U.S. Department 

of Transportation Advisory

Federal Aviation

Administration Circular

 AC 25-22 

 DATE: 3/14/00

CERTIFICATION OF

TRANSPORT AIRPLANE 

MECHANICAL SYSTEMS 

 

______________________________________________ 

Subject: CERTIFICATION OF TRANSPORT 

AIRPLANE MECHANICAL SYSTEMS

Date: 3/14/00 AC No: 25-22 

Initiated by ANM-110 Change: 

_________________________ 

1. PURPOSE. This advisory circular (AC) provides methods acceptable to the Administrator 

for showing compliance with the type certification requirements for transport airplane

mechanical systems and equipment installations. This AC is intended to provide guidance to 

airplane manufacturers, modifiers, foreign regulatory authorities, Federal Aviation 

Administration (FAA) transport airplane type certification engineers and their designees. The 

methods and procedures described herein have evolved over many years and represent current 

certification practice. Like all advisory material, these guidelines are not mandatory and do not 

constitute regulations. They are derived from previous FAA experience in finding compliance 

with the airworthiness requirements and represent methods and procedures found to be 

acceptable by that experience. Although mandatory terms such as “shall” and “must” are used in 

this AC, because the AC method of compliance is not in itself mandatory, these terms apply only 

to applicants who seek to demonstrate compliance by use of the specific method described in 

this AC. 

2. CANCELLATION. Advisory Circular 25-14, High Lift and Drag Devices, dated 

May 4, 1988, is cancelled.

3. APPLICABILITY. This advisory circular contains guidance for the latest amendment of 

the regulations and applies to all transport category airplanes approved under the provisions of 

part 25, for which a new, amended, or supplemental type certificate is requested. 
3/14/00 AC 25-22 

4. RELATED DOCUMENTS.

 a. Related Federal Aviation Regulations. Sections which prescribe requirements for the 

design, substantiation, and certification of transport airplane mechanical systems are for most 

part in Title 14 Code of Federal Regulations (14 CFR) part 25, subpart D - Design and 

Construction, and subpart F - Equipment. Additional sections (and their associated advisory 

circulars where applicable) that prescribe requirements which can have a significant impact on 

the overall design and configuration of mechanical systems are in subpart B- Performance, 

subpart C - Structure, subpart E - Powerplant, and subpart G - Operating Limitations and 

Information. Five advisory circulars are planned. Each AC will address primarily one area of 

regulations for transport category airplanes. They are: Certification of Transport Airplane 

Mechanical Systems (AC 25-XX), Certification of Electrical Equipment Installations (AC 25-

XX), Transport Airplane Propulsion Engine and Auxilliary Power Unit Installation Certification 

Handbook - The Propulsion Mega-AC (AC 25-XX), Certification of Transport Airplane 

Structure (A 25-21), and Transport Airplane Cabin Interiors Handbook (AC 25-17A). Each AC 

will cross reference the other four AC's, as necessary, for coverage of the related regulations. A 

cross reference index is provided as item 7 in appendix 4 of this AC. 

 b. Advisory Circulars (AC’s). This AC can be found and downloaded from the Internet at 

http://www.faa.gov/avr/air/airhome.htm, at the link titled "Advisory Circulars." 

Copies of advisory circulars referenced in this AC may be obtained from the US Department of 

Transportation, Subsequent Distribution Office, SVC-121.23, Ardmore East Business Center, 

3341Q 75th Avenue, Landover, MD 20785. The help line number is 301-322-4961, and the 

fascimile number is 301-386-5394 DOT Warehouse. 

5. BACKGROUND.

a In the past, advisory and guidance information applicable to transport airplane mechanical

systems and equipment installations has been formally published within AC's. However, in 

many instances, policy has been developed and applied to specific certification projects without 

formal publication. This policy appeared in the form of policy memorandums and issue papers 

(or certification review items) which were distributed to the Aircraft Certification Offices. In 

many instances this information was not organized in a way that allowed easy access. This AC 

formalizes existing policy so that the public, FAA personnel, and their designees may have 

access to this information, and contains policy extracted from existing FAA communications

used to provide guidance to applicants and other aircraft certification organizations. 

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3/14/00 AC 25-22 

b. The guidance contained in this document is presented in a format that lists the regulatory 

text, intent of the rule, background of the rule, acceptable compliance methods, and references. 

This AC is considered to be a “living document.” As such, it will be revised to maintain 

currency, such as with the issuance of part 25 rule changes, or the development of substantive 

new guidance. An index of all references listed in this AC is included as an appendix to this 

document. 

c. The methods and procedures described in this AC are only one acceptable means of 

compliance. Any alternative means proposed by the applicant should be given due 

consideration. Applicants are encouraged to use their technical ingenuity and resourcefulness in 

order to develop more efficient and less costly methods of achieving the objective of part 25. 

/s/ Donald L. Riggin

Donald L. Riggin 

Acting Manager 

Transport Airplane Directorate 

Aircraft Certification Service, ANM-100 

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 CONTENTS

Chapter 1. DESIGN AND CONSTRUCTION

Section 1. CONTROL SYSTEMS 

1. Section 25.671 General 1 

 a. Rule Text 1

2. Section 25.672 Stability augmentation and automatic and power-operated 

systems 2

 a. Rule Text 2

3. Section 25.699 Lift and drag device indicator 3 

 a. Rule Text 3

 b. Intent of Rule 3 

 c. Background 3 

 (1) Amendment 25-23 3 

 d. Policy/Compliance 3 

 (1) Advisory Circular (AC) 25-14 4 

 e. Reference 7

4. Section 25.701 Flap interconnection 7 

 a. Rule Text 7

5. Section 25.703 Takeoff warning system 8 

 a. Rule Text 8

 b. Intent of Rule 8 

 c. Background 8 

 d. Policy/Compliance 9 

 e. Reference 9

[6-10] [RESERVED] 9

Section 2. LANDING GEAR 

11. Section 25.729 Retracting mechanism 11

 a. Rule Text 11

 b. Intent of Rule 12 

 c. Background 13 

 (1) Amendment 25-23 13 

 (2) Amendment 25-42 13 

 (3) Amendment 25-72 13 

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 (4) Amendment 25-75 13 

 d. Policy/Compliance 13 

(1) Protection of equipment in wheel wells 13 

(2) Landing gear position indication system 14 

 (3) Flap system/landing gear warning system tie-in 15 

(4) Landing gear position indication system 15 

 (5) Landing gear slush tests 16 

 e. References 17

 f. Definitions 17

12. Section 25.731 Wheels 17 

 a. Rule Text 17

 b. Intent of Rule 17 

 c. Background 17 

 (1) Amendment 25-72 18 

 (2) Harmonization 18

 d. Policy/Compliance 18 

 e. References 19

13. Section 25.733 Tires 19 

 a. Rule Text 20

 b. Intent of Rule 21 

 c. Background 21 

 (1) Amendment 25-23 21 

 (2) Amendment 25-38 21 

 (3) Amendment 25-49 21 

 (4) Amendment 25-52 22 

 (5) Amendment 25-72 22 

 (6) Amendment 25-78 22 

 d. Policy/Compliance 22 

(1) Approval of retreaded tires by similarity 22 

(2) Approval: nonretreadable tires - qualification testing 24 

(3) Approval method for substitute tires 25 

(4) Replacement of bias ply tires with radial ply tires 26 

 (5) Certification of radial tire installation 28 

 (6) Critical conditions and maximum ramp weight definitions 30 

 e. References 31

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14. Section 25.735 Brakes 31 

 a. Rule Text 31

 b. Intent of Rule 33 

 c. Background 34 

 (1) Amendment 25-23 34 

 (2) Amendment 25-48 34 

 (3) Amendment 25-52 34 

 (4) Amendment 25-72 34 

 (5) Amendment 25-92 34 

 (6) Harmonization 35

 d. Policy/Compliance 35 

 (1) Carbon brake refurbishment (redensification) 35 

(2) FAA/JAA certification requirements; 

steel/carbon brakes 37 

 (3) Certification of replacement brakes 37 

 (4) Parking brake testing 40 

(5) Worn brake requirements, non original equipment parts 41 

(6) Determination of allowable worn brake limits 41 

(7) Credit for reverse thrust in worn brake testing 46 

(8) Certification of brake replacement components 47 

(9) Ground equipment - brake cooling unit 51 

(10) Certification of asbestos-free brake linings 52 

(11) Certification of brake components - performance 53 

(12) Approval of brake components utilizing new 

brake friction material 54 

 (13) Maximum quick turnaround times 55 

 e. References 57

Section 3. PERSONNEL AND CARGO ACCOMMODATIONS 

15. Section 25.773 Pilot compartment view 59 

 a. Rule Text 59

 b. Intent of Rule 60 

 c. Background 60 

 (1) Amendment 25-23 60 

 (2) Amendment 25-46 60 

 (3) Amendment 25-72 60 

 (4) Harmonization 60

 d. Policy/Compliance 60 

 e. References 60

[16-25] [RESERVED] 60 

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26. Section 25.783 Doors 61 

 a. Rule Text 61

 b. Intent of Rule 62 

 c. Background 62 

 d. Policy/Compliance 62 

 e. Reference 62

[27-29] [RESERVED] 62 

Section 4. VENTILATION AND HEATING 

30. Section 25.831 Ventilation 63 

 a. Rule Text 63

 b. Intent of Rule 64 

 c. Background 64 

 (1) Amendment 25-41 64 

 (2) Amendment 25-87 65 

 (3) Harmonization 65

 (4) Amendment 25-89 65 

 d. Policy/Compliance 65 

 (1) Ventilation rates 65 

 (2) Smoke evacuation 65

 (3) Ventilation system design 66 

(4) Airplane operation with airconditioning packs off 66 

(5) Fresh air and normal operating conditions 69 

 (6) Ventilation during takeoff 70 

 e. Reference 70

31. Section 25.832 Cabin ozone concentration 71 

 a. Rule Text 71

 b. Intent of Rule 71 

 c. Background 71 

 (1) Amendment 25-56 71 

 (2) Operating Rule 72

 d. Policy/Compliance 72 

 (1) Sea level equivalent 72 

 (2) Ground test 74 

 (3) Use of default retention ratio 74 

 (4) Use of operational limitations 75 

 (5) Meaning of "time-weighted average" 76 

(6) Example of compliance with § 25.832 77 

 e. References 80

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32. Section 25.833 Combustion heating systems 80 

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 a. Rule Text 80

 b. Intent of Rule 80 

 c. Background 80 

 d. Policy/Compliance 80 

(1) Standards for certification of combustion heaters 80 

 e. Reference 83

33. Section 25.841 Pressurized cabins 83 

 a. Rule Text 83

 b. Intent of Rule 84 

 c. Background 84 

 (1) Amendment 25-38 84 

 (2) Amendment 25-87 84 

 (3) Harmonization 85

 d. Policy/Compliance 85 

 (1) Rapid equalization of pressure 85 

(2) Access to a pressurized cargo compartment in the engine

 burst zone during high altitude operation 85 

 (3) Cabin altitude limit; § 25.841(a) 86 

 e. References 87

34. Section 25.843 Tests for pressurized cabins 88 

 a. Rule Text 88

 b. Intent of Rule 88 

 c. Background 88 

 (1) Amendment 25-87 88 

 (2) Harmonization 88

 d. Policy/Compliance 89 

(1) Pressure tests for compliance with § 25.843(a) 89 

 e. Reference 89

Section 5. FIRE PROTECTION 

35. Section 25.851 Fire extinguishers 91 

 a. Rule Text 91

 b. Intent of Rule 92 

 c. Background 92 

 d. Policy/Compliance 92 

 e. References 92

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36. Section 25.854 Lavatory fire protection 92 

 a. Rule Text 92

 b. Intent of Rule 93 

 c. Background 93 

 d. Policy/Compliance 93 

(1) Approval of lavatory smoke detectors 93 

(2) Approval of lavatory fire extinguishers containing 

 agents other than Halon 95

 e. References 95

37. Section 25.855 Cargo or baggage compartments 96 

 a. Rule Text 96

 b. Intent of Rule 96 

 c. Background 97 

 (1) Amendment 25-15 97 

 (2) Amendment 25-32 97 

 (3) Amendment 25-60 97 

 (4) Amendment 25-72 97 

 (5) Amendment 25-93 97 

 (6) Harmonization 97

 d. Policy/Compliance 97 

 (1) Halon concentration 98 

 (2) Class D to class C conversion 98 

(3) Protection of critical equipment in class E 

 cargo compartments 100 

 e. Reference 101 

38. Section 25.857 Cargo compartment classification 101 

 a. Rule Text 101 

 b. Intent of Rule 102 

 c. Background 102 

 (1) Amendment 25-32 102 

 (2) Amendment 25-60 102 

 (3) Amendment 25-93 103 

 (4) Harmonization 103 

 d. Policy/Compliance 103 

(1) Main deck cargo compartment fire protection 

certification procedures 103 

 (2) Class A cargo compartments 105 

 (3) Class E cargo compartments 106 

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(4) Access to Class E cargo compartments 107 

(5) Freighter aircraft with accessible class E compartment 107 

 e. References 108 

39. Section 25.858 Cargo or baggage compartment smoke or fire 

detection systems 108 

 a. Rule Text 108 

 b. Intent of Rule 108 

 c. Background 108 

 (1) Amendment 25-93 109 

 d. Policy/Compliance 109 

 (1) Smoke detection certification testing 109 

 (2) Certification testing for smoke detection 110 

 (3) Smoke detection certification testing 111 

(4) William J. Hughes Technical Center Research and 

 Development 114 

 e. References 114 

40. Section 25.863 Flammable fluid fire protection 114 

 a. Rule Text 114 

 b. Intent of Rule 115 

 c. Background 115 

 d. Policy/Compliance 115 

 e. Reference 115 

41. Section 25.869 Fire protection systems 115 

 a. Rule Text 115 

 b. Intent of Rule 116 

 c. Background 116 

 (1) Amendment 25-72 116 

 d. Policy/Compliance 116 

 e. References 116 

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Chapter 2. POWERPLANT

Section 1. GENERAL 

42. Section 25.943 Negative acceleration 117 

 a. Rule Text 117 

Section 2. FUEL SYSTEM COMPONENTS 

43. Section 25.1001 Fuel jettisoning system 119 

 a. Rule Text 119 

Section 3. POWERPLANT FIRE PROTECTION 

44. Section 25.1183 Flammable fluid-carrying components 121 

 a. Rule Text 121 

45. Section 25.1185 Flammable fluids 122 

 a. Rule Text 122 

46. Section 25.1189 Shutoff means 122 

 a. Rule Text 122 

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Chapter 3. EQUIPMENT

Section 1. GENERAL 

47. Section 25.1301 Function and installation 125 

 a. Rule Text 125 

 b. Intent of Rule 125 

 c. Background 125 

 (1) Harmonization 125 

 d. Policy/Compliance 125 

 (1) JAR ACJ 25.1301(b) 126 

 e. Reference 126 

48. Section 25.1309 Equipment, systems, and installations 126 

 a. Rule Text 126 

 b. Intent of Rule 127 

 c. Background 127 

 (1) Amendment 25-23 128 

 (2) Amendment 25-38 128 

 (3) Amendment 25-41 128 

 (4) Harmonization 128 

 d. Policy/Compliance 128 

 (1) Probability of a fire 128 

 (2) Examples 130 

 (3) Safe and reliable 130 

 e. References 130 

[49-51] [RESERVED] 130 

Section 2. INSTRUMENTS: INSTALLATION

52. Section 25.1325 Static pressure systems 131

 a. Rule Text 131 

 b. Intent of Rule 132 

 c. Background 132 

 (1) Amendment 25-5 132 

 (2) Amendment 25-12 132 

 (3) Amendment 25-41 132 

 d. Policy/Compliance 132 

 e. Reference 132 

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Section 3. SAFETY EQUIPMENT 

53. Section 25 1419 Ice protection 133 

 a. Rule Text 133 

 b. Intent of Rule 133 

 c. Background 133 

 (1) Amendment 25-23 134 

 (2) Amendment 25-72 134 

 d. Policy/Compliance 134 

 e. References 134 

[54-58] [RESERVED] 134 

Section 4. MISCELLANEOUS EQUIPMENT 

59. Section 25.1433 Vacuum systems 135 

 a. Rule Text 135 

 b. Intent of Rule 135 

 c. Background 135 

 (1) Amendment 25-72 135 

 d. Policy/Compliance 135 

 e. References 135 

60. Section 25.1435 Hydraulic systems 135 

 a. Rule Text 135 

 b. Intent of Rule 136 

 c. Background 137 

 (1) Amendment 25-13 137 

 (2) Amendment 25-23 138 

 (3) Amendment 25-41 138 

 (4) Amendment 25-72 138 

 (5) Exemptions 138 

 (6) Harmonization 139 

 d. Policy/Compliance 139 

 (1) Hydraulic fluid contamination 139 

(2) Certification of hydraulic lines for temporary repairs 140 

(3) Use of hydraulic lines for temporary repairs 141 

(4) Hydraulic system certification philosophy 142 

 (5) Interpretation of § 25.1435(a)(4) of 14 CFR 143 

 e. References 144 

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61. Section 25.1438 Pressurization and pneumatic systems 145 

 a. Rule Text 145 

 b. Intent of Rule 145 

 c. Background 145 

 (1) Harmonization 145 

 d. Policy/Compliance 145 

(1) Pneumatic versus pressurization system identification 145 

 e. References 146 

62. Section 25.1439 Protective breathing equipment 146 

 a. Rule Text 146 

 b. Intent of Rule 147 

 c. Background 148 

 (1) Amendment 25-38 148 

 d. Policy/Compliance 148 

(1) Protective breathing equipment usage when approved 

 under TSO-C99 and TSO-C116 148 

(2) Protective breathing equipment serviceability 149 

(3) Protective breathing equipment on the flight deck 149 

(4) Definition of terms used in § 25.1439 150 

 e. References 151 

63. Section 25.1441 Oxygen equipment and supply 151 

 a. Rule Text 151 

 b. Intent of Rule 151 

 c. Background 151 

 d. Policy/Compliance 151 

(1) Compliance with §§ 25.1309(c) and 25.1441(c) 151 

(2) Guidance, availability of flightcrew oxygen 153 

 (3) Oxygen installation for medical use 155 

 e. Reference 156 

64. Section 25.1443 Minimum mass flow of supplemental oxygen 156 

 a. Rule Text 156 

 b. Intent of Rule 157 

 c. Background 157 

 d. Policy/Compliance 157 

 e. Reference 157 

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65. Section 25.1445 Equipment standards for the oxygen distributing system 158 

 a. Rule Text 158 

 b. Intent of Rule 158 

 c. Background 158 

 d. Policy/Compliance 158 

 e. Reference 158 

66. Section 25.1447 Equipment standards for oxygen dispensing units 158 

 a. Rule Text 158 

 b. Intent of Rule 160 

 c. Background 160 

 (1) Amendment 25-41 160 

 (2) Amendment 25-87 160 

 d. Policy/Compliance 160 

(1) Oxygen equipment for forward observer's position 160 

 e. References 161 

67. Section 25.1449 Means for determining use of oxygen 161 

 a. Rule Text 161 

 b. Intent of Rule 161 

 c. Background 161 

 d. Policy/Compliance 162 

(1) Oxygen equipment for observer's position 162 

(2) Determination of passenger oxygen supply 163 

 e. References 163 

68. Section 25.1450 Chemical oxygen generators 163 

 a. Rule Text 163 

 b. Intent of Rule 164 

 c. Background 164 

 d. Policy/Compliance 164 

 e. Reference 164 

69. Section 25.1453 Protection of oxygen equipment from rupture 164 

 a. Rule Text 164 

 b. Intent of Rule 164 

 c. Background 164 

 d. Policy/Compliance 164 

 (1) Flexible plastic oxygen lines 165 

 (2) Use of plastic oxygen lines 165 

 e. Reference 166 

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70. Section 25.1455 Draining of fluids subject to freezing 166 

 a. Rule Text 166 

 b. Intent of Rule 166 

 c. Background 166 

 (1) Amendment 25-23 167 

 d. Policy/Compliance 167 

 e. Reference 167 

71. Section 25.1461 Equipment containing high energy rotors 167 

 a. Rule Text 167 

 b. Intent of Rule 167 

 c. Background 168 

 (1) Amendment 4b-8 (CAR) 168 

 (2) Amendment 4b-12 (CAM) 168 

 (3) Amendment 25-41 168 

 d. Policy/Compliance 168 

 e. Reference 168 

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Chapter 4. OPERATING LIMITATIONS AND INFORMATION

Section 1. OPERATING LIMITATIONS 

72. Section 25.1529 Instructions for continued airworthiness 169 

 H25.4 airworthiness limitations section 169 

 a. Rule Text 169 

 b. Intent of Rule 169 

 c. Background 169 

 d. Policy/Compliance 170 

 e. Reference 170 

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APPENDICES 

Appendix Subject

1. Cross reference for selected sections of part 25, 171 

Section 4b of the civil air regulations (CAR) and 

Section 4b of the civil aeronautics manual (CAM) 

2. Glossary of acronyms 172 

3. Index of advisory circular references 173 

4. Index of incorporated guidance material and references 178 

 (1) FAA memorandums and letters 178 

 (2) FAA technical standard orders (TSO) 181 

 (3) FAA orders 182 

 (4) FAA videos 182 

(5) Society of Automotive Engineers (SAE) documents 183 

 (6) Miscellaneous references 186 

(7) Cross reference to other AC handbooks 189 

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Chapter 1. DESIGN AND CONSTRUCTION

Section 1. CONTROL SYSTEMS 

1. SECTION 25.671 - GENERAL.

a. Rule Text.

