERRIDING JETTISON DEVICE

7.2.1 The overriding jettison device should be consistent in operation at a load within ±5% of the
design value. A sufficient number of specimens should be tested to establish that the strength is within
the design limits. These tests may be made in a testing machine. The device should however also be
included in the rig test so as to establish its replacement life in service.

7.3 FLIGHT TESTS

7.3.1 The satisfactory functioning of the complete installation and the checking of the time for
deployment and jettisoning can be demonstrated during the landing trials of Section 3, Clause 3.13.17 to
3.13.19 There should be at least ten tests with no failure.

7.3.2 At least one of the tests should be made under the most adverse climatic conditions to check the
requirement of Section 3, Clause 3.13.16, that the operation is not affected by ingress of water or
freezing.

7.4 ESTIMATION OF THE SAFE LIFE OF THE PARACHUTE

7.4.1 The safe life of the parachute depends both on the use and the environmental conditions. Tests
to establish satisfactory compliance with the requirements of Section 3, Clause 3.13.12 should be
agreed between the aeroplane and the parachute manufacturers.

7.4.2 It should be assumed in estimating the safe life that the last stream may be made at the
emergency landing speed.

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LEAFLET 26

INTEGRATION OF STORES

INTEGRATION METHODOLOGY

Introduction

The purpose of this Leaflet is to outline the basic methodology for the structures and loads aspects
of the integration of a store onto a British military registered air system, for both an existing or new
store, and an existing or new air system. This Leaflet provides guidance on the approach to be
taken by the parties involved with the structures and loads aspects of integration. In this Leaflet,
the term ‘store’ refers to any launchers and / or adaptors required to carry and launch the store,
including the attachment of these to the pylon, as well as the store itself; the term ‘store’ also
includes Defensive Aids Suites (DAS) that eject expendable stores, and their associated
installation structure. The pylon / weapon carrier is assumed to be part of the air system structure,
and the qualification of the pylon with the store is the responsibility of the Designer/Design
Organisation (DO) for the pylon if different from the aircraft DO.
For the integration process to be effective, close co-operation between the aircraft Designer/DO,
the store supplier and any other party involved (e.g. launcher or pylon suppliers) is essential, with
the respective roles and responsibilities (including responsibilities for interfaces) clearly defined
early in the process. Close co-operation is also required between all the MoD Project Teams
(herein ‘PT’) involved.
It should be noted that there may be several variants of a store, each of which will have to undergo
the integration process. However, this may not be onerous if variants are similar, in terms of usage,
dimensions, mass distribution and imparted loads, as read-across can, legitimately, be used to
reduce the task.

Procedure

At the start of the integration exercise, the specific technical requirements should be defined by an
overall systems integration plan. Each system (e.g. avionics, mechanical systems) should produce
a Qualification Programme Plan (QPP) to define the requirements for demonstrating airworthiness
clearance.
The Designer/DO for the aircraft is responsible for assessing the impact of integrating the store on
the rest of the air system, in their role as aircraft integration authority. As part of this role, the
Designer/DO should provide a definition of the interface, in the Interface Control Document (ICD),
including the mechanical, electrical/software (e.g. Advanced Flight Control Systems need
knowledge of each store’s mass properties and location to determine the correct flight envelope)
and environmental interface between the air system and the store.
The store supplier is responsible for the structural clearance of their store in the specific air system
environment defined by the Designer/DO. Hence, the store supplier must produce qualification
evidence based upon the data provided within the ICD. Where the ‘store’ is a DAS installation, the
aircraft Designer/DO shall be responsible for the clearance of the installation.

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It is the responsibility of the aircraft Designer/DO to provide whoever is responsible for the
qualification of the store / launcher / adaptor (usually the suppliers of these components) with all
the necessary data to complete the qualification process. The aircraft Designer/DO should then
undertake an in-depth vetting of the processes used by the supplier and qualification evidence
provided.
The Designer/DO will then assess the qualification evidence to ensure that the full implications of
the interface to the air system have been understood and addressed and that the safety and
integrity of the air system will not be affected by the carriage / release / jettison of the store.
To ensure co-ordination of stakeholder effort, thereby managing the risks to airworthiness, the
following stage checks should be undertaken:
(a) Stage 1. Determination and agreement of the:

1. Stakeholder responsibilities and the required information exchange;

2. Route to qualification;

3. Structural criticalities and interdependencies of each item.

(b) Stage 2. Derivation, agreement and formal issue of initial design (Phase 0) Loads by
the aircraft Designer/DO and the store supplier.