(a) Each control and control system must operate with the ease, smoothness, and 

positiveness appropriate to its function. 

(b) Each element of each flight control system must be designed, or distinctively 

and permanently marked, to minimize the probability of incorrect assembly that 

could result in the malfunctioning of the system. 

(c) The airplane must be shown by analysis, tests, or both, to be capable of 

continued safe flight and landing after any of the following failures or jamming in 

the flight control system and surfaces (including trim, lift, drag, and feel systems), 

within the normal flight envelope, without requiring exceptional piloting skill or 

strength. Probable malfunctions must have only minor effects on control system 

operation and must be capable of being readily counteracted by the pilot. 

(1) Any single failure, excluding jamming (for example, disconnection or failure 

of mechanical elements, or structural failure of hydraulic components, such as 

actuators, control spool housing, and valves).

(2) Any combination of failures not shown to be extremely improbable, excluding 

jamming (for example, dual electrical or hydraulic system failures, or any single 

failure in combination with any probable hydraulic or electrical failure). 

(3) Any jam in a control position normally encountered during takeoff, climb, 

cruise, normal turns, descent, and landing unless the jam is shown to be extremely 

improbable, or can be alleviated. A runaway of a flight control to an adverse 

position and jam must be accounted for if such runaway and subsequent jamming 

is not extremely improbable. 

(d) The airplane must be designed so that it is controllable if all engines fail. 

Compliance with this requirement may be shown by analysis where that method 

has been shown to be reliable. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 

35 FR 5674, Apr. 8, 1970]

NOTE: This regulation is the subject of a Federal Aviation Regulations/Joint 

Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group may 

recommend revisions to §§ 25.671 and 25.672. The ARAC working group is 

developing Advisory Circular (AC) 25.671-1. 

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3/14/2000 AC 25-22 

2. SECTION 25.672 - STABILITY AUGMENTATION AND AUTOMATIC AND 

POWER-OPERATED SYSTEMS.

a. Rule Text.

If the functioning of stability augmentation or other automatic or power-operated 

systems is necessary to show compliance with the flight characteristics 

requirements of this part, such systems must comply with § 25.671 and the 

following: 

(a) A warning which is clearly distinguishable to the pilot under expected flight 

conditions without requiring his attention must be provided for any failure in the 

stability augmentation system or in any other automatic or power-operated 

system which could result in an unsafe condition if the pilot were not aware of the 

failure. Warning systems must not activate the control systems. 

(b) The design of the stability augmentation system or of any other automatic or 

power-operated system must permit initial counteraction of failures of the type 

specified in § 25.671(c) without requiring exceptional pilot skill or strength, by 

either the deactivation of the system, or a failed portion thereof, or by overriding 

the failure by movement of the flight controls in the normal sense. 

(c) It must be shown that after any single failure of the stability augmentation 

system or any other automatic or power-operated system- 

(1) The airplane is safely controllable when the failure or malfunction occurs at 

any speed or altitude within the approved operating limitations that is critical for 

the type of failure being considered; 

(2) The controllability and maneuverability requirements of this part are met 

within a practical operational flight envelope (for example, speed, altitude, 

normal acceleration, and airplane configurations) which is described in the 

Airplane Flight Manual; and 

(3) The trim, stability, and stall characteristics are not impaired below a level 

needed to permit continued safe flight and landing. 

[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]

NOTE: This regulation is the subject of a Federal Aviation Regulations/Joint 

Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group may 

recommend revisions to §§ 25.671 and 25.672. The ARAC working group is 

developing AC 25.671-1. 

 

NOTE: For policy and guidance on compliance with this requirement, refer to 

AC 25.672-1, Active Flight Controls. 

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3. SECTION 25.699 - LIFT AND DRAG DEVICE INDICATOR.

a. Rule Text.

(a) There must be means to indicate to the pilots the position of each lift or drag 

device having a separate control in the cockpit to adjust its position. In addition, 

an indication of unsymmetrical operation or other malfunction in the lift or drag 

device systems must be provided when such indication is necessary to enable the 

pilots to prevent or counteract an unsafe flight or ground condition, considering 

the effects on flight characteristics and performance.

(b) There must be means to indicate to the pilots the takeoff, en route, approach, 

and landing lift device positions. 

(c) If any extension of the lift and drag devices beyond the landing position is 

possible, the controls must be clearly marked to identify this range of extension. 

[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]

 b. Intent of Rule. This rule prescribes standards for providing visual indication to the pilot 

of the high lift and drag device(s) surface positions for the takeoff, enroute, approach, and 

landing conditions. 

c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR), to replace part 4b of the Civil Air Regulations (CAR). Sections 4b.323(e) 

and 4b.323(f) of the CAR basically covered trailing edge devices. These sections became

§§ 25.699(a) and 25.699(b) of 14 CFR for wing flap position indicator. 

 (1) Amendment 25-23 (April 8, 1970) expanded the rule to specifically cover all high 

drag devices or high lift devices (such as slots, spoilers, and slats), and added the requirement for 

position indication for all lift and drag devices with separate cockpit controls. The title was 

revised accordingly.

 d. Policy/Compliance Methods. Indicators that show the position of the lift or drag devices 

must be provided. If asymmetric extension of lift or drag devices could result in an unsafe 

condition, an indication of each individual control surface is required. The indication must 

clearly identify each position setting. The controls should be designed so that inadvertent 

extension beyond the landing position is not possible. For guidance on compliance with this 

requirement, refer to AC 25-14, dated 5/4/88, which is incorporated below. 

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3/14/2000 AC 25-22 

AC No.: AC 25-14 HIGH LIFT AND DRAG DEVICES 

Initiated by: ANM-110 Date: 5/4/88 

 1. PURPOSE. This advisory circular (AC) sets forth an acceptable means of 

compliance with the provisions of part 25 of the Federal Aviation Regulations (FAR) dealing 

with the certification requirements for high lift and drag devices. Guidance information is 

provided for showing compliance with structural and functional safety standards for high lift and 

drag devices and their operating systems. The intent of the requirements and some acceptable 

means of compliance are discussed. Other means are acceptable if they meet the intent of the 

regulations. 

 2. RELATED FAR SECTIONS. The contents of this AC are considered by the FAA in 

determining compliance of flaps, slats and drag devices with part 25. Related sections are 

25.301, 25.303, 25.333(e), 25.345, 25.457, 25.571, 25.671, 25.672, 25.697, 25.699, 25.701, 

25.703, and 25.1309. 

 3. BACKGROUND. For several years, special consideration has been given to high lift 

and drag devices to ensure that malfunction or failure will not result in an unsafe condition. 

These considerations are consolidated and incorporated in this AC. 

 4. STRUCTURAL REQUIREMENTS. The structure of high lift and drag devices must 

be designed to comply with the damage tolerance requirements of § 25.571, Amendment 25-45, 

of the FAR. The design should incorporate features which would provide a high probability of 

detection of any damage before the damage causes loss of the surface from the airplane. High 

lift and drag components to be evaluated under the requirements of § 25.571 typically include all 

structure which contributes significantly in reacting applied flight and actuation loads. Examples 

of such structure are the flap or slat surfaces, support linkages or tracks, hinges, fittings and 

attachments. 

 5. CONTROL SYSTEM REQUIREMENTS. The control system for high lift and drag 

devices must be designed to comply with the requirements of § 25.671. For the purpose of 

compliance with § 25.671, the control system ends where the control surface attaches to fixed 

structure such as the wing or fuselage. Examples of elements to be evaluated under the 

requirements of § 25.671 are linkages, hinges, cables, pulleys, quadrants, valves, actuator 

components, track rollers, movable tracks, bearings, and hydraulic or electrical systems. In 

accordance with § 25.671, the airplane must be shown to be capable of continued safe flight and 

landing without requiring exceptional pilot skill or strength following the failure of any single 

mechanical element or any combination of failures not shown to be extremely improbable, 

excluding jamming. Following this failure or combination of failures, the remaining structure 

4 
3/14/2000 AC 25-22 

must be able to withstand the loads defined by §§ 25.333 and 25.345. These are considered 

ultimate loads for this condition. If the surfaces are automatically or power operated, the control 

system must also be designed to meet the requirements of § 25.672. 

 6. DETAIL DESIGN CONSIDERATIONS. 

a. Unless the airplane has safe flight characteristics with the functionally related 

high lift or drag devices retracted on one side and extended on the other, the motion of the 

devices opposite sides of the plane of symmetry must be synchronized by a mechanical 

interconnection or approved equivalent means as required by § 25.701. The criteria of § 25.701 

are considered equally applicable to high lift and drag devices. 

(1) The surface interconnection must be designed for the loads resulting when the 

surfaces on one side of the plane of symmetry are jammed and immovable while the surfaces on 

the other side are free to move and the full power of the surface actuating system is applied. The 

flight loads from § 25.345 acting on the surface must be considered in combination with the 

actuating system loads (including system inertia loads). This is considered a limit load

condition. 

(2) In showing compliance with the interconnection requirements of § 25.701, all 

possible jam locations in the drive and support system should be considered. The surface 

mechanical interconnection must be able to withstand the jam condition and preclude any unsafe 

asymmetrical condition. The interconnection system is comprised of all elements which react 

the drive output from the actuator source to the jam point. These elements may include 

structures, interconnection linkages, and drive system components. When the interconnection is 

the only means to prevent an unsafe asymmetrical condition, the loads associated with the jam

conditions are considered limit loads and require a 1.5 factor of safety. A factor of safety less 

than 1.5 may be used when a reliable (i.e., a probability of failure of 10-3 or less) and 

independent means is used, in addition to the mechanical interconnection, to prevent unsafe 

asymmetry of a high lift system. The alternate system should detect the jam and shut down the 

drive system before the loads from any jam condition are reacted by the mechanical 

interconnection. The factor of safety may be as low as 1.25 if the probability of failure of the 

alternate system is 10-5 or less; however, it should not be less than 1.25 unless the alternate 

system is found to be equivalent to a mechanical interconnection. When a torque limiter is used, 

the torque tolerance limit should be used to react the required load rather than the nominal or set 

torque. A torque limiter should not be located in the drive system in a position where the limiter 

itself would allow an unsymmetrical configuration if a jam occurred. 

(3) An equivalent means of compliance with the requirements for a mechanical 

interconnection system may be substantiated using a systems safety analysis. Guidelines for 

performing a systems safety analysis are given in AC 25.1309-1, System Design Analysis. 

5
3/14/2000 AC 25-22 

b. Where failures in the drive system can result in uncommanded extensions or 

retractions of the high lift or drag devices, a positive means should be provided to limit the 

movement of the affected surfaces. This may be accomplished through irreversible drive 

actuators, no-back devices, redundancy in the drive system, or other equivalent means. 

c. In determining loads on high lift devices during actuation, it may be necessary to 

consider friction loads in the actuating system, which may be reasonably expected to occur in 

service. Flap tracks and rollers for instance, are often subjected to ice and slush which may offer 

high resistance to flap actuation. Each design should be evaluated to determine its susceptibility 

to friction in the mechanism and any loads associated with such resistance should be accounted 

for and applied in combination with normal operating loads. 

d. In evaluating the effects of failures or jamming of high lift surfaces, the effects 

of skewed surfaces on the operation of adjacent surfaces should be evaluated. Damage to 

adjacent structures and systems due to skewing of the surface should also be evaluated. 

 7. INDICATING AND WARNING SYSTEMS.

a. Indicating systems for high lift and drag devices must provide visual indication 

to the pilot of the surface positions for the takeoff, enroute, approach, and landing conditions. 

The position sensors should be located such that they show a direct indication of failure 

conditions. There should be independent monitoring of each functionally related set of surfaces 

(i.e., a set of surfaces on each side of the plane of symmetry that is driven by a common actuator, 

or is synchronized by some other means to ensure symmetric actuation) for which a failure will 

require an action or procedural change by the flightcrew. For instance, a functionally related flap 

set which is in an unsymmetrical configuration about the fuselage centerline would require an 

indication to the pilot of the unsymmetrical condition before takeoff. The indication to the 

flightcrew need not indicate the specific surface which has failed, but must clearly reflect the 

abnormal configuration (§ 25.699). The cockpit surface position display must also clearly 

distinguish a fault which was caused by a high lift "asymmetrical" deployment from a high lift 

"disagree" condition. A "disagree" condition exists when the high lift surface is stopped at a 

position different than the position commanded-by the pilot through the flap selection switch or 

handle. This distinction will aid the pilot in using proper procedures to further deploy, retract, or

leave the high lift or drag devices for continued flight. 

b. The takeoff warning system required by § 25.703 should sense the position of 

each functionally related set of high lift devices (symmetric about the airplane centerline) and 

provide aural warning during the initial portion of the takeoff roll if any set is not in an approved 

takeoff position. 

 8. FLIGHT LOADS MEASUREMENT. Notwithstanding the advancements in 

analytical methods used in predicting loads on airplane structures, accurate prediction of loads on 

wing leading edge and trailing edge high lift devices continues to be a problem. It is, therefore, 

advisable to verify the loads on these surfaces by conducting flight loads surveys regardless of 

the level of confidence in the overall loads program. 

6 
3/14/2000 AC 25-22 

 9. AIRPLANE CONTROLLABILITY. It should be shown by analysis, and where 

necessary by ground, simulation or flight tests, that the airplane has adequate stall margins and 

controllability to sustain the failure conditions addressed in paragraphs 4, 5, and 6 of this AC 

without requiring exceptional flightcrew skill or strength. It should also be demonstrated that no 

hazardous change in altitude or attitude will develop during transition to the unsymmetric 

condition considering likely transition rates. 

END OF ADVISORY CIRCULAR 25-14 

 e. Reference. None. 

4. SECTION 25.701 - FLAP INTERCONNECTION.

a. Rule Text.

(a) Unless the airplane has safe flight characteristics with the flaps or slats 

retracted on one side and extended on the other, the motion of flaps or slats on 

opposite sides of the plane of symmetry must be synchronized by a mechanical 

interconnection or approved equivalent means. 

(b) If a wing flap or slat interconnection or equivalent means is used, it must be 

designed to account for the applicable unsymmetrical loads, including those 

resulting from flight with the engines on one side of the plane of symmetry 

inoperative and the remaining engines at takeoff power. 

(c) For airplanes with flaps or slats that are not subjected to slipstream 

conditions, the structure must be designed for the loads imposed when the wing 

flaps or slats on one side are carrying the most severe load occurring in the 

prescribed symmetrical conditions and those on the other side are carrying not 

more than 80 percent of that load. 

(d) The interconnection must be designed for the loads resulting when 

interconnected flap or slat surfaces on one side of the plane of symmetry are 

jammed and immovable while the surfaces on the other side are free to move and 

the full power of the surface actuating system is applied. 

[Amdt. 25-72, 55 FR 29777, July 20, 1990]

NOTE: For policy and guidance on compliance with this requirement, refer to 

Advisory Circular (AC) 25-14, High Lift and Drag Devices, incorporated into this 

AC in § 25.699 paragraph d. 

7
3/14/2000 AC 25-22 

5. SECTION 25.703 - TAKE OFF WARNING SYSTEM.

 a. Rule Text.

A takeoff warning system must be installed and must meet the following 

requirements: 

(a) The system must provide to the pilots an aural warning that is automatically 

activated during the initial portion of the takeoff roll if the airplane is in a

configuration, including any of the following, that would not allow a safe takeoff: 

(1) The wing flaps or leading edge devices are not within the approved range of 

takeoff positions. 

(2) Wing spoilers (except lateral control spoilers meeting the requirements of 

§ 25.671), speed brakes, or longitudinal trim devices are in a position that would 

not allow a safe takeoff. 

(b) The warning required by paragraph (a) of this section must continue until- 

(1) The configuration is changed to allow a safe takeoff; 

(2) Action is taken by the pilot to terminate the takeoff roll; 

(3) The airplane is rotated for takeoff; or 

(4) The warning is manually deactivated by the pilot. 

(c) The means used to activate the system must function properly throughout the 

ranges of takeoff weights, altitudes, and temperatures for which certification is 

requested. 

[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]

 b. Intent of Rule. This rule prescribes a requirement for a takeoff warning system to warn 

the pilots during the initial portion of the takeoff roll if the airplane is in a configuration that 

would prevent successful completion of the takeoff. 

 c. Background. This section was not addressed in either part 4b of the Civil Air 

Regulations (CAR) or part 25 of Title 14, Code of Federal Regulations (14 CFR) when it was 

originally codified. This rule was introduced with Amendment 25-42, dated January 16, 1978. 

It was realized that wing flaps and associated leading edge devices posed a special problem

because some airplanes have take off flap settings that vary with weight, altitude, temperature, 

and runway length. A warning system that accounts for these variables would be extremely 

complex, and would still require the pilot to enter the proper data. In the interest of reliability, 

the rule requires the system to give a warning when the flaps or leading edge devices are not 

within the approved range of takeoff positions, e.g., when the pilot has not placed the flaps in an 

approved takeoff position or has retracted the flaps inadvertently, or if the flaps fail to move 

from the retracted position in response to a control input. Additionally, longitudinal trim devices 

(such as movable stabilizers) and drag devices (such as ground and flight spoilers) may be 

positioned inadvertently or due to failure conditions in a configuration that may prevent 

successful completion of a takeoff. The regulation applies to the stabilizer, speed brake systems, 

and rudder trim as well. 

8 
3/14/2000 AC 25-22 

d. Policy/Compliance Methods. For policy and guidance on compliance with this 

requirement, refer to Advisory Circular (AC) 25-14, High Lift and Drag Devices, incorporated 

into this AC in § 25.699 paragraph d. 

 e. Reference. The address for ordering the latest revision of the advisory circular listed 

below can be found in the Appendix to this AC. 

AC 25.703-1, Takeoff Configuration Warning Systems. 

6-10. [RESERVED] 

9
3/14/2000 AC 25-22 

THIS PAGE INTENTIONALLY LEFT BLANK. 

10 
3/14/2000 AC 25-22 

Section 2. LANDING GEAR 

11. SECTION 25.729 - RETRACTING MECHANISM.

 a. Rule Text.

(a) General. For airplanes with retractable landing gear, the following apply: 

(1) The landing gear retracting mechanism, wheel well doors, and supporting 

structure, must be designed for- 

(i) The loads occurring in the flight conditions when the gear is in the retracted 

position, 

(ii) The combination of friction loads, inertia loads, brake torque loads, air loads, 

and gyroscopic loads resulting from the wheels rotating at a peripheral speed 

equal to 1.3 VS (with the flaps in takeoff position at design takeoff weight),

occurring during retraction and extension at any airspeed up to 1.6 VS1 (with the 

flaps in the approach position at design landing weight), and 

(iii) Any load factor up to those specified in § 25.345(a) for the flaps extended 

condition. 

(2) Unless there are other means to decelerate the airplane in flight at this speed, 

the landing gear, the retracting mechanism, and the airplane structure (including 

wheel well doors) must be designed to withstand the flight loads occurring with 

the landing gear in the extended position at any speed up to 0.67 VC . 

(3) Landing gear doors, their operating mechanism, and their supporting 

structures must be designed for the yawing maneuvers prescribed for the airplane 

in addition to the conditions of airspeed and load factor prescribed in paragraphs 

(a)(1) and (2) of this section. 

(b) Landing gear lock. There must be positive means to keep the landing gear 

extended, in flight and on the ground.

(c) Emergency operation. There must be an emergency means for extending the 

landing gear in the event of- 

(1) Any reasonably probable failure in the normal retraction system; or 

(2) The failure of any single source of hydraulic, electric, or equivalent energy 

supply. 

(d) Operation test. The proper functioning of the retracting mechanism must be 

shown by operation tests. 

(e) Position indicator and warning device. If a retractable landing gear is used, 

there must be a landing gear position indicator (as well as necessary switches to 

actuate the indicator) or other means to inform the pilot that the gear is secured 

11
3/14/2000 AC 25-22 

in the extended (or retracted) position. This means must be designed as follows: 

(1) If switches are used, they must be located and coupled to the landing gear 

mechanical systems in a manner that prevents an erroneous indication of ''down 

and locked'' if the landing gear is not in a fully extended position, or of ''up and 

locked'' if the landing gear is not in the fully retracted position. The switches may 

be located where they are operated by the actual landing gear locking latch or 

device. 

(2) The flightcrew must be given an aural warning that functions continuously, or 

is periodically repeated, if a landing is attempted when the landing gear is not 

locked down.

(3) The warning must be given in sufficient time to allow the landing gear to be 

locked down or a go-around to be made. 

(4) There must not be a manual shutoff means readily available to the flightcrew 

for the warning required by paragraph (e)(2) of this section such that it could be 

operated instinctively, inadvertently, or by habitual reflexive action. 

(5) The system used to generate the aural warning must be designed to eliminate 

false or inappropriate alerts. 

(6) Failures of systems used to inhibit the landing gear aural warning, that would 

prevent the warning system from operating, must be improbable. 

(f) Protection of equipment in wheel wells. Equipment that is essential to safe 

operation of the airplane and that is located in wheel wells must be protected 

from the damaging effects of- 

(1) A bursting tire, unless it is shown that a tire cannot burst from overheat; and 

(2) A loose tire tread, unless it is shown that a loose tire tread cannot cause 

damage. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 

35 FR 5676, Apr. 8, 1970; Amdt. 25-42, 43 FR 2323, Jan. 16, 1978; Amdt. 25-72, 

55 FR 29777, Jul. 20, 1990; Amdt. 25-75, 56 FR 63762, Dec. 5, 1991]

 b. Intent of Rule. This rule provides minimum design and certification requirements for 

landing gear actuation systems to address: 

(1) Structural integrity for the nose and main landing gear, retracting mechanism(s), 

doors, gear supporting structure for actuation loads, maneuvering loads, and yawing flight 

condition loads. 

(2) Emergency means to extend gear under certain failure conditions. 

(3) Downlock and uplock design. 

(4) Gear up-and-locked and down-and-locked position indications and aural warning. 

12 
3/14/2000 AC 25-22 

(5) Protection of essential equipment located in the wheel well from a bursting tire or 

loose tire tread. 

 (6) Function demonstration. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Sections 4b.334 

and 4b.334-2 of the CAR, became § 25.729 of 14 CFR for landing gear retracting mechanism. 

 (1) Amendment 25-23 (April 8, 1970) added a wheel rotational speed based on a 

factored takeoff speed of 1.3 Vs to be used for load computations under § 25.729(a)(1)(ii) and 

changed the reference from § 25.345 to § 25.345(a) under § 25.729(a)(1)(iii).

 (2) Amendment 25-42 (January 16, 1978) clarified the rule and made minor editorial 

changes to § 25.729(e)(3). 

 (3) Amendment 25-72 (July 20, 1990) amended the rule. It made editorial changes and 

deleted reference to § 25.67(e) under § 25.729(e)(4), since § 25.67 no longer existed. 

 (4) Amendment 25-75 (December 5, 1991) revised § 25.729(e)(2) through (e)(6) to 

state objectives without stating how the requirements were to be met; thus allowing 

manufacturers to use their ingenuity in designing systems to minimize the occurrence of 

nuisance and inappropriate aural warnings. 

 d. Policy/Compliance Methods. Guidance addressing flight testing used to demonstrate 

compliance with this section may be found in Advisory Circular (AC) 25-7A, Flight Test Guide 

for Transport Category Airplanes, chapter 4, section 4, paragraph 52, issued March 31, 1998. 

Also see AC 25.963-1, Fuel Tank Access Covers, section 4, for compliance to § 25.729(f)(2) 

 (1) Protection of Equipment in wheel wells, § 25.729(f)(1). The following is extracted 

from an FAA memorandum dated December 4, 1997, which addresses whether wheel/tire 

assemblies containing thermal fuse plugs are sufficient to comply with the requirements of

§ 25.729(f)(1). 

13
3/14/2000 AC 25-22 

(a) The intent of the regulation is to protect essential equipment from the effects of 

a tire burst, regardless of the cause of the burst; overheat is just one way of causing a tire burst. 

Other ways a tire burst may be caused are: 

 1 foreign object damage. 

 2 under-inflation. 

 3 over-inflation. 

 4 overload. 

 5 an abnormal wheel component such as a melted/defective fuse plug. 

Additionally, compliance with § 25.1309(b) is required for a continued safe flight and landing. 

A tire burst must be considered regardless of its probability of occurrence. 

(b) As cited in the preamble to Amendment 25-78: “A tire burst, as referred to in 

§ 25.729(f), is a sudden, sometimes violent, venting of the pressure from within a tire, usually 

associated with a flaw in the tire, foreign object damage, or tire overheat/overload. The FAA 

assumes that tire bursts will occasionally occur, given the severe operating environment of

airplane tires, and the fact that certain tire damage may go undetected until tire failure. With this 

in mind, equipment installed in wheel wells is evaluated at the time of certification to determine 

its ability to withstand the effects of a bursting tire. Analyses and laboratory tests are performed 

to identify critical areas, and design changes are often made to ensure that a single tire burst will 

not cause loss of critical functions.” 

 (2) Landing Gear Position Indication System - "Backup Requirement" (Section 

25.729(e)). The following is extracted from an FAA memorandum dated July 12, 1988, which 

addresses whether a backup gear position system is always required. 

(a) The failure of the landing gear position indicating system does not prevent the 

continued safe flight and landing of the airplane. Further, service history has shown that gear-up 

landings on transport category airplanes are not catastrophic. Compliance with the requirements 

of § 25.729 may be accomplished by a straightforward engineering assessment of the design. A 

numerical probability analysis is not required. 

(b) If a review of the certification basis of the airplane reveals that a second 

method of landing gear position indication (e.g., a viewing port) was required due to the 

characteristics of the design, a replacement for the now-unusable viewing ports would have to be 

provided. The replacement system could be of any type (electrical, mechanical, etc.), as long as 

it provided a back-up with reliability and functionality not less than that of the system approved 

under the original type certification basis. 

(c) If the back-up indication system was installed at the manufacturer's option, and 

was not required by the type certification basis, removing the system now could result in crew 

confusion or other operational problems. Although it is not recommended to remove an optional 

backup system (to avoid confusion in the cockpit), it would be acceptable provided the airplane 

continues to meet the original type certification basis requirements. Replacing the viewing ports 

with a system that would be similar in function to the original is encouraged, i.e., a redundant 

14 
3/14/2000 AC 25-22 

means of determining that the landing gear is indeed down if there is any uncertainty from using 

the primary system. The determination of the reliability of the system need not be as rigorous as 

in (b), above, as long as the replacement system can be shown to perform its intended function. 

(d) Service difficulty reports relating to the landing gear indication system should 

be reviewed, and if the primary electrical system has required the use of the back-up viewing 

system to an inordinate degree, the back-up system would then become more important to the 

overall reliability of the indication system, and would be considered to be necessary under the 

provisions of § 21.21(b)(2). 

 (3) Flap System/Landing Gear Warning System Tie-In. The following is extracted 

from an FAA memorandum dated December 19, 1983. The memo addressed whether it was an

acceptable design to provide two approach flap settings and two landing flap positions without 

including the “landing gear-up” warning required by § 25.729(e)(4), redesignated as 

§ 25.729(e)(2), effective at Amendment level 25-72. 

(a) The FAA originally determined that providing no warning in this configuration 

was unacceptable. However, it was noted that there were transport category aircraft already 

certificated with approved alternate "selectable" means of changing the gear-up warning onset 

for alternate approach/landing flap configurations and that "uniformity" of certification 

warranted that this concept be approved if the proposed alternate means was judged acceptable 

by the cognizant Aircraft Certification Office. The concept was the use of a guarded, manually 

operated selection switch (location not specified) and the associated proposed Airplane Flight 

Manual procedures for the use of this system be provided. 

(b) FAA flight test personnel evaluated this proposed alternate design and found it 

acceptable. Based on these events, it was determined that the proposed installation is acceptable, 

meets the intent of § 25.729(e)(4), and provides an acceptable equivalent level of safety. 

 (4) Landing Gear Position Indication System. The following policy is extracted from

an FAA memorandum dated June 3, 1983, which addresses whether other regulations need to be 

considered when finding compliance with § 25.729(e) (e.g., §§ 25.1301 and 25.1309). 

(a) Section 25.729(e) provides standards for landing gear position indicating 

systems. This section is necessary, but not sufficient, for certification of these systems. This 

section only establishes the requirement for such a system; however, it is necessary to consider 

the requirements of § 25.1301 (that requires the system to be of a kind and design appropriate to 

its intended function and function properly when installed), and §§ 25.1309(a) and 25.1309(b) 

(that require failures which reduce the capability of the aircraft or crew to cope with adverse 

operating conditions to be "improbable" and those that prevent the continued safe flight and 

landing of the airplane to be "extremely improbable"). In summary, all sections of part 25 

contained in the certification basis should be reviewed for applicability. 

(b) In addition, failures that result in presenting hazardously misleading indication 

to the crew must be improbable. An example of a hazardously misleading failure condition 

15
3/14/2000 AC 25-22 

would be if the gear indicating portions of the landing gear system indicate that all gear are 

down/locked when in actuality, all gear are not fully down and locked. 

(c) If the failure evaluation process of the landing gear system (including the gear 

position indication) shows that a gear-up landing will have catastrophic consequences, then the 

occurrence of this condition should be shown to be extremely improbable. However, if the 

evaluation shows that a gear up landing would only result in a hazardous condition, then the 

occurrences of this condition would fall in the improbable category. 

(d) Reviews of transport category accident/incident reports from the FAA files 

covering the five years prior to this memorandum have not listed a catastrophic result due to 

failures in the landing gear and/or indicating system. Further, providing a second independent 

landing gear position indicating system may not improve the overall safety of' the aircraft and 

may, in some instances, lead to a more confused condition on the part of the crew in trying to 

sort out which indicating system should be used. 

 (5) Landing Gear Slush Tests. The following policy is extracted from an FAA 

memorandum dated April 22, 1983. This memo addresses the need for slush tests to ensure that 

the landing gear can be extended for an aircraft with a single hydraulic system and no auxiliary 

power source. 

(a) There is no specific certification requirement for slush tests to evaluate landing 

gear retraction/extension systems during wet, freezing conditions. Service history indicates this 

can be a problem, even on large transport category airplanes with multiple hydraulic systems. 

The FAA has not made such tests a certification requirement because there is no practical means 

of preventing accumulation of slush in the wheel well area, short of prohibiting operation during 

slushy conditions. Slush tests on a new airplane maintaining original design tolerances and 

properly lubricated joints would be of questionable value since the problem is often associated 

with poor maintenance practices. 

(b) In several reported instances of failure of the landing gear to extend, even with 

hydraulic pressure applied, the problems were usually corrected by lubrication of rotating joints. 

During slushy runway operation, water can enter rotating joints and freeze after cold soaking at 

altitude. The problem is usually reported during cold winter months and during low surface 

temperatures. 

(c) The design should be evaluated for possible accumulation of ice on the uplock 

mechanism and rotating joints, which could prevent manual operation of the emergency 

extension system. If, after careful review of a particular landing gear design, it appears that slush 

testing is necessary, such tests could be required under §§ 21.21(b)(2) and 25.729(d). If tests are 

deemed necessary, the tests should be conducted with little or no lubrication in all rotating joints. 

16 
3/14/2000 AC 25-22 

 e. References. The addresses for ordering the latest revision of advisory circulars and 

other referenced documents listed below can be found in the Appendix to this AC.

AC 20-34D, Prevention of Retractable Landing Gear Failures. 

AC 23.729-1, Landing Gear Doors and Retraction Mechanism. (For information only). 

AC 25-7A, Flight Test Guide for Transport Category Airplanes. 

AC 43.13-1A, Acceptable Methods, Techniques and Practices - Aircraft Inspection and 

Repair. 

AC 25.963-1, Fuel Tank Access Covers. 

SAE AIR-4566 - Crashworthiness Landing Gear Design. 

SAE ARP-1311A - Landing Gear- Aircraft. 

 f. Definitions. For definitions of VS, VS1, and VC, see 14 CFR part 1, section 1.2, titled 

Abbreviations and symbols. 

12. SECTION 25.731 - WHEELS.

 

 a. Rule Text.

(a) Each main and nose wheel must be approved. 

(b) The maximum static load rating of each wheel may not be less than the 

corresponding static ground reaction with- 

(1) Design maximum weight; and 

(2) Critical center of gravity. 

(c) The maximum limit load rating of each wheel must equal or exceed the 

maximum radial limit load determined under the applicable ground load 

requirements of this part. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 

55 FR 29777, Jul. 20, 1990]

 b. Intent of Rule. This rule requires use of approved wheels, either approved under an 

applicable Technical Standard Order (TSO) e.g., TSO-C26c, or approved as part of the type 

design for the airplane. Wheels must satisfy both a design static (1g) load and design limit 

landing or taxiing load determined under the applicable ground load requirements (§§ 25.471 

through 25.511). Standards for a tire installed on a wheel are contained in § 25.733. Standards 

for a brake installed on a wheel are contained in § 25.735. A TSO approval is not an approval to 

install wheels on the airplane. If the airframe manufacturer decides to install equipment 

approved under a TSO, they must conduct the applicable airplane certification tests and obtain 

FAA approval for the installation. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Sections 4b.335(a) 

and 4b.335(b) of the CAR became §§ 25.731(a) and 25.335(b) of 14 CFR for wheels. 

17
3/14/2000 AC 25-22 

 (1) Amendment 25-72 (July 20, 1990) amended the rule to become compatible with 

§ 25.25 (which had been amended). This amendment provides for weights that are in excess of 

takeoff weight, such as ramp weights, provided that the compliance with applicable structural 

requirements, including wheel strength, is demonstrated at the higher weights. 

 (2) Harmonization. This regulation is the subject of a Federal Aviation 

Regulation/Joint Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group is revising §§ 25.731 

and 25.735, and is developing a new advisory circular AC 25.735-1, and a new TSO-C135 for 

transport category airplanes, replacing the applicable parts of the existing TSO-C26c. 

 d. Policy/Compliance Methods. For guidance on compliance, the following material is 

extracted from Advisory Circulars (AC) 27-1, Certification of Normal Category Rotorcraft, and 

AC 29-2B, Certification of Transport Category Rotorcraft. It is applicable to certification of

transport category airplanes. While not specifically prepared for transport category airplanes, 

this policy material has been used for demonstrating compliance with § 25.731. For additional 

guidance, see AC 25-7A, Flight Test Guide for Transport Category Airplanes, chapter 4, section 

4, par 53, issued March 31, 1998. 

(1) The structural design loads data shall contain both a static load and a landing and 

taxiing load for each wheel. These loads are determined by virtue of compliance with the 

standards of §§ 25.731(b) and 25.731(c). The ratings of the wheel shall not be exceeded. 

TSO-C26c contains minimum performance standards for TSO approval of aircraft wheels and 

wheel-brake assemblies. Ratings are assigned in accordance with this performance standard. 

(2) If a wheel selected for an aircraft design has TSO-C26c approval, the wheel 

manufacturer will supply the rating to the aircraft manufacturer. Each wheel shall be marked as 

prescribed which includes a listing of the TSO number. Even though a wheel is TSO approved, 

the application on the aircraft (loads imposed on the wheel) requires proof that the rating is not 

exceeded. Wheel rating must not be less than airplane maximum radial load limits. 

(3) If a wheel selected for an aircraft design is not approved under TSO-C26, the 

necessary data, both detail design and assembly drawings and qualification tests and test report 

data, will be required to comply with the standards contained in part 25. Design control and 

inspections will be accomplished as a part of the aircraft type design. Structural substantiation 

and any appropriate qualification tests shall be accomplished. See §§ 25.471 through 25.497 for 

the ground load conditions. 

(4) The Tire and Rim Association Inc., issues a yearbook listing aircraft tire and rim

sizes and ratings. The dimensions and contours for aircraft wheel rims are contained in the 

yearbook. 

 e. References. The addresses for ordering the latest revision of advisory circulars, 

technical standard orders, and other referenced documents listed below can be found in the 

Appendix to this AC.

18 
3/14/2000 AC 25-22 

Section 25.731 Wheels and § 25.735 Brakes, part 25 of 14 CFR. 

AC 25-7A, Flight Test Guide for Transport Category Airplanes. 

AC 27-1, Certification of Normal Category Rotorcraft. 

AC 29-2A, Certification of Transport Category Rotorcraft. 

AC 43.13-1A, Acceptable Methods, Techniques and Practices-Aircraft Inspection and

Repair, Chap. 8, L. G. Equipment.

TSO-C26, was an acceptable standard to Civil Aeronautics Administrator, but was not 

part of the regulations. 

TSO-C26a, was a part of the regulations, under part 514 of the Regulations of the 

Administrator, as § 514.72. 

TSO-C26b, was part of the regulations as § 37.172 of 14 CFR. 

TSO-C26c, was part of the regulations as § 37.172 of 14 CFR. 

TSO-C26c with Addendum I. The TSO was removed from regulations (§ 37.172, 

Amendment 25-52, 1980) and became a voluntary standard. 

SAE AIR-811B, Disposition of Wheels which have been Overheated. 

SAE ARP-1322, Overpressurization Release Devices. 

SAE ARP-1786, Wheel Roll on Rim Criteria for Aircraft Application. 

SAE AS-707B, Thermal Sensitive Inflation Pressure Release Devices for Tubeless 

Aircraft Wheels. 

13. SECTION 25.733 - TIRES.

 a. Rule Text.