(c) Stage 3. Determination, agreement and formal issue of updated (Phase 1) Loads by
both the store supplier and the aircraft Designer/DO.

(d) Stage 4. Determination, agreement and formal issue of an initial air system clearance
by both the store supplier and the aircraft Designer/DO.

(e) Stage 5. Review and formal issue of a production air system clearance by both the
store supplier and the aircraft Designer/DO.

REFERENCES

1. Tornado Stores Integration Generic Structural Qualification Programme Plan, BAe-WSS-QP-TOR-

ACA-001900, Issue 2, June 2002.

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LEAFLET 27

INTEGRATION OF STORES

DESCRIPTION OF DESIGN CONSIDERATIONS & LOADING ACTIONS

Introduction
The purpose of this leaflet is to outline the structural issues to be considered during the integration
of a store onto a British military registered air system for either an existing, or new store, and an
existing or new air system. This leaflet provides guidance on the design considerations and loading
actions to be considered, as well as a detailed description of the process by which these loading
actions can be addressed within the overall integration task.
In this leaflet, the term ‘store’ includes Defensive Aids Suites (DAS) that eject expendable stores,
and their installation structures as well as any launchers and / or adaptors required to carry and
launch the store, including the attachment of these to the pylon, as well as the store itself. The
pylon / weapon carrier is assumed to be part of the air system structure, and the qualification of the
pylon with the store is the responsibility of the Design Organisation (DO) for that pylon if different
from the aircraft DO.

KEY AREAS OF STRUCTURAL QUALIFICATION

In an assessment of the key areas, the store suppliers and the aircraft Designer/DO should
include, but not be limited to, the following areas.
Design factors, usage & ALLOWABLES:

(a) Definition of terms, e.g. reserve factors (RFs), margin of safety.

(b) Static strength factors:

(1) Proof.

(2) Ultimate.

(c) Fatigue strength factors:

(1) Representatively tested structure.

(2) Untested structure.

(3) Monitored / Unmonitored structure.

(d) In-Service Usage.

(1) Operational Loads Measurement (OLM)

(e) Material allowables (static & fatigue):

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(1) Variability.

(2) Environmental performance.

Loading actions:
(a) Inertia:

(1) Manoeuvre.

(2) Ejection / launch / jettison.

(3) Crash.

(4) Shock.

(5) Arrested landings / catapult launch / ski-jump.

(6) Landings (including VTOL on applicable Naval Platforms)

(7) Trampling of arrester cables.

(8) Repaired runway operation.

(b) Aerodynamic (total / local):

(1) Manoeuvre.

(2) Unsteady – gust / buffet.

(3) Gun / missile blast.

(c) Vibration:

(1) Normal (Natural & Forced)

(2) Gunfire.

(3) Acoustic noise.

(d) Impact damage:

(1) Birdstrike.

(2) Foreign Object Debris (FOD)

(3) Effect of ejected missile debris on air system surfaces.

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(e) Specific store effects:

(1) Trailing.

(2) Emission of corrosive substances.

(3) Heavy stores – some Helicopters may require “running” take-off and may not be
cleared for such.

(4) Size and position of store may affect ground clearance – Issue for both Fixed
Wing Aircraft and Helicopters. For instance: landing on sloping surface and carrier
landings (due to maximum compression of the oleo).

DESIGN CONSIDERATIONS
(a) Static strength.

(b) Fatigue strength.

(c) Stiffness.

(d) Material / manufacturing process.

(e) Functionality and geometric clearance considerations.

(f) Robustness.

(g) Likelihood of water ingress.

SPECIFIC CRITICAL AREAS FOR CONSIDERATION

(a) Ejector Release Unit (ERU) / Minimum Area Crutchless Ejector (MACE) and
associated back-up structures.

(b) Pylons, Stub wings and Chin mounted installations.

(c) Load input points.

(d) Major structural joints.

(e) Wings / fins.

(f) Special features (e.g. ram air turbine, flight refuelling probe) and the associated back-
up structure.

(g) System failure cases.

(1) Inadvertent operation or deployment of store control surfaces.

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(2) Hang up / hang fire.

GENERAL DATA

(a) Mass, moments of inertias and centre of gravity (c.g.) limits and tolerances.

(b) Geometries.

(c) Motor Thrust.

AEROELASTIC INSTABILITY
(a) Flutter.

(b) Helicopter Ground Resonance>

(c) Limit Cycle Oscillation (e.g. aileron buzz)

ASSESSMENT OF CRITICALITIES

Based on all available information, the structural criticality of the store, the air system and the
interface between the two should be assessed by using techniques such as FMECA, FTA or
FTEA. On the basis of this assessment, the key areas to be addressed are described in the
following paragraphs.