(a) When a landing gear axle is fitted with a single wheel and tire assembly, the 

wheel must be fitted with a suitable tire of proper fit with a speed rating approved 

by the Administrator that is not exceeded under critical conditions and with a 

load rating approved by the Administrator that is not exceeded under- 

(1) The loads on the main wheel tire, corresponding to the most critical 

combination of airplane weight (up to maximum weight) and center of gravity 

position, and 

(2) The loads corresponding to the ground reactions in paragraph (b) of this 

section, on the nose wheel tire, except as provided in paragraphs (b)(2) and (b)(3)

of this section. 

(b) The applicable ground reactions for nose wheel tires are as follows: 

(1) The static ground reaction for the tire corresponding to the most critical 

combination of airplane weight (up to maximum ramp weight) and center of 

gravity position with a force of 1.0g acting downward at the center of gravity. 

This load may not exceed the load rating of the tire. 

(2) The ground reaction of the tire corresponding to the most critical combination 

of airplane weight (up to maximum landing weight) and center of gravity position 

combined with forces of 1.0g downward and 0.31g forward acting at the center of 

gravity. The reactions in this case must be distributed to the nose and main 

19
3/14/2000 AC 25-22 

wheels by the principles of statics with a drag reaction equal to 0.31 times the 

vertical load at each wheel with brakes capable of producing this ground 

reaction. This nose tire load may not exceed 1.5 times the load rating of the tire. 

(3) The ground reaction of the tire corresponding to the most critical combination 

of airplane weight (up to maximum ramp weight) and center of gravity position 

combined with forces of 1.0g downward and 0.20g forward acting at the center of 

gravity. The reactions in this case must be distributed to the nose and main 

wheels by the principles of statics with a drag reaction equal to 0.20 times the 

vertical load at each wheel with brakes capable of producing this ground 

reaction. This nose tire load may not exceed 1.5 times the load rating of the tire. 

(c) When a landing gear axle is fitted with more than one wheel and tire 

assembly, such as dual or dual-tandem, each wheel must be fitted with a suitable 

tire of proper fit with a speed rating approved by the Administrator that is not 

exceeded under critical conditions, and with a load rating approved by the 

Administrator that is not exceeded by- 

(1) The loads on each main wheel tire, corresponding to the most critical 

combination of airplane weight (up to maximum weight) and center of gravity 

position, when multiplied by a factor of 1.07; and

(2) Loads specified in paragraphs (a)(2), (b)(1), (b)(2), and (b)(3) of this section 

on each nose wheel tire.

(d) Each tire installed on a retractable landing gear system must, at the maximum 

size of the tire type expected in service, have a clearance to surrounding structure 

and systems that is adequate to prevent unintended contact between the tire and 

any part of the structure or systems. 

(e) For an airplane with a maximum certificated takeoff weight of more than 

75,000 pounds, tires mounted on braked wheels must be inflated with dry nitrogen 

or other gases shown to be inert so that the gas mixture in the tire does not 

contain oxygen in excess of 5 percent by volume, unless it can be shown that the 

tire liner material will not produce a volatile gas when heated or that means are 

provided to prevent tire temperatures from reaching unsafe levels. 

[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt. 25-38, 41 FR 55467, Dec. 20, 

1976; Amdt. 25-49, 44 FR 68745, Nov. 29, 1979; Amdt. 25-72, 55 FR 29777, 

Jul 20, 1990, as amended by Amdt 25-78, Mar 29, 1993]

 b. Intent of Rule. This rule specifies type certification requirements for both design and 

performance of tires used on transport category airplanes. The tire must be of proper fit and have 

approved speed and load ratings for a particular airplane application. The maximum static 

ground reaction for the condition specified must not exceed the maximum static load rating of 

each tire. Retractable gear system tires must have adequate clearance from surrounding structure 

and systems. The tire inflation medium is to be an inert gas to avoid explosions. Tires installed 

on landing gear axles with multi-wheels (main wheel tires only), must have a 7% load margin 

included in their rating. Tire performance standards are contained in Technical Standard Order 

(TSO) TSO-C62. A TSO approval is not approval to install tires on the airplane. The airframe

manufacturer/user must conduct the applicable airplane certification tests and receive FAA 

approval for installation. 

20 
3/14/2000 AC 25-22 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Sections 4b.336(a) 

and 4b.336(b) of the CAR became §§ 25.733(a) and 25.733(b) of 14 CFR for tires. 

 (1) Amendment 25-23 (April 8, 1970) added tire speed rating as a limitation governing 

the acceptance of tires for use on particular airplanes, identified the FAA Administrator to 

approve load and speed ratings in lieu of The Tire and Rim Association, and added the term

"suitable" to "the tire of proper fit." 

 (2) Amendment 25-38 (December 20, 1976) added § 25.733(c) to require that for 

retractable landing gear, the design must account for tire production tolerances and size increases 

expected in service. The tire should have adequate clearance to prevent jamming or interference 

of landing gear mechanisms and equipment/structure. 

 (3) Amendment 25-49 (November 29, 1979) improved minimum performance 

standards applicable to landing gear main and nose wheel tires (by revising § 37.167/TSO-C62b 

to TSO-C62c) and added more comprehensive design standards covering tire loads and speed 

ratings (§ 25.733). This amendment also specified a date (December 31, 1979) after which tire 

manufacturers could no longer identify their tires as approved under earlier standards (TSO-C62, 

C62a, C62b). 

(a) This amendment evolved as follows: The minimum performance standards for 

tires were set forth in the technical standard order (§ 37.167 "Aircraft Tires,” TSO-C62b) which 

was unchanged since it became effective in 1962. A series of accidents and incidents involving

large commercial jet airplanes, particularly wide body types, in the mid 1970's involved failures 

of tires, wheels, brakes, and antiskid devices. The FAA intensified its surveillance efforts with 

respect to aircraft tires and began an analysis of the failures and potential corrective actions. 

This led to joint FAA-Industry meetings in 1976 and 1977 that resulted in a proposed set of 

revised and updated standards (Notice 79-7) to reflect the latest technology and to meet operating 

conditions. 

(b) This amendment was a response to recommendations from the National 

Transportation Safety Board (NTSB). The NTSB issued a series of recommendations and 

advisory actions affecting tires in the areas of design standards, qualification testing, quality 

control during manufacture, and operational limits. 

21
3/14/2000 AC 25-22 

 (4) Amendment 25-52 (June 2, 1980) revoked part 37, § 37.167 Aircraft Tires - TSOC62c. Subpart A of 14 CFR part 37 became Subpart O of 14 CFR part 21. Subpart B of 14 CFR 

part 37 was eliminated, and the technical performance standards were made available at FAA 

Headquarters. 

 (5) Amendment 25-72 (July 20, 1990) deleted the requirement to consider the effects of 

inertia, from § 25.733(a)(1) because such effects were determined to be negligible for constant 

speed taxi and takeoff conditions. In addition, § 25.733(c)(1) changed from a reference to 

§ 25.733(a)(1) to a direct statement of the loading conditions for the sake of clarity. 

 (6) Amendment 25-78 (March 29, 1993) added paragraph § 25.733(e) to require that 

for airplanes with a maximum certified takeoff weight of more than 75,000 pounds, the tires 

mounted on braked wheels be inflated with dry nitrogen or other inert gases so that the tire does 

not contain oxygen in excess of 5 percent by volume to prevent tire explosions. The 75,000 

pounds weight limit was based on a review of the service difficulty reports indicating that tire 

explosions, as opposed to tire bursts, tend to occur on the larger, heavier airplanes. The 5 

percent by volume limit for oxygen content was based on a series of laboratory tests indicating 

that an abrupt auto-ignition could occur for oxygen concentrations of 10 percent or more. 

(a) There had been several cases where tire explosions had occurred in transport 

category operations. A tire explosion differs from a tire burst, which can occur when an 

overheated or over inflated tire fails and releases the high pressure air contained therein. 

Protection against tire burst is required under § 25.729(f). A tire explosion is the result of a 

chemical reaction occurring when gases released from overheated tire material mix with oxygen 

in the inflation air and ignite. In 1987, the FAA issued an airworthiness directive (AD 87-08-09) 

requiring use of nitrogen for tire inflation to ensure that the tires on braked wheels of airplanes 

do not contain more than 5 percent oxygen. Amendment 25-78 was intended to accomplish the 

same purpose for new airplanes. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Approval of Retreaded Tires by Similarity. The following is extracted from an 

FAA memorandum dated March 1, 1995, that addresses whether “qualification by similarity” 

basis could be used for approving the production of retreaded aircraft tires, using a modified 

tread compound, in various sizes. 

(a) Background. A tire manufacturer had qualified several retreaded tires of 

different sizes in accordance with the stipulations of advisory circular (AC) 145-4, Inspection, 

Retread, Repair, and Alteration of Aircraft Tires (dynamometer testing). The manufacturer 

wished to produce other retreaded tires of different sizes and ratings, based on similarity. In the 

past, every tire size had to be tested for such an approval. 

22 
3/14/2000 AC 25-22 

(b) Supportive Information. The following supportive information was provided: 

 1 Nature of the Change The formulation for the tread compound used to 

retread commercial aircraft tires had been modified as follows: 

 (aa) Polymer: A proprietary description of the change. 

(bb) Filler: A proprietary description of the change. 

 2 Test Results. The change in tread compound had been proven by dynamic 

qualification testing on a sample of tire sizes. The various tire sizes were representative of the 

tire sizes listed for production approval in the applicant's letter. Qualification Test Reports 

(QTR’s) were submitted to the FAA. The QTR’s were approved by the cognizant Flight 

Standards District Office. 

 3 Physical Properties of the Tread Material. The physical properties 

measured in a laboratory showed that the new compound had a specific percent lower hysteresis 

value, an indicator of heat generation. This characteristic provided an indication of tread 

durability, and related to the qualification of tires that had been retreaded using the new 

compound. Also, this lower modulus and a lower hardness of the new compound indicated a 

better resistance to scuffing and tearing actions to which the tire is subjected in service.

 4 Static Test Results per AC 145-4. Because the tire design remained the 

same, all static measurements (overall diameter, section width, static-loaded radius, skid depth, 

and static unbalance) were not affected by this change. Tread weight remained within the retread 

process capability. 

 5 In-Service Evaluation. Tires retreaded with this new material had been 

evaluated in service with several different airlines. No problems had been reported by the 

operators and no tires had been prematurely removed from service due to tread-related 

conditions, such as chunking or tread separation. Field inspection and examination of returned 

worn casings, performed by qualified personnel, had indicated good tread integrity and wear 

characteristics. Based on cycle data gathered (number of landings per retread), the length of the 

maintenance interval remained within a reasonable percentage of the baseline tire. 

 6 Rationale Summary. The tire sizes selected to demonstrate the 

continuation of performance were representative of the tire sizes for which production approval 

was requested. At least one tire size was qualified in each of the tire design types presently in 

service. For each tire application (nose and main landing gear), at least one tire was dynamically 

qualified. With one tire size, the highest speed rating in the list of proposed tire sizes was 

successfully qualified. With another tire size, the tested sample included the highest rated 

pressure tire. With yet another tire size, the largest skid depth in the list of proposed sizes was 

23
3/14/2000 AC 25-22 

dynamically tested. Within their own class of tires, these were, historically, the most difficult 

sizes to meet the qualifications of the AC 145-4 requirements. They had often been used in the 

past to prove process or material changes. 

 7 In view of the substantiating factors detailed above, it was requested that 

all of the additional tire sizes listed in their letter be authorized for production on the basis of 

“qualification by similarity.” 

(c) Approval. The FAA carefully considered the information provided above, 

and determined that there was sufficient merit to authorize production of the additional tire sizes 

for retreading, based on ”qualification by similarity.” 

(2) Approval of Nonretreadable Tires - Qualification Testing. The following is 

extracted from an FAA memorandum dated August 18, 1988, which addresses the definition and

application of "major and minor" changes as they relate to products with a TSO versus aircraft 

installation approval. 

(a) TSO tire dynamometer testing, alone, is not considered acceptable for 

installation approval except for the most straightforward changes. This memorandum also 

addresses a potentially unsafe condition that may have resulted from retreading tires that were 

designed and approved as nonretreadable. 

 1 Based on evidence presented, including the adverse service history, it 

appears that the development of the lightweight and nonretreadable tires should have been 

classified as a "major change," which would have necessitated flight testing on the candidate 

airplane(s) prior to approval. 

(b) It was noted that the TSO was granted based on dynamometer testing, and that 

the tire was presumably shown to meet the required load-speed-time curve, and other airplane 

weight and load parameters. However, FAA policy has been that dynamometer testing is not a 

suitable substitute for airplane tests for any but the most straightforward changes. 

(c) The memorandum recommended that a suitable test program be implemented 

as soon as possible that would demonstrate that these tires are appropriate for their intended 

application. The test program that was established in January 1984, for the introduction of radial 

tires, presented later in paragraph (4)(b), is also appropriate in this case since the effect of these 

tires on airplane performance, landing gear dynamics and antiskid operation, wheel/fuse plug 

integrity, brake energy absorption, and wheel-well tire burst, is unknown. These lower weight, 

and therefore lower strength, tires may have a profound effect on the airplane performance, the 

wheel/brake integrity, or the life of the tires. The documentation that accompanies 

nonretreadable tires, once they are approved, must clearly show that the tires are not to be 

retreaded. 

(d) There is concern that there have been many changes to wheels, tires, and 

brakes that have been considered to be "product improvements," and approved under the TSO 

24 
3/14/2000 AC 25-22 

system as "minor changes," that should have been demonstrated on the airplane prior to their 

approval. 

(3) Approval Method for Substitute Tires. The following is extracted from an FAA 

memorandum dated April 14, 1988, which addresses the certification considerations that should 

be taken into account when the installation approval of a new tire on a type certificated aircraft is 

requested. 

(a) A substitute tire must be shown to be adequate for its intended use. In some 

cases, the proposed tire will be listed in the airframe manufacturer's specification documents, and 

can be installed without further investigation. The airframe manufacturer should be contacted to 

determine if data exist regarding the acceptability of the proposed tire to be substituted on the 

airplane. 

(b) If the replacement tires are of the same basic type and construction as the 

currently approved tires, but they have not been previously approved on the airplane, the 

following items should be considered. This list is not meant to imply that flight tests would be 

mandatory; rather, that the items should be discussed with the applicant and any discrepancy or 

concerns addressed. 

 1 The TSO load-speed-time curves should be compared to the airplane 

envelope to establish the suitability of the proposed tire for installation on the airplane. If the tire 

is not TSO authorized, it would be the applicant's responsibility to demonstrate that the tire 

meets all TSO requirements. 

 2 The physical dimensions of the mounted tire should allow adequate 

clearance to the wheel well and its contents, landing gear structure, gear doors, etc. 

Consideration should be given to new tires as well as retreads and possible growth effects of 

retreading. 

 3 If the weight of the proposed tire is significantly different from the 

original, or from other tires approved on the airplane, landing gear retraction times or emergency 

gear extension operation may be affected. A flight test should be conducted if there is a concern 

that any weight difference would be large enough to affect gear actuation times. 

 4 The spring rates for the proposed tire should be compared to other tires 

approved for the specific airplane at the expected load and deflection, to determine any adverse 

effect on the loads on the landing gear or any adverse dynamic coupling between the landing 

gear and tires. 

25
3/14/2000 AC 25-22 

 5 The tire rolling radius, tire pressure, tire deflection, and tire footprint 

should be compared to other tires approved on the airplane, to determine the effect on water 

spray patterns or tire heating. If the tire pressure is adjusted to obtain a certain deflection, the 

effect on tire load carrying capability should be investigated. 

 6 Uninstrumented functional landings should be conducted. If antiskid is 

installed, these functional landings should be conducted on both wet and dry runways to 

determine if any adverse effects on the system have been introduced by the proposed tire. 

 7 If comparison of the proposed tire to other tires approved on the airplane 

reveals any significant differences in the factors noted above, or if there is evidence of

significantly increased or decreased tire rolling resistance or tire μ, performance landings should 

be conducted. Increased tire rolling resistance could impact takeoff performance, decreased 

rolling resistance could affect fuse plug integrity, and a different tire μ could affect stopping 

performance. 

(4) Replacement of Bias Ply Tires with Radial Ply Tires. The following is extracted 

from an FAA memorandum dated February 7, 1984. 

(a) This memorandum provided certification criteria for installation of radial tires 

on type certificated aircraft which only had bias tire approval. The radial tire had obtained TSO 

approval and this memorandum outlined the installation considerations. 

(b) Certification Program for Replacing Bias Tires with Radial Tires, part 25 

Transports; All weight Categories, dated Jan. 12, 1984. A bias ply to radial ply tire change is not 

considered minor. The following certification program is designed for a TSO approved radial 

tire that is to replace a bias tire of equivalent rating. 

 1 The tire must be TSO-C62c approved, or equivalent as stated in 

§ 25.733(c), and rated for the speed and gross weight under consideration. The radial tire must 

have been tested to the equivalent of the TSO, i.e., loads, speeds, and energies, etc. 

 2 The vertical spring rate for the radial tire should be determined and 

compared to the original bias tire in order to assess the difference in energy absorption 

characteristics. The airplane vertical load factors should not exceed the original design values 

with the radial tire installed. Also, spin-up and spring-back loads should be evaluated if there are 

significant differences in the weight and moments of inertia of the two tire and wheel assemblies. 

(For guidance on compliance refer to §§ 25.479, 25.723, and 25.725.) 

 3 The protection of the equipment in the wheel wells should be reassessed as 

the radial tire will have different failure burst characteristics than the bias tire. This assessment 

should cover tire air blast, as well as loose tread. (For guidance on compliance refer to 

§ 25.729(f).) 

26 
3/14/2000 AC 25-22 

 4 Emergency gear extension should be reviewed and, if necessary, 

reassessed if tire size or weight changes are involved. (For guidance on compliance refer to 

§ 25.729(c).) 

 5 A maximum Rejected Takeoff (RTO) energy stop using the 3 mile taxi, 3 

full stops criteria of AC 25-7A, Flight Test Guide for Transport Category Airplanes, chapter 4, 

section 4, paragraph 55, should be conducted. 

(aa) The tire pressure is to be set to the highest value appropriate for the 

maximum takeoff weight for which approval is being sought. Pressure should be set before taxi 

and with cold tires. This test is specified to confirm (at maximum takeoff weight) acceptable tire 

structural integrity, tire anti-skid dynamics, airplane braking performance, effectiveness of the 

fuse plugs, and to assess the directional control compatibility of the airplane if the fuse plugs 

melt late in the RTO run. In addition, the fire limitations of AC 25-7A, paragraph 55c(ii), or 

latest revision, are to be observed. 

 6 A sufficient number of RTO's and landings should be conducted to 

substantiate that the aircraft performance with radial tires is as good as, or better than, originally 

demonstrated with bias tires. Tests should be conducted in accordance with AC 25-7A, 

paragraph (2)(ii)(c), or latest revision. 

 7 The fuse plug integrity no-melt tests should be conducted in accordance 

with AC 25-7A, paragraph 55c(7), unless the applicant demonstrates the energies absorbed by 

the brake with a radial tire are no greater than the energies absorbed in the same brake when 

using the bias tire when comparing equivalent flywheel energies on the dynamometer. The fuse 

plug integrity test should also be run if the brake or tire temperatures of the radial 

tire/wheel/brake assembly are greater than the brake or tire temperatures of the bias 

tire/wheel/brake assembly when comparing equivalent flywheel energies on the dynamometer. 

 8 A sufficient number of flight tests should be conducted on wet and dry 

runways to establish that the anti-skid and/or autobrakes will function compatibly and acceptably 

with the radial tires. 

 9 A high speed takeoff at the highest speed for which approval is being 

sought, and sufficient taxi and turn tests to assess dynamic stability should be conducted. 

 10 If the airplane is approved for unimproved runways, additional evaluations 

should be conducted if there is an appreciable change in the tire footprint. 

 11 Water spray characteristics should be reviewed and, if necessary, 

reassessed. 

 12 Tire intermix, radial/bias, should be considered based on an evaluation on 

a case by case basis. For example, some of these items (1-11) may not apply for a nose tire 

27
3/14/2000 AC 25-22 

intermix. Any additional considerations deemed necessary by the cognizant ACO should be 

addressed. 

 13 If there is a significant tire tread change from one radial tire to another 

radial tire, the airplane performance should be considered and if necessary reassessed.

 14 Additional testing may be required if certification for an increase in 

airplane weight or kinetic energy is requested. 

(5) Certification of Radial Tire Installation. The following is extracted from an FAA 

memorandum dated March 28, 1984, which addresses as to why is it necessary to consider items

(4)(b)1. through (4)(b)14. addressed in the preceding paragraph, and provides historical 

information on the reasons for certain items contained in the radial tire certification program. It 

also explains some of the reasons why dynamometer testing alone is not considered adequate for 

installation approval of replacement tires. 

(a) The "Certification Program for Replacing Bias Tires with Radial Tires" dated 

January 12, 1984, is still the FAA recommended program. A bias ply tire to radial ply tire 

change is not considered minor. Dynamometer tests are not an acceptable substitute for airplane 

tests for tires, wheels, and brakes. This position, which is supported by FAA certification 

experience, is taken in AC 25-7A, chapter 4, section 4, paragraph 55. 

 1 For example, the FAA was requested to certificate a medium transport 

with three sets of tire/wheel/brake combinations for an increased gross weight. The 

manufacturer wished to certificate these combinations based on extrapolated data from existing 

lower energy data of one set, and to the higher TSO authorized energy levels of all three sets; 

however, the FAA insisted a full testing program be done. All three sets failed to achieve the 

projected braking coefficient of friction (μ). In addition, the set that was supposed to have the 

greatest margin, the set designated for future growth potential, almost failed due to shearing of 

the brake rotor lugs. The wheels, tires and fuseplugs functioned properly. 

(b) Another example was the certification of a new model large transport. The 

tires had received TSO authorization, but were failing at the high speed takeoffs within the TSO 

performance envelope. Going further back, the FAA has had similar experiences relative to 

fuseplugs, anti-skid brake lock-up, and so forth. These experiences confirm that the 

dynamometer is not a satisfactory tool in predicting performance, compatibility, fuseplug 

melt/no-melt, etc., for tires, wheels, and brakes. The dynamometer is even questionable when it 

is used for trend data, as demonstrated by the medium transport example above. It is, however, 

useful as background information. 

(c) That the dynamometer does not accurately simulate actual conditions may be 

due to its inherent inability to adequately simulate the following: 

 1 Actual runway conditions (i.e., round versus the flat surface of the 

runway, surface conditions and so forth). 

28 
3/14/2000 AC 25-22 

 2 Thermodynamic interaction (i.e., conduction. radiation, and convection 

between adjacent tires, wheels, and brakes, between the tire and the runway surface, and the 

effects of the air). 

 3 Airplane/System dynamics (i.e., the interaction between the airplane and 

supporting gear and systems, control systems and so forth). 

(d) Adequate means of simulation could probably be developed with new test 

criteria (if comparisons are to be made) and new techniques. The end result would most likely 

be a test fixture on a track with a complete landing system including supporting structure, 

systems duplicating the number of landing gear, and so on, which would be capable of intricate 

simulations of takeoff, landing and braking dynamics. However, it would probably be less 

expensive to test the airplane as is currently the practice. 

(e) Another problem is establishing the adequacy of the tire relative to the TSO 

data and to additional tests. The airframe manufacturer may request that the tire, wheel, and 

brake manufacturers conduct specific tests beyond the scope of the TSO. Side loading and 

overloading are good examples of additional tests. We can never be sure after completing the 

airplane tests that the tire (wheel or brake also) reliability is dependent totally on the TSO tests or

on the TSO test and the additional tests specified. We are aware of at least one case where two 

tires manufactured by two different companies were installed on the same airplane and operated 

within the same TSO envelopes, yet one tire type failed while the other passed. The successful 

tire type had additional side loads specified which were not part of the TSO testing. As a 

repetitive statement, the FAA position is that TSO data is not a satisfactory substitute for tire 

wheel or brake tests on the airplane. 

(f) If the radial ply tire is heavier than the bias tire, it must be determined whether 

that is significant. Lighter tire weight can affect some emergency extension systems as some of 

these systems use the landing gear weight to overcome the unlocking residual forces to provide 

enough inertia to extend the gear down, over center, and to lock the gear in place. Heavier 

weight may have an effect on retraction times. Radial tires may weigh as much as 1/4 less than 

the bias tires they would replace. 

 (g) Rejected takeoff tests are not only to prove the capability of the brakes, but 

also to prove the landing gear system and its interaction with associated components as well. 

These tests are to evaluate performance, dynamics, and system adequacy. As stated earlier, the 

dynamometer tests of the TSO are not an acceptable equivalent or substitute for the airplane 

tests. We also note that an applicant has stated that: "The radial tire exhibits less drag, and 

would therefore result in a slightly longer stopping distance at the same brake pressure (i.e., 

brake torque)." The maximum RTO is conducted at the maximum brake pressure, therefore, the 

braking distance would appear to be longer, based on an applicant's comment. In addition to 

other considerations, the maximum RTO and other performance tests should help establish how 

much longer the braking distance is going to be. 

29
3/14/2000 AC 25-22 

(h) Note that for the same bias tire dynamometer test energy, the brake energy 

increases with a decrease in energy absorbed by the radial tire; therefore, the heat being delivered 

to the fuseplug through the wheel will probably increase. It is not known how much or how 

significant that increase will be. How the fuseplugs are going to act (no-melt/melt) should be 

established on the airplane prior to certification. 

(i) A sufficient number of flight tests should be conducted on wet and dry 

runways to establish that the anti-skid and the landing gear system will function compatibly and 

acceptably with the radial tires. 

(j) It should be established prior to certification that the airplane tire will behave 

structurally and dynamically on the airplane as well as it did on the dynamometer at high speed 

conditions. Because dynamic stability may be more of a function of gear dynamics coupled to 

the tire dynamics, airplane tests should be conducted, as the dynamometer is insufficient in 

simulating airplane and landing gear system dynamics. 

(6) Critical Conditions & Maximum Ramp Weight Definitions. The following is 

extracted from an FAA telegraphic message dated January 26, 1981, which addresses definitions

of "critical conditions" and "maximum ramp weight" as they relate to § 25.733. 

(a) Section 25.733(a) Critical Conditions. The critical conditions are those 

conditions where the takeoff or landing speed extremes that have been established by evaluating 

takeoff or landing performance for temperature, altitude, aircraft weight, aircraft configuration, 

runway conditions, engine performance, etc., may approach the tire speed rating. The critical 

conditions in this case are the takeoff or landing performance conditions that establish maximum 

tire speed. 

(b) Section 25.733(a)(1); (b)(1); (b)(3);Maximum ramp weight. The maximum

ramp weight is the maximum weight allowed on the airplane (includes airplane weight, fuel, oil, 

baggage, etc.). This definition is used interchangeably with maximum taxi weight. 

30 
3/14/2000 AC 25-22 

 e. References. The addresses for ordering the latest revision of advisory circulars, 

technical standard orders, and other referenced documents listed below can be found in the 

Appendix to this AC. 

AC 20-97A, High-Speed Tire Maintenance and Operational Practices. 

AC 43.13-1A, Acceptable Methods, Techniques and Practices - Aircraft Inspection and 

Repair, chap. 8, L. G. Equipment. 

AC 145-4, Inspection, Retread, Repair and Alterations of Tires.

FAA Order 8000.54, Process Specifications for Retreading Tires and Inspection 

Procedures Manuals for Tire Retreaders. 

FAA Order 8000.64, Qualification of Aircraft Radial Tires for Use on Aircraft and for 

Retreading. 

FAA's Tire Approval Process Video, MTS 422/422.1, 42:15 Minutes. 

TSO-C62c, of § 37.167 14 CFR. This TSO was removed from regulations (§ 37.167, 

Amendment 25-52, 1980), and became a voluntary standard. 

TSO-C62c, With Addendum I. This TSO was removed from regulations (§ 37.167, 

Amendment 25-52, 1980), and became a voluntary standard. 

 TSO-C62d, Tires. 

SAE AIR-1904 - Tire Spray Suppression - Airplane, Design Consideration and Testing. 

SAE ARP-1322 - Overpressurization Release Devices. 

SAE ARP-4834 - Recommended Practice for Retreaded Aircraft Tires - Radial and Bias. 

SAE AS-707B - Thermal Sensitive Inflation Pressure Release Devices for Tubeless 

Aircraft Wheels. 

SAE-AS 1188 - Aircraft Tire Inflation-Deflation Equipment. 

SAE AS-4833 - Aircraft New Tire Standard- Bias and Radial. 

14. SECTION 25.735 - BRAKES.

 a. Rule Text.

(a) Each brake must be approved. 

(b) The brake system and associated systems must be designed and constructed so 

that if any electrical, pneumatic, hydraulic, or mechanical connecting or 

transmitting element (excluding the operating pedal or handle) fails, or if any 

single source of hydraulic or other brake operating energy supply is lost, it is 

possible to bring the airplane to rest under conditions specified in 25.125, with a 

mean deceleration during the landing roll of at least 50 percent of that obtained 

in determining the landing distance prescribed in that section. Subcomponents 

within the brake assembly, such as brake drum, shoes, and actuators (or their 

equivalents), shall be considered as connecting or transmitting elements, unless it 

is shown that leakage of hydraulic fluid resulting from failure of the sealing 

elements in these subcomponents within the brake assembly would not reduce the 

braking effectiveness below that specified in this paragraph. 

(c) Brake controls may not require excessive control force in their operation. 

31
3/14/2000 AC 25-22 

(d) The airplane must have a parking control that, when set by the pilot, will 

without further attention, prevent the airplane from rolling on a paved, level 

runway with takeoff power on the critical engine. 

(e) If antiskid devices are installed, the devices and associated systems must be 

designed so that no single probable malfunction will result in a hazardous loss of 

braking ability or directional control of the airplane. 

(f) The design landing brake kinetic energy capacity rating of each main wheelbrake assembly shall be used during qualification testing of the brake to the 

applicable Technical Standard Order (TSO) or an acceptable equivalent. This 

kinetic energy rating may not be less than the kinetic energy absorption 

requirements determined under either of the following methods: 

(1) The brake kinetic energy absorption requirements must be based on a rational 

analysis of the sequence of events expected during operational landings at 

maximum landing weight. This analysis must include conservative values of 

airplane speed at which the brakes are applied, braking coefficient of friction 

between tires and runway, aerodynamic drag, propeller drag or power-plant 

forward thrust, and (if more critical) the most adverse single engine or propeller 

malfunction. 

(2) Instead of a rational analysis, the kinetic energy absorption requirements for 

each main wheel-brake assembly may be derived from the following formula, 

which must be modified in cases of designed unequal braking distributions. 

KE = 0.0443 WV2 

 N 

whereKE = Kinetic energy per wheel (ft.-lb.);

W = Design landing weight (lb.); 

V = Airplane speed in knots. V must be not less than VS0 , the power off stalling 

 speed of the airplane at sea level, at the design landing weight, and in the 

 landing configuration; and 

N = Number of main wheels with brakes. 

(g) The minimum stalling speed rating of each main wheel-brake assembly (that 

is, the initial speed used in the dynamometer tests) may not be more than the V 

used in the determination of kinetic energy in accordance with paragraph (f) of 

this section, assuming that the test procedures for wheel-brake assemblies involve 

a specified rate of deceleration, and, therefore, for the same amount of kinetic 

energy, the rate of energy absorption (the power absorbing ability of the brake)

varies inversely with the initial speed. 

(h) The rejected takeoff brake kinetic energy capacity rating of each main wheelbrake assembly that is at the fully worn limit of its allowable wear range shall be 

used during qualification testing of the brake to the applicable Technical 

Standard Order (TSO) or an acceptable equivalent. This kinetic energy rating 

may not be less than the kinetic energy absorption requirements determined under 

either of the following methods: 

32 
3/14/2000 AC 25-22 

(1) The brake kinetic energy absorption requirements must be based on a rational 

analysis of the sequence of events expected during an accelerate-stop maneuver. 

This analysis must include conservative values of airplane speed at which the 

brakes are applied, braking coefficient of friction between tires and runway, 

aerodynamic drag, propeller drag or powerplant forward thrust, and (if more 

critical) the most adverse single engine or propeller malfunction. 

(2) Instead of a rational analysis, the kinetic energy absorption requirements for 

each main wheel brake assembly may be derived from the following formula, 

which must be modified in cases of designed unequal braking distributions: 

KE = 0.0443 WV2 

 N 

whereKE = Kinetic energy per wheel (ft.-lb.);

W = Design landing weight (lb.); 

V = Airplane speed (knots);

N = Number of main wheels with brakes; and 

W and V are the most critical combination of takeoff weight and ground speed 

obtained in a rejected takeoff. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 

35 FR 5676, Apr. 8, 1970; Amdt. 25-48, 44 FR 68742, Nov. 29, 1979; Amdt. 25-

72, 55 FR 29777, Jul. 20, 1990, as amended by Amdt. 25-78, 58 FR 11781, 

Feb. 26, 1993; Amdt. 25-92, 63 FR 8320, Feb. 18, 1998]

b. Intent of Rule. 

(1) This rule requires use of approved brakes and wheel assemblies, either approved 

under an applicable technical standard order (TSO) e.g., TSO-C26c, or approved under the type 

certificate for the airplane. The existence of TSO approval with the article displaying required 

markings does not automatically constitute the authority to install and use the article on an 

airplane. 

(2) It is the responsibility of those desiring to install this article to determine that the 

aircraft operating conditions are within the TSO standards. Additional requirements may be 

imposed based on airplane specifications, wheel and brake design, and quality control 

specifications. The airframe manufacturer/user must conduct the applicable airplane certification 

tests and receive FAA approval. 

(3) In addition to brake performance and safety requirements, the rule provides 

standards for systems and equipment associated with brakes, e.g., control mechanisms and antiskid systems. The brake and wheel assembly must have proper energy and load ratings for taxi, 

takeoff, refused takeoff, and landing. The braking system must have acceptable pilot-control 

forces, adequate parking brake capability, emergency braking capacity and directional control, 

and protection against overpressure, overtemperature, and fire. 

33
3/14/2000 AC 25-22 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Sections 4b.337 

(a)(1), 4b.337(a)(2) and 4b.337(a)(3), 4b.337(b), 4b.337(c), 4b.337(d), 4b.335(c), and 4b.335(d) 

of the CAR, respectively became §§ 25.735(a), 25.735(b), 25.735(c), 25.735(d), 25.735(e), 

25.735(f) and 25.735(g) of 14 CFR, for brakes/braking systems. 

 (1) Amendment 25-23 (April 8, 1970) deleted reference to a military specification 

(MIL-B-8075) as the means of compliance for antiskid devices under § 25.735(e) to allow any 

other acceptable means of compliance. In addition, proper units of "knots" were added to stall 

speed under § 25.735(f)(2). 

 (2) Amendment 25-48 (November 29, 1979) revised the technical standard order 

TSO-C26b for aircraft wheels and wheel-brake assemblies and related type certification 

requirements for airplane brakes. The revised standards in § 37.172 - TSO-C26c incorporated an 

updated and improved minimum performance standard for the design and construction of aircraft 

wheels and brakes. Under § 25.735(b), the incorrect reference to § 25.75 was replaced by a 

correct reference to § 25.125. Under § 25.735(f)(2), the numerical constant 0.0442 was 

corrected as 0.0443, and the letter "N" was appropriately redefined as the number of main wheels 

with brakes. Under § 25.735(f)(2), the term Vso in the formula was replaced with "V" such that 

V must not be less than Vso. Under § 25.735(g), the term Vso was replaced by V to be 

consistent with terminology used under § 25.735(f)(2). 

 (3) Amendment 25-52 (June 9, 1980) eliminated § 37.172 Aircraft wheels and brakes -

TSO-C26c from the regulations. TSO-C26c was made available at FAA Headquarters in the 

Office of Airworthiness. 

 (4) Amendment 25-72 (July 20, 1990) revised the text of the last sentence in 

§ 25.735(b) to clarify the intent. 

 (5) Amendment 25-92 (February 18, 1998) added § 25.735(h) requiring that the 

maximum rejected takeoff kinetic energy capacity rating of the aircraft brakes be determined 

with the brake at the fully worn limit of its allowable wear range. The requirement evolved from 

an National Transportation Safety Board (NTSB) recommendation as follows: In 1988, a large 

transport category airplane experienced an 86% maximum kinetic energy (KE) rejected takeoff 

(RTO) in a dispatch configuration in which eight of the ten brakes were worn close to the 

maintenance limits. The eight brakes failed early in the braking run and the airplane overran the 

runway. As a result, the FAA reviewed the methodology used in the determination of allowable 

brake wear limits for transport category airplanes. It was determined that brake wear limits 

should be established during certification to ensure that fully worn brakes will function properly 

during a maximum KE RTO. A series of airplane specific airworthiness directives were issued 

between 1989 and 1994 to establish brake wear limits using the new criteria. These criteria were 

applied to transport category airplanes that exceeded a gross weight of 75000 pounds. The test 

criteria to establish brake wear limits are discussed later under Policy/Compliance item d.(6). 

The regulation was updated to reflect this new criteria. 

34 
3/14/2000 AC 25-22 

 (6) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group has recommended 

revisions to §§ 25.731 and 25.735, and has developed a new advisory circular (AC) 25.735-1. A 

new TSO-C135 for transport category airplanes, replacing the applicable parts of the existing 

TSO-C26c, was also developed. The proposed rule change, the proposed AC, and the proposed 

TSO package were published in the Federal Register, August 10, 1999, inviting public 

comments. The proposed harmonized AC 25.735-1 will not, however, replace all policy matter 

contained in this AC because several issues are not addressed in AC 25.735-1. 

 d. Policy/Compliance Methods. Flight and ground testing procedures for demonstrating 

compliance with the requirements of § 25.735 may be found in AC 25-7A, Flight Test Guide for 

Transport Category Airplanes. 

(1) Carbon Brake Refurbishment (redensification). The following is extracted from an 

FAA memorandum, dated April 11, 1996, which provides guidance regarding approval of an 

alternate process for redensification of carbon brake disks during brake overhaul. The 

redensification process involves a 2-for-1 brake disk refurbishment. The term 2-for-1 is 

associated with a re-building of the brake where two worn rotors are resurfaced, placed in an 

oven to have the carbon "redensified" and riveted together to form a new rotor, which is then 

placed in the brake as part of the re-building of the assembly. An approved repair shop that 

overhauls brakes in accordance with the original equipment manufacturer (OEM) overhaul 

procedures, requested the FAA to approve, as a minor change, the option of using a different 

carbon brake redensification process developed by the shop and its associates. The only 

redensification option presently available is to return the brake components to the OEM. The 

proposed process was initially applied to carbon brake disks used on a military aircraft. 

(a) Due to continued in-service problems associated with carbon brakes and their 

complex processing techniques, the FAA contacted major U.S. manufacturers to solicit 

background information regarding carbon brake overhaul procedures including the 2-for-1 

refurbishment and redensification process. All brake manufacturers consider carbon brake 

redensification a complex and proprietary process which can affect brake performance. 