STATIC LOADING
To qualify the aircraft structure, the aircraft Designer/DO must make an assessment of the
loads induced by the store and compare these loads to the known strength of the air system
structure, as defined in the appropriate strength envelopes. The Designer/DO should then be
in a position to provide a statement confirming that the loads induced by new store lie within
the existing strength of the structure. If the loads are shown to lie outside the known strength
envelope for an item of structure, then a review to identify the way forward will be required.
The possible solutions could include an expansion of the existing strength envelope or the
imposition of flight limitations to ensure that the critical strength envelope is not exceeded.
Any such solution must be qualified.

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NEW DESIGNS OF STORE / AIR SYSTEM - The clearance of a store to withstand the static
loads experienced during carriage on the air system is generally carried out in two phases.
The initial step is to assess the integrity of the store airframe interface structure and airframe
back up structure under preliminary Phase 0 loads provided by the Designer/DO, with the
loading requirements defined by the Designer/DO. Store suppliers should be involved to
ensure that their requirements are provided for. If this preliminary analysis shows all RFs to
be significantly larger than 1.0, then further analysis may not be required. If the Phase 0
analysis shows a more refined study is required, then the Designer/DO will need to define
the Phase 1 loads. The Phase 1 loads will usually be presented as two components, the
steady loads derived from the wind tunnel data and the unsteady loads generally based on
measured flight data from existing stores. During the derivation of Phase 1 loads,
consideration should also be given to store failure cases and hammershock loads as dictated
by the flight controls and position of the store.
After each assessment of their structure, under both the Phase 0 and the Phase 1 loads, the
suppliers must provide a strength statement along with the supporting calculations for their
structure. All Strength Statements will be vetted by the Designer/DO before being referenced
in the overall store integration Structural Strength Statement/evidence, which is then used to
produce the required flight clearance documentation.
In the case of significant structural redundancy and prior to the issue of the Phase 0 and
Phase 1 loads, the method of generating the interface loads between each item in the store
attachment system must be agreed between all the suppliers involved. Following the
generation of the store loads by the Designer/DO, the correct interface loads can then be
derived and each company will produce their Strength Statement.
To assist in the assessment of dynamic loading situations, the Designer/DO may require
additional information as part of the Phase 1 loads, such as end of rail velocities or ejector
ram velocity profiles.
The format and method of transferring the Phase 0 and Phase 1 loads will be discussed with
each supplier so that the most appropriate format can be adopted.
If the new store is structurally and aerodynamically similar to an existing store, it may be
acceptable to obtain the required loading data by means of read-across to similar air system
/ store combinations. However, evidence of the similarity shall be provided to support such a
read-across.
EXISTING AIR SYSTEMS - Where stores are to be integrated onto an existing air system,
the qualification process should be similar to that for a new air system. The Designer/DO, as
Integration Authority, should define the static (and fatigue) loading environment in the ICD
provided to the store supplier, so that the impact of the environment on the store can be
assessed.
‘OFF THE SHELF’ STORES - Where the store has already been qualified for another
application (either from another air system or a ‘generic’ specification ), it will be necessary to
decide early in the integration process who is the store Designer/DO, responsible for the
interpretation of this qualification in terms of the specific environment defined by the aircraft
Designer/DO.

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STATIC QUALIFICATION EVIDENCE - Evidence substantiating the static clearance of the

store shall be provided in a Static Type Record / Structural Design Record which should be
cross-referenced in the aircraft Static Type Record or equivalent document.