Experience at three (3) brake manufacturers of original equipment has shown that a very minor 

carbon processing change (heat treat operation) can have a significant effect on brake 

performance and airplane structural compatibility. By the criteria established in AC 25-7A, 

Flight Test Guide for Transport Category Airplanes, paragraph 55, it has been determined that 

redensification is a major change. 

(b) For additional guidance on compliance with this requirement, refer to the latest 

revision of AC 25-7A, chapter 4, section 4, paragraph 55. 

(c) The typical dynamometer test program conducted by an OEM for an 

overhauled brake configuration can be quite substantial. An airplane manufacturer usually 

requires that the OEM conduct testing to obtain FAA TSO-C26c approval and re-qualify the 

brake before airline introduction and retrofit. This full scale brake requalification effort depends 

35
3/14/2000 AC 25-22 

on the type and degree of modification(s) done to the brake to obtain the overhauled 

configuration. 

 (d) An airplane manufacturer pays particular attention to potential differences in 

brake torque, energy capacity, vibration, brake sensitivity, dynamic response, and structural 

strength. A brake requalification test program including the conditions below would be 

accomplished using a wheel/brake assembly in the most adverse service condition. 

 1 Design Landing Stops (Normal and Overload Energy levels). 

 2 Wet Brake Landing Stops (effect of moisture content in carbon). 

 3 New Brake Rejected Takeoffs (RTO's) at 60%, 80%, and 100% of 

maximum brake kinetic energy). 

 4 Worn Brake RTO's. 

 5 Wet Brake RTO's. 

 6 High Power/Torque Repeatability RTO. 

 7 Torque Sensitivity Testing. 

 8 Service Landing Wear Testing. 

 9 Fuse Plug No-Melt RTO series to evaluate disk warpage and thermal 

characteristics. 

 10 Brake Rolling Drag. 

 11 Brake Peak Torque Survey. 

 12 Static Low Speed Torque. 

 13 Vibration Testing. 

 14 Frequency Response. 

 15 Torque Recovery. 

 16 Hysteresis. 

 17 Static Structural Torque. 

 18 Dynamic Structural Torque Test. 

 19 Static Brake Torque to evaluate park brake hold capability. 

 20 Carbon Disk Oxidation Resistance. 

 21 Heat Sink Contamination Testing. 

 22 Intermix of New and Overhauled disks. 

 23 Brake Thermal Modeling. 

(e) The test results are then compared to the original new brake qualification test 

results to determine equivalence of the overhauled brake configuration. When changes include 

different heat sink material processing or are done by a supplier other than the OEM, it may be 

36 
3/14/2000 AC 25-22 

necessary to perform airplane/brake system flight testing. This will validate the Airplane Flight

Manual performance and the brake control system compatibility for the overhauled brake 

configuration. 

(f) Some additional considerations are as follows: 

 1 The FAA does not accept military airplane equipment test data for 

compliance with transport category airplane regulations. 

 2 The applicant has not conducted any tests whatsoever on the modified 

brakes. 

 3 The applicant may need to apply for a new TSO and possibly a 

supplemental type certificate (STC). 

(2) FAA/JAA certification requirements: steel/carbon brakes. The following is 

extracted from an FAA letter dated April 9, 1996, which addresses obtaining parts manufacturer 

approval (PMA) for manufacturers located outside of the United States, TSO procedures, and 

brake certification requirements for steel and carbon brakes. 

(a) There are no differences in the certification requirements between carbon and 

steel brakes. The TSO-C26c and AC 25-7A, Flight Test Guide for Certification of Transport 

Category Airplanes, are still applicable. In addition, each airframe manufacturer has specific 

vibration control criteria in their wheel and brake specification. This criteria is crucial to the 

brake dynamics, the particular airplane installation, and structural integrity of the airplane. 

(b) There are no known differences on the carbon brake certification requirements 

between the FAA and the Joint Aviation Authorities (JAA). You are, however, advised to verify 

this with the JAA. 

(3) Certification of Replacement Brakes. The following is extracted from an FAA 

letter dated May 19, 1992, which addresses policy on aircraft brake replacement. The 

information was requested by a foreign airworthiness authority and is provided in the advisory 

circular format. 

ADVISORY INFORMATION FOR BRAKE REPLACEMENT

 (a) Purpose. To provide advisory information and guidance for a non-U.S. 

manufacturer concerning an acceptable means of showing compliance with Title 14, Code of 

Federal Regulations (14 CFR), applicable to the replacement of brakes for part 25 airplanes. 

37
3/14/2000 AC 25-22 

(b) Related Regulations and Documents. 

 1 Sections. The related sections of 14 CFR, part 25, are as follows: 

 § 21.305 Approval of materials, parts, process, and appliances. 

§ 21.502 Approval of materials, parts, and appliances. 

 § 21.607 General rules governing holders of TSO authorizations. 

 § 21.609 Approval for deviation. 

§ 21.617 Issue of letters of TSO design approval: import appliances. 

 § 25.109 Accelerate-stop distance.

 § 25.125 Landing. 

 § 25.143 Controllability and maneuverability. 

 § 25.231 Longitudinal stability and control. 

§ 25.233 Direction stability and control. 

 § 25.493 Braked roll conditions. 

 § 25.735 Brakes. 

 § 25.1301 Function and Installation. 

 § 25.1309 Equipment, systems, and installations. 

 2 Advisory Circulars (AC's). The addresses for ordering the latest revision 

of the advisory circulars and technical standard orders listed below can be found in the Appendix 

to this AC. 

AC 21-23, Airworthiness Certification of Civil Aircraft Engine, Propellers, and 

Related Products. 

AC 25-7A, Flight Test Guide for Certification of Transport Category 

Airplanes. 

TSO-C26c, Aircraft Wheels and Wheel-Brake Assemblies with Addendum I. 

 3 Industry Documents. These documents provide additional information, 

guidance, and/or standards. The address to order the documents listed below can be found in the 

Appendix to this AC. 

SAE ARP-597C, Wheels and Brakes, Supplementary Criteria for Design 

Endurance - Civil Transport Aircraft. 

SAE ARP-813A, Maintainability Recommendations for Aircraft Wheels and 

Brakes. 

SAE ARP-1064B, Brake Dynamics. 

SAE AS-1145A, Aircraft Brake Temperature Monitor System (BTMS). 

SAE ARP-1619, Replacement and Modified Brakes and Wheels. 

38 
3/14/2000 AC 25-22 

 (c) Airworthiness Requirements. 

 1 Applicable Requirements. Replacement of brakes must be evaluated 

against the applicable part 25 airworthiness requirements. They are §§ 25.109, 25.125, 25.143, 

25.231, 25.233, 25.493, 25.735, 25.1301, 25.1309, of Title 14 Code of Federal Regulations 

(14 CFR), and TSO-C26c. 

 2 Substantiation Procedures: 

(aa) The design and its design deviations should be substantiated by 

conducting the necessary analytical investigations, laboratory testing, and airplane testing to 

ensure that the change can be made without impairing airplane braking or rolling performance. 

The recommended substantiation procedures for replacement brakes are based on parts approval 

and their impact on the prior certification. Depending upon the type and extent of change and 

the certification basis of the airplane, engineering judgment will be used to determine which of 

the substantiation procedures discussed in AC 25-7A, Flight Test Guide for Certification of 

Transport Category Airplanes, chapter 4, section 4, paragraph 55, should be considered for 

replacement brake installations in addition to the laboratory tests required by the TSO. 

 (bb) Aircraft brakes are approved under the Technical Standard Order 

(TSO) system (TSO-C26c). A TSO-C26c authorization indicates that the brake and its 

associated wheel have met certain requirements relating to strength, energy absorption, 

durability, construction, performance, and manufacture. In addition, the holder of a TSO 

authorization has demonstrated certain quality control procedures and the ability to produce 

duplicate parts from approved drawings. 

(cc) A TSO authorization, by itself, is generally considered adequate to 

show that the wheel/brake combination is satisfactory for flight testing on an airplane, and would 

be safe for use during subsequent certification testing. However, it is not sufficient to show 

satisfactory performance within the operational envelope of the airplane, due to the welldocumented differences (load distribution on the landing gear, heat dissipation, brake efficiency, 

runway crown, etc.) between laboratory dynamometer testing and full-scale airplane testing. 

(d) Continued Airworthiness/Brake Wear Limits. 

 1. During previous certification activity, the determination of airplane 

performance and the demonstration of maximum brake kinetic energy capability was based on 

testing brakes at or near their new state. In 1988, a transport category airplane experienced a 

rejected takeoff (RTO) at 86% of the brake's maximum kinetic energy capability. The airplane 

went off the end of the runway. An investigation determined that eight of the ten brakes were 

near the maximum allowable wear limits before the RTO and were unable to absorb the required 

energy, thus contributing to the accident. 

 2 Since this accident, the FAA has reviewed the methodology used in the 

determination of allowable wear limits and airplane performance for transport category airplanes. 

39
3/14/2000 AC 25-22 

The FAA has issued airworthiness directives (AD's) that require more conservative brake wear 

limits on various models of US and foreign transport category airplanes having a U.S. type 

certificate. 

(4) Parking Brake Testing. The following is extracted from an FAA message dated 

March 20, 1992, which addresses the intent and application of § 25.735(d). 

(a) The parking brake requirement was added to the CAR under § 4b.397 of the 

CAR on September 1, 1949, and changed to § 4b.337(c) on July 20, 1950. Section 4b.337(c) of 

the CAR became § 25.735(d) of Title 14, Code of Federal Regulations (14 CFR). The guidance 

material in § 4b.337-1 indicated this was to be a demonstration to determine that sufficient 

braking was maintained to prevent the airplane from rolling on a paved runway while takeoff 

power was applied on the most critical engine. 

(b) The FAA certification offices agree that the test is intended to demonstrate that 

the parking brake will hold with the power of the critical engine at takeoff power and the other 

engine(s) at idle. The effects of propeller wash or propeller/engine torque, in addition to 

asymmetric thrust and aircraft weight, should be used in determining the critical engine and 

vertical loads on the main landing gear. The vectored thrust point can be outside the propellers 

axis of rotation. Therefore, propeller rotation and engine inclination are a factor in determining 

the critical engine. 

(c) To test the maximum torque for the parking brake, the airplane should be 

configured at its highest gross weight with the flaps retracted. The critical engine for a four 

engine airplane is usually one of the outboard engines. For a two engine executive jet airplane 

with near centerline engines, or a twin engine propeller airplane with propellers rotating the 

opposite direction, it may not make a difference which is the critical engine. For some executive 

jets, where there wasn't considered to be a critical engine relative to § 25.735(d), the 

manufacturers tested by putting both engines at takeoff power. This is not being tested according 

to the rule, but it should lead to the same conclusion. 

(d) For consistency, the following test procedure is recommended: 

Paint a radial stripe(s) on the wheels. With the airplane sitting on dry pavement at the airplane's

maximum takeoff weight, the tires pressurized normally for this weight, the flaps retracted and 

control surfaces centered (neutral), conduct the parking brake test as follows: 

 1 Set the parking brake. 

 2 Apply takeoff power to the critical engine (critical relative to § 25.735(d)) 

leaving the other engine(s) at idle. 

 3 An external observer, looking at the white stripe(s) on the wheels, verifies 

the wheels are not turning, i.e., the brakes are locked. The airplane may skip or shear tire tread 

or the tire may slip around the wheel, but the wheels must not rotate. 

40 
3/14/2000 AC 25-22 

(e) If it has not been established which engine is critical, conduct a separate test 

for each engine. If it has been determined there isn't a critical engine for a twin engine airplane, 

a test with either engine at takeoff power will suffice. 

(5) Worn Brake Requirements for Non-Original-Equipment Parts. The following is 

extracted from an FAA letter dated February 18, 1992, which addresses an original equipment 

manufacturer (OEM) inquiring whether a non-OEM was required to follow the same procedures 

and conduct the same tests as the OEM. The FAA response was affirmative and explained 

below. 

(a) An FAA issued AD proposes brake wear limits for specific series airplanes. 

This proposal required that all landing gear brakes be inspected for wear and be replaced if the 

wear limits prescribed in this proposal are not met, and that the new wear limits be incorporated 

into the FAA-approved maintenance inspection program. An original equipment manufacturer 

asked that the AD require the following for non-original equipment (NOE) certified parts: 

 1 NOE manufacturers submit worn brake RTO test results for all brake 

assemblies using their parts. 

 2 Brake assemblies containing NOE parts be given new part numbers and 

TSO markings. 

 3 Installation approval and notification of all operators and overhaul shops 

be required concerning these approvals, and 

 4 Compliance with the terms of the AD be required for anyone using or 

furnishing parts for brakes. 

(b) The FAA position was that regulations, orders and advisory material provide 

guidance concerning changes to approved parts, components and assemblies and that events 1, 2, 

3, and 4 occur. As an example, a brake assembly was approved with NOE rotors under a 

supplemental type certificate (STC) project. The Aircraft Certification Office, which issued the 

STC, required the company manufacturing the brake assembly meet all applicable worn brake 

RTO requirements. 

(6) Determination of Allowable Worn Brake Limits. The following is extracted from

an FAA memorandum dated February 23, 1990, which provides guidance regarding worn brake 

capability determination. A policy letter dated March 2, 1990, contained the same information. 

Issue papers were used to administer the same policy for new certification programs until 

Amendment 25-92 was issued in February 1998. 

(a) This memorandum established criteria for worn brake testing for in-service 

transport category airplanes. Worn-brake accountability determination was intended to validate

brake wear limits with respect to brake energy capacity only, and was not meant to account for 

41
3/14/2000 AC 25-22 

any reduction in brake force due solely to the wear state of the brake. This criteria is intended 

for transport category airplanes used in part 121 operation, and which have a certificated 

maximum takeoff weight in excess of 75,000 pounds. 

(b) It should be noted that this worn-brake accountability determination will 

validate brake wear limits with respect to brake energy capacity only, and is not meant to 

account for any reduction in brake force due solely to the wear state of the brake. Any reduction 

in brake force (or torque) that may develop over time as a result of brake wear is to be evaluated 

and accounted for as part of the rulemaking project described in (c) below. The accident that 

prompted this investigation into brake wear is believed to have been caused by inadequate brake 

energy-absorbing capability. The issue of loss of brake effectiveness with brake wear, and the 

resulting increase in runway lengths required for a RTO is addressed under Amendment 25-92. 

(c) The following dynamometer test guidelines are the result of a series of 

meetings. These test procedures may be used for the determination of acceptable airplane brake 

wear limits.

 1 Acceptable Test Brakes.

 (aa) Either airplane-worn or mechanically-worn brakes may be used. 

Mechanically-worn is defined as not being airplane-worn, e.g., machined or dynamometer-worn. 

If mechanically-worn brakes are used, it must be shown that they can be expected to provide 

similar results to airplane-worn brakes. 

(bb) Each test brake shall be subjected to a sufficient number and type of

stops to ensure that the brake's performance is representative of in-service use. 

 2 Wear State of the Test Brake. 

(aa) Degree of Wear. The degree of wear of the test brake shall be 100 

percent. One hundred percent worn is defined as that degree of wear which the applicant intends 

to allow before the brake is to be removed for overhaul. At the overhaul limit the brake will not 

be fully worn out, but will contain sufficient braking capability to meet the energy absorption 

requirements discussed below. The chosen test brake shall be such that the wear-in conditions 

produce a brake ready to perform the RTO test at the correct wear setting. If a brake to be tested 

is worn less than 100 percent, an acceptable method of extrapolation to the fully worn state must 

be provided prior to test. 

(bb) Definition of Degree of Wear. The degree of wear shall be defined 

in terms of the linear, axial direction dimension relating to the allowable wear of the brake as 

commonly determined by noting wear pin extension. 

(cc) Distribution of Wear (Applicable only to mechanically-worn brakes). 

The proportioning of the wear through the brake for the various friction pairs shall be based on 

42 
3/14/2000 AC 25-22 

either: Service experience on the test brake or an appropriate equivalent brake, or Dynamometer 

wear test data. 

 3 Brake Energy Level to Set 100 Percent Wear Limits. 

(aa) It will be necessary to conduct the dynamometer test with an initial 

energy value prior to the RTO test that is analogous to that generated during the airplane 

certification flight test of that brake (e.g., energy due to taxi stops). 

(bb) The dynamometer RTO maximum energy level may be based on the 

Airplane Flight Manual limiting conditions, including, if desired by the applicant, the effects of: 

 Engine reverse thrust, following normal procedures for power 

setting, cutback speed, and 

The recommended number of reversers to be used with a critical 

engine inoperative, and the demonstrated transition times achieved in or equivalent to an RTO 

flight test. 

(cc) The dynamometer test should be conducted to duplicate the airplane's

speed versus time performance as closely as practicable. 

(dd) The use of reverse thrust shall be limited to a determination of the 

energy to be absorbed by the brake during the dynamometer test. No thrust reverser credit shall 

be given for the purpose of determining runway distance requirements or airplane stopping 

performance. 

 4 Power Level. The test shall be conducted at either of the conditions 

below, provided that the test is conducted at the condition which more closely represents the 

actual braking conditions obtainable on the airplane. The intent of these procedures is to 

simulate actual airplane conditions as closely as possible: 

(aa) The maximum brake pressure, or 

(bb) The maximum tire drag or brake torque consistent with the airplane's

hydraulic system and any antiskid and/or torque limiter pressure limitations that would occur on 

the airplane during an equivalent RTO operation. 

43
3/14/2000 AC 25-22 

(cc) In order to assure that the test results are accurate, repeatable, and 

simulate actual airplane conditions as closely as possible, the dynamometer tests must be 

conducted on an inertial-wheel type dynamometer, using a suitable tire/wheel/brake assembly. If 

use of a shaft-type dynamometer is desired, it must be shown prior to any testing, that the shafttype dynamometer will provide results equivalent to that provided by the inertial-wheel type 

dynamometer. 

 5 Final Condition Definition. 

(aa) A full-stop demonstration is not required for the worn-brake RTO 

test. The test brake pressure may be released at a dynamometer speed of up to 20 knots to 

facilitate a detailed post-test inspection of the brake. The dynamometer test may be started at a 

slightly higher speed so that the test may be terminated at 20 knots or less, provided that the data 

submitted for each test show that the energy absorbed by the brake during a test that is

terminated at 20 knots or less, is equal to the energy that would have been absorbed if the test 

had been started at the proper speed and continued to zero ground speed. 

(bb) There shall be no wheel burst as a result of this test. 

 6 Data Requirements. 

(aa) As a minimum, the following technical data shall be obtained for 

each dynamometer test conducted: 

 Brake torque (or force).

 Brake pressure. 

 Time.

 Road wheel speed. 

 Road wheel distance. 

 Dynamometer inertia equivalent. 

(bb) The absorbed dynamometer kinetic energy shall be computed based 

on measured data. Additional data may also be obtained to aid in interpolating and extrapolating 

test results. 

(cc) A test report shall be prepared which, as a minimum, shall include: 

A detailed description of the test article (e.g., component part numbers, 

individual disk measurements, wear pin measurements, etc.), 

 The test procedures, 

 The test results, and 

 If possible, photographs of brake rotors/stators before and after the 

dynamometer test. 

(dd) In order to receive credit for the effects of (N-1) thrust reversers, 

where N is the number of reversers available at dispatch, thrust reverser effectiveness must be 

44 
3/14/2000 AC 25-22 

substantiated. Any existing validated reverser performance data are to be submitted for FAA 

review and approval. This is required for each engine family for each airplane model. 

 7 Interpretation of Data. Any adjustment of energy levels or allowable wear 

from the dynamometer test shall be based on a review of the test data, inspection of brake 

hardware after test, and subsequent analysis. 

(aa) An extrapolation of wear data or energy data up to 5 percent of the 

test values shall be permissible. 

(bb) An interpolation of data up to 20 percent of the test values shall be 

permissible to establish energy levels from multiple dynamometer tests which are within this 

range from the target condition. 

 8 Acceptance of Prior Tests. 

(aa) Worn-brake RTO tests which have been conducted successfully prior 

to the adoption of these procedures may be acceptable. These tests need not be repeated solely to 

gather test data specified here. 

 (bb) These worn-brake dynamometer test guidelines are a recommended 

test procedure, and, as such, represent one means, but not necessarily the only means, of 

determining acceptable maximum brake wear. The FAA would consider deviations from these 

guidelines if it can be shown that the proposed procedure is appropriate and would produce 

equivalent results. 

(cc) The dynamometer guidelines listed above indicate that credit for 

reverse thrust, using the recommended number of thrust reversers with a critical engine 

inoperative, is to be allowed in the determination of the maximum brake energy on the 

dynamometer. The effect of inoperative thrust reversers due to Minimum Equipment List (MEL) 

dispatch must also be accounted for. 

(dd) If a particular brake is affected by the Airworthiness Directive (AD), 

and is later tested on the dynamometer and shown to require a different wear pin cut-back than 

that specified in the AD, then the AD should be amended at that time to reflect the correct wear 

pin length. The continued airworthiness aspects of this statement apply to all brakes whether or 

not an AD was issued. 

(ee) On any given airplane model there may be many different brake 

configurations that are quite similar, and identified by different dash numbers. It may be 

acceptable for the manufacturer to demonstrate acceptable brake wear limits on one brake 

representative of a brake family, and extend that brake wear limit to others of the same family by 

analysis. Each case will have to be evaluated on its own merits. 

45
3/14/2000 AC 25-22 

(ff) There may be situations in which a brake modification or brake 

installation is proposed by someone other than the airframe manufacturer or the original brake 

manufacturer. In these cases, if the brake modification is judged to be a major change involving 

the friction couple (rotor/stator), or in any way could affect brake energy capacity, a 

determination of the effects of brake wear in accordance with these procedures must be 

accomplished. The applicant may claim the same brake wear limit as the original brake 

manufacturer, and this may be acceptable, but this would have to be demonstrated by analysis or 

test. 

(gg) A project to amend part 25 to include a determination of the effects 

of brake wear on brake force has been initiated. At this time it has not been determined what the 

format for this determination will be. For the present time, any certification project involving 

new brakes or a substantial redesign of existing brakes (major change) should be conducted as 

described in AC 25-7A (formerly Order 8110.8), i.e., airplane flight tests should be conducted 

using new brakes. An adjustment to the allowable brake wear limits should then be made, if 

applicable, using the procedures described above. 

(7) Credit for Reverse Thrust in Worn Brake Testing. The following is extracted from

an FAA memorandum dated April 28, 1989, which established the acceptability of reverse thrust 

credit for the purposes of establishing worn brake limits only. 

(a) During certification of aircraft brake systems, a demonstration of stopping 

performance at the maximum takeoff weight and speed expected in service is conducted. This 

RTO test shows that the brake has the capacity to absorb the required kinetic energy, and may 

limit certain performance information in the Airplane Flight Manual. This test is conducted with 

new brakes and without reverse thrust or other deceleration means other than spoilers. 

(b) Accident data has shown that the brake-wear limits and brake design in general 

may be inadequate to provide the required stopping performance at maximum RTO energy levels 

with worn brakes. 

(c) An adopted rule AD that reduces allowable brake wear prior to overhaul on the 

XYZ model airplane has been issued. This AD was based on dynamometer testing, and provides 

for a brake overhaul interval such that sufficient brake mass remains at overhaul to absorb the 

certificated RTO energy and allow the airplane to stop on the runway. 

 (d) FAA/Industry meetings were held to develop a dynamometer test protocol 

prior to any testing, so as to minimize costs and delays. A major issue, which arose during these 

meetings, is credit for the effects of reverse thrust on the kinetic energy that must be absorbed by 

the brakes. Because of the way brake energy for the initial dynamometer tests was calculated, 

46 
3/14/2000 AC 25-22 

the initial AD included credit for (N-1) thrust reversers, i.e., 2-engine reverser credit for the 

3-engine aircraft. As you know, the FAA has been reluctant in the past to grant credit for reverse 

thrust relating to braking performance or brake kinetic energy determination. We remain 

unconvinced that reverse thrust credit should be granted in the case of airplane braking 

performance, because reverse thrust has not been shown to be sufficiently reliable for this 

purpose. However, the FAA has no valid reason to deny the use of reverse thrust in the 

determination of brake kinetic energy for brake wear limit tests. Reverse thrust will certainly be 

used in any RTO (if available), and the probability of a high-energy, field-length-limited RTO 

combined with an inoperative reversers is sufficiently low to allow relief in this area. It will be 

acceptable to allow credit for the effects of (N-1) thrust reversers in the determination of 

allowable wear limits in dynamometer tests discussed above. 

(e) It should be noted that this use of reverse thrust applies to kinetic energy 

determination for the purpose of establishing worn brake limits only. The FAA will not consider 

allowing credit for reverse thrust in the determination of accelerate-stop distances or braking 

performance. 

(8) Certification of Brake Replacement Components. The following is extracted from

FAA letters and memorandums dated March 16, 1989, December 27, 1988, and March 16, 1988, 

which address the certification data/testing required for obtaining FAA approval for replacement 

parts of an existing FAA approved brake installation. 

(a) In early 1989, the FAA issued an STC to provide replacement brake parts for 

certain part number brakes on a specific airplane. Throughout the STC review process, the FAA 

was acutely aware that granting this STC would be precedent-setting. Brakes to date had been 

designed and approved as a system, involving the wheel/brake assembly, the tire, and the 

airplane landing gear and brake systems. This approval involved a component of a previously 

approved system, and as such, the approval method for this modification was necessarily 

different from that applied to the original manufacturer. 

(b) The key issue to be addressed was the amount of product development and 

testing that must be conducted in order to demonstrate that the proposed parts are satisfactory for 

their intended use. An original brake manufacturer must perform considerable research and 

development in order to design an aircraft brake with the best compromise of low weight, 

durability, useful life, and performance. The interface between the airplane structure and 

systems and the proposed brake must be investigated, and this sometimes requires much 

dynamometer and airplane flight testing. Early in-service experience with the new brake may 

reveal the need for minor "fine tuning" in the brake design. 

(c) If another manufacturer later desires to produce any of the components of an 

approved assembly under an STC or a Parts Manufacturer Approval (PMA), he or she must show 

that the airplane, as modified to incorporate the new components, continues to comply with the 

regulations incorporated by reference in the type certificate for that airplane. This does not 

necessarily entail following the same approval path as that of the original manufacturer. In many 

cases, he or she need only duplicate the part in question, and if the part can be shown to be 

47
3/14/2000 AC 25-22 

"identical" to the original part, FAA approval can be granted on that basis. This method of 

approval of replacement parts has a long, successful history and will continue to be used when 

appropriate. The original manufacturer of a TSO-authorized part does enjoy certain advantages 

over another manufacturer that seeks to modify that part. He may make modifications to the part 

based on the experience gained during the product development without extensive testing, and 

have the change approved as a "minor" change. The same change proposed by another applicant 

might be judged to be a "major" change, which would necessitate considerably more testing. 

(d) In this case, the STC applicant originally sought approval of their part based on 

identicality with the original part. A finding of identicality could not be made, however, due to 

some small differences between the parts. The differences existed and could not be ignored, but 

were judged to be so insignificant as to obviate the need for the full spectrum of ground and 

flight tests discussed in AC 25-7A, Flight Test Guide for Transport Category Airplanes (formerly 

Order 8110.8), chapter 4, section 4, paragraph 55. It was agreed by the FAA and the STC 

applicant that a back-to-back comparison of the original brake and the company's modified brake 

on a flywheel-type dynamometer at the ultimate energy capacity of the original brake would be 

an acceptable test to show equivalent performance. New OEM brakes would be tested at 

successively higher energies until brake failure occurred, or was imminent; the applicant's 

modified brake would then be tested at this energy, with a second modified brake tested for a 

demonstration of repeatability. 

(e) Although, by its very nature this dynamometer test alone could not support a 

"major" brake change (as discussed in AC 25-7A, formerly Order 8110.8), it would demonstrate 

that the ultimate kinetic energy (KE) capacity of the replacement brake part and the OEM brake 

part were equivalent under identical laboratory test conditions. This, together with the near 

identicality of the two components, was the basis for the FAA approval. No attempt was made 

to duplicate any conditions relating to on-airplane brake testing. The well-documented 

differences between dynamometer testing and airplane testing prevented any such comparison. 

It is possible that this series of dynamometer tests was a more severe test than a single RTO, 

since the brake was tested to destruction in the laboratory under controlled conditions, and its 

behavior could be more closely monitored than on the airplane. The FAA was aware of the 

torque and deceleration window mentioned by the OEM, and the replacement part was shown to 

perform within that envelope satisfactorily. 

(f) The OEM was concerned that the friction couple was a key factor in the 

determination of the performance of the brake, and the FAA concurred. In future applications 

for similar approvals, either by the STC applicant or other manufacturers, the FAA expects to 

apply the test regimen discussed in FAA Order 8110.8. The dynamometer test used in this case 

would only be used if the situation were essentially the same as this case ; i.e., a part that is so 

nearly identical with the original that there is little question that the part would perform the same

as the original, and the FAA was satisfied with a demonstration of manufacturing competence by

the STC applicant. 

(g) Another issue raised by the OEM was the unilateral FAA approval of the 

replacement part without the concurrence of the airframe manufacturer. This is the normal 

48 
3/14/2000 AC 25-22 

method of FAA approval for STC applications. Data supplied to the FAA by the STC applicant 

is, in most cases, proprietary to the applicant, and the FAA is obliged not to reveal details of the 

application to another manufacturer. The FAA is in a position to have access to pertinent data 

from the original airframe manufacturer, and questions of compatibility can be answered without 

involving the original manufacturer. It is unreasonable to expect assistance from a manufacturer 

if that assistance would benefit a competitor or interfere with their own interests. There are 

numerous examples of major modifications to airplanes that have been completed without the 

assistance or cooperation of the original airframe manufacturer. In this particular case, the 

replacement part manufacturer unsuccessfully attempted to obtain technical data from both the 

airframe and brake manufacturer, as well as from the FAA, under the Freedom of Information 

Act. 

(h) During a meeting between the OEM and the FAA, the issues of product 

support, configuration management, and product maintenance were discussed at some length. 

Product improvements are a normal part of any manufactured product, and we expect the OEM 

brake that was the basis for the STC approval will continue to evolve as the market changes and 

service history dictates. In each case, the burden will be on the STC holder to demonstrate that 

their STC is satisfactory as a replacement for any new OEM brake part numbers, if they choose 

to make such an application, and that applicability must be demonstrated, either by the 

dynamometer test method discussed above, or other test requirements dictated by the particular 

situation under consideration. It is possible that future changes to the OEM brake under 

consideration here may make the replacement part STC obsolete. 

(i) We concur with the OEM’s statement that the performance of the proposed 

replacement brake component must necessarily be identical to the original. In many other areas, 

the FAA will accept a demonstration of performance that is "equal to or greater than" the 

original, since many of our requirements are intended to be minimum performance standards. In 

this case, improved brake performance may actually be a detriment, because landing gear 

structure, and brake and antiskid systems are often tuned to brake performance, and a change in 

braking μ or other brake characteristics may have an adverse affect on the airplane itself. If the 

performance of the replacement part is shown to be different in any way from the original, the 

applicant would be required to show that the effect on the airplane would be acceptable. 

(j) In 1987 the FAA convened a special technical review team to examine the 

facts surrounding the original application for approval of replacement brake components. The 

team reviewed the brake approval procedures that had been applied in the past, including TSO 

approvals, dynamometer testing, and airplane flight testing. The team also considered 

information presented by the STC applicant. The consensus of the team is as follows: 

 1 The distinction between “major” and “minor” changes to brakes and brake 

systems, as discussed in AC 25-7A, chapter 4, section 4, paragraph 55, (formerly FAA Order 

8110.8), is valid, and should continue to be followed. The discussion and examples given in 

paragraph 55 are generally adequate to determine the proper course of testing to be followed in 

the certification of brake changes. 

49
3/14/2000 AC 25-22 

 2 The applicant had failed to establish identicality of the proposed 

replacement part with the originally approved brake component. There was also the possibility 

that the change could impact the airplane stopping performance, or the kinetic energy absorption 

characteristics of the brake. Such a design change would normally lead to a classification as a 

major change requiring a demonstration of brake performance on the airplane, including a 

maximum energy RTO. 

 3 The review team concluded, however, that the proposed replacement 

brake part could be properly considered a minor change, on the basis of demonstrated 

equivalency, as supported by additional analysis and tests. This means that the testing 

requirements of a major change need not apply, if functional airplane landings and comparative 

dynamometer tests could be successfully completed. Given the close similarity of the 

replacement part and the original, the inherent simplicity of the brake part (as compared to the 

stator with its friction linings, for example), and the applicant's long experience in producing 

acceptable brake rotors for the U. S. Air Force, it was determined that the procedure proposed by 

the FAA for dynamometer testing of the applicants part was acceptable. This procedure contains 

the following items: 

(aa) Use of new stator and rotor parts in both the applicant's and OEM’s 

brakes for each test will be required, in order to minimize test configuration variables. If rebuilt 

or in-service components other than these fail during testing, it should be realized that the results 

of the test may be questionable. Suspect tests would be carefully scrutinized by the FAA and 

retesting may then be necessary. 

(bb) A series of tests may be necessary to find the ultimate KE level of 

the OEM brake, as agreed upon in the original test plan. For each succeeding run, the KE will be 

increased by 5% over the previous run until the ultimate KE level is determined, i.e., brake 

failure has occurred, or is imminent. The initial kinetic energy level for this series of tests will 

be at the discretion of the applicant. After each test run, the brake will be examined and a 

determination will be made whether additional tests at higher energies are necessary. 

(cc) A minimum of two test runs at this ultimate energy level must be 

conducted on the OEM brakes for validation of this level prior to conducting the applicant's 

comparison tests. These two runs must show similar results. This provides confidence similar to 

that obtained from multiple brake assemblies on actual airplane tests. 

50 
3/14/2000 AC 25-22 

(dd) At least one run at this ultimate energy level must then be run on the 

applicant's brake assembly, i.e., a back-to-back demonstration of the two brakes. The need for 

additional test runs on the applicant's brake will be determined after the first test. If the 

applicant's brake shows good correlation to the OEM brake, additional tests will not be

necessary. 

(ee) The correlation factor between the OEM and the applicant's tests 

must be such that the results of the applicant's tests can be shown to be equivalent to or better 

than the OEM tests. 

(ff) Constant maximum braking force pressure appropriate to the airplane

braking system must be applied during the tests. 

(gg) Fuse plugs may be released or the tire deflated after each test run, to 

reduce the risk to test personnel. 

(hh) The functional landings already completed were also a required part 

of this approval. 

 4 This additional dynamometer testing in lieu of actual airplane testing was 

applied in this case due to the unusual circumstances relating to the applicant's application. The 

applicant had considerable manufacturing experience with brake parts, indicated by long and 

trouble-free service history in the Air Force inventory. While this Air Force service history was 

not directly applicable to commercial operations, and would not in itself be a basis for FAA 

approval, it did indicate that the applicant's parts performed adequately in service. This gave a 

measure of confidence that the ultimate energy dynamometer testing to be performed would be 

successfully completed. 

 5 It should not be inferred from this action taken on this specific brake part 

proposed by this applicant that this dynamometer test procedure will be applicable to any other 

application, either from this applicant or other manufacturers. Each request for approval should 

be judged on its own merits. 

(9) Ground Equipment - Brake Cooling Unit. The following is extracted from an FAA 

letter dated December 27, 1988, which addresses certification of an aircraft wheel/brake cooling 

unit. 

(a) This type of equipment is not approved as part of the aircraft type certificate. 

The FAA does not impose any technical standards on most ground-support equipment that is 

used to support airline operations. Ground service trucks, baggage carts, and the device 

proposed by the applicant fall into this category. The only exception to this is if the particular 

item of ground equipment performs a critical function, or could cause damage to the airplane if 

used 

51
3/14/2000 AC 25-22 

incorrectly. Examples of the latter include portions of jacking equipment which must mate with 

jacking fittings on the airplane, towing bars that attach to the nose gear, and fuel systems on fuel 

trucks to ensure cleanliness of the fuel. 

(b) Airlines that choose to use the wheel/brake cooling device proposed by the 

applicant would be responsible for showing that the device is appropriate for its intended use in 

their operations and would not pose a hazard to their personnel or the airplane. In addition, 

performance credit for reduced ground turn-around time (required time for brake cooling prior to 

the next takeoff) cannot be granted without an evaluation of the effectiveness of the device by 

the appropriate FAA Aircraft Certification office. 

(10) Certification Requirements for Asbestos-Free Brake Linings. The following is 

extracted from an FAA letter dated December 16, 1987, which addresses guidance for 

certification of replacement brakes with asbestos-free linings, based only on dynamometer and 

TSO testing.

(a) The testing requirements in AC 25-7A, Flight Test Guide for Transport 

Category Airplanes (formerly FAA Order 8110.8, paragraph 89(b)), relate to the addition of a 

highly modified brake design to an existing type certificated airplane. A highly modified brake 

is defined in this section as one which contains new or modified parts which in turn will cause a 

significant variance in the kinetic energy absorption characteristics of the brake. It has been our 

experience that airplane performance cannot be adequately predicted from the dynamometer 

testing specified in the TSO. While it may be satisfactory to assign some performance penalty to 

compensate for this, there remains the question of the effect of the change in friction materials on 

fuse plug performance, ground handling, and brake sensitivity. AC 25-7A, chapter 4, section 4, 

paragraph 55, issued March 31, 1998, contains relevant portions of canceled FAA Order 8110.8, 

and additional information. 

(b) In addition, the coefficient of friction wear rates of the new materials may 

prove to be different from the original. In order to adequately demonstrate the performance of a 

major change to the brakes or brakes control system, a maximum kinetic energy RTO, as 

described in AC 25-7A (formerly Order 8110.8), would normally be required, as well as six 

functional takeoffs and landings. In addition, fuse plug substantiation test (melt and no-melt) 

would be required (if applicable), unless it can be demonstrated that heat transfer from the brake 

stack to the wheel is identical to the original. Also, antiskid-on and antiskid-off tests (if 

applicable), qualitative taxi tests, and wet runway tests (antiskid-on, if applicable) would be 

required. These test procedures are described in AC 25-7A (formerly Order 8110.8, paragraph 

89(f)). 

(c) It is noted that numerous older airplanes may use these new asbestos-free 

brakes. The intent of the test program discussed above is to validate the brake change, as it 

affects each type design under consideration. It is important to note that each airplane type 

which uses these brakes provides a different operating environment, considering landing gear 

geometry, dynamic response, weight, rate of energy absorption, etc. As stated in AC 25-7A 

(formerly Order 8110.8, paragraph 89(b)), a sufficient number of conditions to verify the existing 

52 
3/14/2000 AC 25-22 

approved performance levels on each type design would be required. The actual content of the 

test program would depend on many factors, including the similarity of the old and new designs, 

the relative performance envelopes of the demonstrated brake and candidate airplane, effect of 

the new brake design on fuse plug performance (if applicable), presence or absence of antiskid 

systems, and possible effects of the change on ground handling. It is also possible that the 

addition of these new brakes to certain existing airplanes could be considered to be a minor 

change, in which case AC 25-7A (formerly Order 8110.8, paragraph 89(b)) would apply, and 

only uninstrumented functional landings would be required. It was recognized that it may be 

impossible to duplicate test data that may have been created over 20 years ago, especially if the 

airplane in question is not equipped with an antiskid system. In this case, it may be appropriate 

to apply a suitable performance penalty to allow for this uncertainty. In any event, engineering 

judgment should be applied to each individual situation so that the intent of the test program

discussed above will be followed. 