FATIGUE LOADING
The supplier must determine the life of the store in the fatigue-loading environment defined
by the Designer/DO. The fatigue spectrum is defined by the Designer/DO after an
assessment of the aircraft’s actual and intended usage contained within the appropriate
Statement of Operating Intent and Usage (SOIU) and consultation with the user community
within the MoD PT. This spectrum is defined in the Fatigue Loading Spectrum (FLS)
document appropriate to that store. The fatigue loads should cover both quasi-static and
dynamic loading effects. The store suppliers must then produce a fatigue statement which
will be vetted by the Designer/DO prior to the issue of the overall store integration fatigue
clearance statement. If the store has a short required service fatigue life then a full fatigue
assessment may not be required with the life of the store being assured by adequate static
margins. In the case of DAS, the fatigue loads are generated from the frequency and ejection
pattern of the expendable stores, and this is information is not contained within the SOIU.
Therefore, such anticipated or actual usage information should be obtained from the PT.
The service fatigue life should be clearly stated in the Fatigue Type Record / Structural
Design Record or equivalent document so that carriage life or usage limits can be imposed.
For structure cleared by calculation alone, and structure for which the loading is not
monitored in service, appropriate additional life and stress factors should be used. These are
specified in Def Stan 00-970 Part 1, Section 3 - Structure - Clause 3.2. Exceptions to these
would be where parts of a weapon are designed to Def-Stan 07-85 or where the use of
Foreign Standards has been agreed by the MoD PTL.
If the fatigue qualification of a store is not completed by the commencement of development
flight trials then it may be possible to provide a limited interim development fatigue clearance.
This would be based on the fatigue data available at that time and an assessment of the
static RFs for the store and its attachment.
Analysis and testing of the effect of multiple store carriage on the air system should be
considered. This is required to investigate the changes in load due to different store
configurations as well as in-flight store usage i.e. the affect of changing loads following
release and the store(s) passing other stores.
RESPONSIBILITY FOR FATIGUE QUALIFICATION OF ‘OFF THE SHELF’ STORES -
Where the store has already been qualified to another spectrum (either from another air
system or a ‘generic’ spectrum), it will be necessary to decide early in the integration process
who is the store Designer/DO, responsible for the interpretation of this qualification in terms
of the specific environment defined by the aircraft Designer/DO.
FATIGUE QUALIFICATION EVIDENCE - Evidence substantiating the fatigue clearance of
the store shall be provided in a Fatigue Type Record / Structural Design Record which
should be cross-referenced in the aircraft Fatigue Type Record or equivalent document.

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MASS
Mass, c.g. and inertia properties and tolerances for all store variants will be monitored based
on documentation provided by the supplier and / or weighings as necessary. Mass and c.g.
data will be passed to the aircraft Designer/DO for use in the derivation of loads, flutter, flight
mechanics and safe separation clearances. If mass property tolerance data are not initially
available from the suppliers, the aircraft Designer/DO may specify the appropriate values to
use.

ENVIRONMENT
The environmental qualification of the store should normally occur in two stages; an initial
clearance for development flying, based on environmental levels read across from other
stores, and a final production clearance, based on environmental levels obtained from an
instrumented version of the store. It will be necessary to ensure that an instrumented store is
representative of a service-use store (e.g. in terms of mass distribution, stiffness).
The definition of the environment should include, but not be limited to, the following areas:

(a) Normal vibration:

(1) defined by the Aircraft Designer/DO using Flight Test data, when available.

(2) for Helicopters the Platform Designer/DO will define the forced vibration
frequencies and Power Spectral Densities (PSDs).

(b) Acoustic noise;

(c) Gunfire transmitted vibration;

(d) Gunfire blast pressure;

(e) Shock;

(f) Normal, lateral and longitudinal acceleration levels, the latter two being particularly
relevant for different ship-borne take-offs and landings.

The route to environmental clearance should include, but not be limited to, the following:

(a) Programme Launch.

(b) Assessment of SOIU and criticalities.

(c) Definition of Ground Loads (i.e. Ground Resonance issues)

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(d) Definition of vibration test points and times.

(e) Definition of: in-flight vibration levels, in-flight gunfire vibration levels, definition of
gunfire blast pressures, definition of in-flight acoustic noise levels.

(f) Issue of development clearance statements.

(g) Preliminary clearance.

(h) In-flight measurements and ground measurements for Ground Resonance data.

(i) Final vibration levels defined.

(j) Production qualification statements issued.

(k) Production clearance.

The Supplier should provide a response to the environmental levels provided in an

Environmental Qualification Statement, which will be required for both the Preliminary and
the Production clearances.

BIRDSTRIKE
The requirement for a birdstrike assessment of a store will be defined by the store
Designer/DO. It is expected that a risk assessment will be done and that a qualitative
statement from the supplier outlining the likely consequences of a birdstrike on the integrity
of the store and its structural attachments will be the minimum required to fulfil the birdstrike
qualification requirements. The birdstrike qualification will need to be examined for each
different application or if the carriage environment changes (e.g. due a change in tactics or
role).
LIGHTNING STRIKE
Integration of the store lightning strike protection with that of the air system shall be carried
out.
PLUME ASSESSMENT
The supplier should provide the aircraft Designer/DO with details of any rocket motor plume,
to allow an assessment of implications of that plume on the air system structure. The plume
pressure information will be supplied to the Designer/DO and the air system skin
temperatures when exposed to the plume will be assessed. The Designer/DO should then
determine the integrity of all areas affected by the plume, such as the tailplane, adjacent
stores or under fuselage panels, under the required loading. Qualification of the effects of the
plume on the launcher will be the responsibility of the supplier of that equipment, unless
otherwise agreed.

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The plume assessment will need to include an assessment of any other release debris such
as solid propellant, corrosive substances, cartridge casings or umbilical connections.
HAMMERSHOCK
The store supplier will be responsible for assessing whether hammershock loading will have
a significant effect on the store’s structure. Furthermore, the store supplier will provide data
to the aircraft Designer/DO who will be responsible for assessing whether hammershock
loading from the store will have a significant effect on the aircraft structure.
UNDER-FUSELAGE STORES
For stores that are carried under fuselage and either are or become non-jettisonable, advice
should be provided by the aircraft Designer/DO regarding the risks associated with wheels-
up landings also for helicopters landing on sloping surfaces. This advice should be based on
assessment of the strength of the store and experience gained from other stores.
RELEASE OF ADJACENT STORES
The aircraft Designer/DO will be responsible for providing a release trajectory assessment of
the store when carried in close proximity to other stores to ensure their safe release.
DYNAMIC LOADING – CARRIAGE (BUFFET) AND RELEASE
The store supplier will be responsible for providing store geometry data to the aircraft
Designer/DO in order to enable the Designer/DO to undertake an assessment of the
aerodynamic interactions for the platform environment.
FLUTTER
At the commencement of the integration exercise, the store supplier will provide the aircraft
Designer/DO (in their role as aircraft integration authority) with data relating to the mass and
inertias of the store. The aircraft Designer/DO will use this in combination with other data,
such as pylon stiffness, to make flutter predictions for the air system. These may be all that is
required if the results are acceptable and the store is dynamically similar to existing stores
used on the air system. Otherwise, a ground resonance test will be performed, the
mathematical model updated and a flutter analysis undertaken. This will be used to produce
an initial limited flutter clearance and will support the test flying of an instrumented store / air
system combination to gradually expand the flutter envelope and provide additional data for
the analysis process.
This process will be used to demonstrate sufficient margins between the maximum operating
parameters and any flutter conditions, and the Designer/DO in conjunction with the store
supplier will then produce a flutter clearance for the store.
MONITORING OF STORES CARRIAGE
It is strongly recommended that some method of monitoring stores carriage be implemented,
in order to maximise the potential life of both the store and the pylon / pylon back-up
structure, whilst ensuring that safety is not compromised.

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The objectives of the monitoring are to ensure that the qualified fatigue/vibration clearance of
the store / pylon / pylon back-up structure is not exceeded and to aid in the effective
management of available fatigue life. In addition, depending on the level of monitoring
undertaken, any exceedance of the qualified static strength of the store / pylon / pylon back-
up structure can be substantiated and recorded - hence appropriate action can be taken.
This monitoring may take a number of forms depending upon the degree of criticality, for
example:-

(a) Monitoring of carriage hours and the station upon which the store was carried.

(b) Monitoring of carriage hours in given SPCs and the station upon which the store was
carried.

(c) Monitoring of carriage hours, roll rates, Nz spectrum and the station upon which the
store was carried.

(d) Actual load measurement in the pylon / pylon back-up structure and / or
accelerometers in the store to measure accelerations / vibration.

In the first three examples above, it will be necessary to have validated the loads and the
environment at every station on which the store will be carried in every cleared combination
of adjacent stores in order to be able to compare the monitored exposure to that cleared
during the qualification of the store.

HELICOPTER ASPECTS
The external carriage of weapons on helicopters may pose particular problems that need to
be addressed.
The mass of the weapons will lower the rigid body frequencies of the helicopter when running
on the ground so that the ground resonance clearance must be revisited to ensure that
stability margins are still adequate when stores are carried.
The added mass of the weapons may also cause changes to the airframe resonance
frequencies that bring them closer to the frequencies of the forcing loads from the rotors,
resulting in increased vibration. A flight vibration survey must be done to check that vibration
is acceptable.
Some weapons carried by helicopters may have high levels of drag, particularly those
adapted from weapons originally designed for carriage by man or ground vehicle. The
increased drag will have an effect on rotor loads that will either reduce performance or
increase rotor and transmission fatigue loads if performance is maintained. Where such an
effect seems possible, a flight load survey on an instrumented aircraft must be done to
assess whether fatigue lives or Release to Service limitations need to be adjusted.

140
 DEF STAN 00-970 PART 13/11
SECTION 4

REFERENCES

1. Tornado Stores Integration Generic Structural Qualification Programme Plan, BAe-WSS-QP-TOR-

ACA-001900, Issue 2, June 2002.

141
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