(d) While the performance of the new brake must be no less than that of the 

originally approved brake, it also should not be greater; e.g., a higher brake torque is desirable 

for reducing stopping distance but may be undesirable for structural integrity of other 

components of the landing gear. Original landing gear designs are based on structural analysis, 

which could be adversely affected by a brake system change. 

(11) Certification Requirements for Brake Components - Performance. The following 

is extracted from an FAA letter dated October 26, 1987, which addresses "major" and "minor" 

changes to aircraft brake components and explains FAA policy on this subject. 

 (a) Measured accelerate-stop distance, landing distance, and functional flight tests 

may be required depending upon an evaluation of the individual characteristics of each brake 

system change and in order to establish aircraft landing and RTO certification performance 

levels for a replacement or modified brake. The type and magnitude of flight tests required will 

depend on whether or not a requested change to the brake involves a change in heat sink capacity 

or torque requirements of the original certificated brake. In addition, such tests will also depend 

on whether or not an increase in the FAA certificated performance level is desired by the 

applicant, and whether the proposed change to the brake would affect the landing gear static or 

dynamic loads. 

(b) The flight test procedures applicable to brake certification, among others, are 

contained in AC 25-7A, Flight Test Guide for Transport Category Airplanes (formerly FAA 

Order 8110.8). The testing requirements in AC 25-7A (formerly Order 8110.8, paragraph 89(b)), 

relate to the addition of a highly modified brake design to an existing type certificated airplane. 

A highly modified brake is defined in this section as one which contains new or modified parts 

which in turn will cause a significant variance in the kinetic energy absorption characteristics of

the brake. The FAA concurs that a change to either the rotors or stators in a previously approved 

brake could meet this definition, and be considered a major change. In order to adequately 

demonstrate the performance of a major change to the brakes or brake control system, a 

maximum kinetic energy RTO, as described in AC 25-7A (formerly Order 8110.8), would 

normally be required, as well as six functional takeoffs and landings. In addition, fuse plug 

53
3/14/2000 AC 25-22 

substantiation tests (melt and no-melt) would be required (if applicable), unless it can be 

demonstrated that heat transfer from the brake stack to the wheel is identical to the original. 

Also, antiskid-on and antiskid-off tests (if applicable), qualitative taxi tests, and wet runway tests 

(antiskid-on, if applicable) would be required. Order 8110.8 has been canceled. Relevant 

portions of Order 8110.8 and additional information have been included in AC 25-7A, chapter 4, 

section 4, paragraph 55. 

(c) Minor brake changes that cannot affect airplane braking performance may 

require functional landings. This may be required to verify airplane-pilot-brake-antiskid 

combination compatibility. Normally, five non-instrumented functional landings are considered 

sufficient to verify this compatibility. A change to thicker friction material, heavier heat sink 

elements, or other structural improvements generally fall into this minor change category, and 

the effect of the weight change on landing gear performance and other airplane effects would be 

evaluated. If the change to the brake resulted in a lighter weight heat sink, the effects of this 

change would be carefully evaluated to ensure that the thermal characteristics of the brake, 

wheel, axle, fuse plugs, or other landing gear structure is not adversely affected. 

(d) Technical Standard Order (TSO) approval of a modification to a brake is only 

the first step towards approval for the brake to be installed on an airplane. In each case, the 

modification must be shown to have no adverse effect on airplane performance. Depending on 

the magnitude of the proposed change, this demonstration may consist of functional tests only, a 

full test program, or a selection of the various test procedures described in AC 25-7A (formerly 

Order 8110.8). Enough flexibility exists in FAA procedures to allow for a thorough evaluation 

of a proposed brake change without causing undue burden on the applicant. In any event, if 

there is a question whether or not a given modification to a brake will result in a change to the 

certificated performance of the airplane, FAA policy has been to demonstrate the performance 

change on the airplane, or, at the very least, to impose a suitable performance penalty so as to 

cover any eventuality. There are no plans to relax this requirement. 

(12) Approval of Brake Components Utilizing New Brake Friction Material. The 

following is extracted from an FAA memorandum dated July 27, 1987, which addresses 

guidance regarding certification of brake components utilizing new brake friction material. 

(a) This is in response to questions regarding an approval for new replacement 

brake pads. For the purpose of approving a replacement brake pad, the FAA may issue a design 

approval by either a Parts Manufacturer Approval (PMA) or a Technical Standard Order 

Authorization (TSOA) followed by a Supplemental Type Certificate (STC) for installation, or a 

procedure using an STC followed by a PMA for production. Each of these approval methods is 

discussed below. Once a PMA approval is given, the applicant could then mark the brake 

assembly with his own nameplate, indicating the method of approval (STC number), without 

altering the existing TSO nameplate. 

(b) Issuance of a PMA. If the applicant intends to show identicality with existing 

TSO-authorized parts, he must consider not only the form and fit of the parts, but also the 

chemical and/or metallurgical characteristics (including microstructure, surface and cross54 
3/14/2000 AC 25-22 

sectional hardness, and material spectrum analysis), and process and manufacturing 

specifications. Previous experience with various brake manufacturers has shown that it is 

extremely difficult to duplicate the physical and chemical characteristics of brake lining 

materials. For example, two linings with similar hardness or other measurable characteristics 

could have been produced using different process specifications. If the applicant satisfactorily 

shows that his parts are indeed identical to the originals, the change could then be considered to 

be a "minor change," as discussed in AC 25-7A (formerly Order 8110.8, paragraphs 89(b, (c)). 

Five non-instrumented landings would be sufficient to demonstrate compatibility between the 

airplane, pilot, brake, and anti-skid systems on the airplane. With respect to the TSO, a 

statement of conformance and submittal of the data showing identicality would be sufficient. 

Order 8110.8 has been canceled. Relevant portions of Order 8110.8 and additional information 

have been included in AC 25-7A, chapter 4, section 4, paragraph 55. 

(c) Issuance of a TSOA+STC or an STC+PMA. If the applicant is unable to 

demonstrate identicality, the change must then be considered to be a major change, since the 

coefficient of friction and wear rates may be different, and the new parts could, therefore, cause a 

significant variance in the heat sink capacity, the torque characteristics, or the kinetic energy 

absorption characteristics of the brake. It would be necessary, therefore, for the applicant to 

demonstrate compliance with the pertinent parts of TSO-C26c, including the 100 landing stops 

and 1 accelerate-stop described in Table III of § 4.2 of the TSO. 

(d) In addition to the TSO requirements, a maximum kinetic energy Rejected 

Takeoff (RTO), as described in AC 25-7A (formerly Order 8110.8), must be conducted in the 

airplane, as well as six functional takeoffs and landings. In addition, fuse plug substantiation 

tests (melt and no-melt) must be conducted, unless the applicant can demonstrate that heat 

transfer from the brake stack to the wheel is identical to the original. Also, antiskid-on and 

antiskid-off tests, qualitative taxi tests, and wet runway tests (antiskid-on) must be conducted. 

These test procedures are described in AC 25-7A (formerly Order 8110.8, paragraph 89), and are 

required to adequately demonstrate the performance of the new lining material . 

(e) If the applicant successfully demonstrates that the performance of the brake 

with his parts installed is identical to the original, he may mark the brake as described above. If 

the demonstrated performance is greater than that of the original, however, an FAA evaluation 

by the cognizant ACO is essential, since original landing gear designs are based on structural 

analysis which could be adversely affected by a brake system change. 

(13) Maximum Quick Turnaround Times. The following is extracted from an FAA 

letter dated August 19, 1983, which addresses a request to intermix brakes from different 

manufacturers for testing of the fuse plug integrity and receive certification credit for each 

brake/wheel system. 

(a) It is fundamental FAA policy to test the complete configuration of an airplane 

system that is to be certificated for transportation service. In this case the applicant proposed to 

certificate two unmixed configurations, the Original Equipment Manufacturer 1(OEM1) and the 

Original Equipment Manufacturer 2 (OEM2) wheel/brake systems. In general, the FAA does not 

55
3/14/2000 AC 25-22 

certificate unmixed configurations based solely on a test of mixed configurations. Further, 

certification of the mixed configuration requires extensive testing, beyond the test outlined in AC 

25-7A (formerly FAA Order 8110.8). Order 8110.8 has been canceled. However, the relevant 

portions of the Order and additional information have been included in AC 25-7A, chapter 4, 

section 4, paragraph 55. 

(b) Precedent tests previously accepted by the FAA have been cited. These cases 

were judged on their merits, as must be done here. Approval of the eutectic melt temperature 

reduction was based on dynamometer tests and an analysis. In that case, wherein the 

wheel/brake configuration was fixed and test date/analysis were provided, the FAA judged 

equivalency in a single system element (fuse plugs). The FAA did not consider it appropriate to 

compare the equivalency finding in the foregoing case to that of mixing wheel/brake 

configurations for establishing new turnaround limitations. 

(c) The OEM1 and OEM2 wheel/brake configurations are “similar in design and 

performance" was cited as justification to allow mixed wheel/brake fuse plug turnaround testing. 

The FAA noted that the wheel brake designs were intended to mount on the same axle using the 

same hydraulic and antiskid/autobrake systems and brake torque reaction rods. Both wheel/ 

brake assemblies have double heat shields and the tolerance of the OEM1 fuse plug is inside the 

tolerance of the OEM2 fuse plug. The nominal difference is not considered significant. 

(d) The structural arrangement in detail, however, is different in a number of 

places, including in and around the fuse plugs. The FAA must assume the variations in metal 

mass and configuration will affect heat transfer. Also, there are differences in the hydraulic 

actuators. The brakes themselves are made up of proprietary materials that will exhibit different 

thermal reactions. The wheels are similar in appearance because they have to be 

interchangeable; however, it has not been established that they are similar relative to the fuse 

plug performance. 

(e) "Significant margin exists for higher certificated no melt energy levels" is also 

proposed. It is the FAA's understanding that the applicant intends to use this margin, derived 

from dynamometer tests minus airplane test data available, to increase the fuse plug "no melt" 

energy. The FAA has determined this increase is a valid reason to substantiate by test both 

configurations, i.e., OEM1 and OEM2, with the wheels and brakes in the unmixed state. 

(f) The applicant proposed that twice as many fuses be instrumented as in the 

basic certification. This would give useful empirical data, although the FAA does not require 

fuse plug instrumentation. In the final analysis, the success of the test is judged by whether the 

fuse plug integrity is maintained. 

(g) In light of the foregoing, the FAA has determined that each wheel/brake 

configuration should be tested separately for the "no melt" condition to provide a technically 

defensible minimum wheel/brake fuse plug sample size. This decision is consistent with FAA 

policy published in AC 25-7A, Flight Test Guide for Transport Category Airplanes, chapter 4, 

section 4, paragraph 55. This information was in Order 8110.8, paragraphs 89(a), (b), and (c), 

56 
3/14/2000 AC 25-22 

which has been canceled. However, the relevant portions of the Order and additional 

information have been included in AC 25-7A, 

(h) It has been suggested that the proposed intermix and test procedure ensures 

minimal data scatter between brakes. If this statement means that the test data scatter between 

the two tire/wheel configurations will be minimized because all the data will be collected on one 

test run versus two, which may occur at separate times of the day or different days under varying 

conditions, then we can concur. However, this does not lend credence to the selection of an 

inadequate sample size. The proposal does not address what occurs if differences in brake 

performance are observed; or how to establish a significant difference (which could affect 

turnaround times), when comparing one brake system to another. 

 e. References. The addresses for ordering the latest revision of advisory circulars, 

technical standard orders, and other referenced documents listed below can be found in the 

Appendix 4 to this AC. 

TSO-C26, was an acceptable standard to CAA but not part of the regulations. 

TSO-C26a, was part of regulations as § 514.72 of the CAR. 

TSO-C26b, was part of regulations as § 37.172. 

TSO-C26c, was part of regulations as § 37.172 of 14 CFR. 

TSO-C26c with Addendum I, was removed from regulations (§ 37.172, Amendment

25-52, 1980) and became a voluntary standard. 

AC 25-7A, Flight Test Guide for Transport Category Airplanes (formerly FAA 

Order 8110.8). 

AC 43.13-1A, Acceptable Methods, Techniques and Practices- Aircraft Inspection and 

Repair; L. G. Equipment.

FAA Letter, Dynamometer Test Guidelines to Establish Brake wear Limits. 

FAA's Wheels and Brakes Approval Process Videos, parts 1 and 2. 

SAE AS-1145A - Aircraft Brake Temperature Monitor System. 

SAE ARP-1907 - Automatic Braking Systems Requirements. 

SAE AIR-1064C - Brake Dynamics. 

SAE ARP-1070B - Design and Testing of Antiskid Brake Control. 

SAE AIR-1739 - Information on Antiskid Systems. 

SAE ARP-813B - Maintainability Recommendations for Aircraft Wheels and Brakes. 

SAE ARP-1619 - Replacement and Modified Brakes and Wheels. 

SAE AS-483A - Skid Control Equipment. 

SAE ARP-862A - Skid Control Performance. 

SAE AIR-1934A - Use of Carbon Heat Sink Brakes on Aircraft. 

SAE ARP-597C - Wheels and Brakes, Supplementary Criteria for Design Endurance. 

57
3/14/2000 AC 25-22 

THIS PAGE INTENTIONALLY LEFT BLANK. 

58 
3/14/2000 AC 25-22 

Section 3. PERSONNEL AND CARGO ACCOMMODATIONS 

15. SECTION 25.773 - PILOT COMPARTMENT VIEW.

 a. Rule Text. 

(a) Nonprecipitation conditions. For nonprecipitation conditions, the following 

apply: 

(1) Each pilot compartment must be arranged to give the pilots a sufficiently 

extensive, clear, and undistorted view, to enable them to safely perform any 

maneuvers within the operating limitations of the airplane, including taxiing, 

takeoff, approach, and landing. 

(2) Each pilot compartment must be free of glare and reflection that could 

interfere with the normal duties of the minimum flight crew (established under 

§ 25.1523). This must be shown in day and night flight tests under 

nonprecipitation conditions. 

(b) Precipitation conditions. For precipitation conditions, the following apply: 

(1) The airplane must have a means to maintain a clear portion of the windshield, 

during precipitation conditions, sufficient for both pilots to have a sufficiently 

extensive view along the flight path in normal flight attitudes of the airplane. This 

means must be designed to function, without continuous attention on the part of 

the crew, in −

(i) Heavy rain at speeds up to 1.6 VS1 with lift and drag devices retracted; and 

(ii) The icing conditions specified in § 25.1419 if certification with ice protection 

provisions is requested. 

(2) The first pilot must have −

(i) A window that is openable under the conditions prescribed in paragraph (b)(1)

of this section when the cabin is not pressurized, provides the view specified in 

that paragraph, and gives sufficient protection from the elements against 

impairment of the pilot's vision; or 

(ii) An alternate means to maintain a clear view under the conditions specified in 

paragraph (b)(1) of this section, considering the probable damage due to a severe 

hail encounter. 

(c) Internal windshield and window fogging. The airplane must have a means to 

prevent fogging of the internal portions of the windshield and window panels over 

an area which would provide the visibility specified in paragraph (a) of this 

section under all internal and external ambient conditions, including precipitation 

conditions, in which the airplane is intended to be operated. 

(d) Fixed markers or other guides must be installed at each pilot station to enable 

the pilots to position themselves in their seats for an optimum combination of 

59
3/14/2000 AC 25-22 

outside visibility and instrument scan. If lighted markers or guides are used they 

must comply with the requirements specified in § 25.1381. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, Apr. 8, 

1970; Amdt. 25-46, Oct. 30, 1978; Amdt. 25-72, Jul. 20, 1990]

b. Intent of rule. The intent of this rule is for the flight deck windshield to provide 

sufficient external vision to permit the pilot to safely perform any maneuvers within the 

operating limits of the aircraft and, at the same time afford an unobstructed view of the flight 

instruments and other critical components and displays from the same eye position. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Section 4b.351 of 

the CAR became § 25.773 of 14 CFR for pilot compartment view. 

 (1) Amendment 25-23 (April 8, 1970) added a new § 25.773(c) to require means to 

prevent fogging of the internal portions of the windshield and window panels which can occur 

under certain temperature and humidity conditions and create a hazardous situation. 

 (2) Amendment 25-46 (October 30, 1978) added a new § 25.773(d) requiring means 

(such as fixed markers) to guide the pilot to determine an optimum seat position for proper 

visibility. 

 (3) Amendment 25-72 (July 20, 1990) revised § 25.773(b)(1)(i) for clarity and 

consistency with actual certification practice by specifying that all lift and drag devices, e.g., 

slats and spoilers as well as flaps, must be retracted. In addition, section 25.773(b)(2) was 

revised to allow alternate means of compliance to maintain clear vision as opposed to an 

openable sliding window being the only means. 

 (4) Harmonization. Parts of this regulation (§ 25.773(b)(2) and (b)(4)) will be the 

subject of a Federal Aviation Regulations/Joint Aviation Requirements (FAR/JAR) 

harmonization effort under the Aviation Rulemaking Advisory Committee (ARAC). The ARAC 

harmonization working group may adopt parts of both FAR and JAR as well as revise 

AC 25.783-1. 

 d. Policy/Compliance Methods. For guidance with this requirement, refer to the preamble 

of this rule and Advisory Circular (AC) 25-7A, Flight Test Guide for Certification of Transport 

Category Airplanes, and AC 25.773-1, Pilot Compartment View for Transport Category 

Airplanes. 

 e. References. None. 

16 - 25. [RESERVED] 

60 
3/14/2000 AC 25-22 

26. SECTION 25.783 - DOORS.

 a. Rule Text. 

(a) Each cabin must have at least one easily accessible external door. 

(b) There must be a means to lock and safeguard each external door against 

opening in flight (either inadvertently by persons or as a result of mechanical 

failure or failure of a single structural element either during or after closure). 

Each external door must be openable from both the inside and the outside, even 

though persons may be crowded against the door on the inside of the airplane. 

Inward opening doors may be used if there are means to prevent occupants from 

crowding against the door to an extent that would interfere with the opening of 

the door. The means of opening must be simple and obvious and must be 

arranged and marked so that it can be readily located and operated, even in 

darkness. Auxiliary locking devices may be used. 

(c) Each external door must be reasonably free from jamming as a result of 

fuselage deformation in a minor crash. 

(d) Each external door must be located where persons using them will not be 

endangered by the propellers when appropriate operating procedures are used. 

(e) There must be a provision for direct visual inspection of the locking 

mechanism to determine if external doors, for which the initial opening movement 

is not inward (including passenger, crew, service, and cargo doors), are fully 

closed and locked. The provision must be discernible under operational lighting 

conditions by appropriate crewmembers using a flashlight or equivalent lighting 

source. In addition, there must be a visual warning means to signal the 

appropriate flight crewmembers if any external door is not fully closed and 

locked. The means must be designed such that any failure or combination of 

failures that would result in an erroneous closed and locked indication is 

improbable for doors for which the initial opening movement is not inward. 

(f) External doors must have provisions to prevent the initiation of pressurization 

of the airplane to an unsafe level if the door is not fully closed and locked. In 

addition, it must be shown by safety analysis that inadvertent opening is extremely 

improbable. 

(g) Cargo and service doors not suitable for use as emergency exits need only 

meet paragraphs (e) and (f) of this section and be safeguarded against opening in 

flight as a result of mechanical failure or failure of a single structural element. 

(h) Each passenger entry door in the side of the fuselage must meet the applicable 

requirements of §§ 25.807 through 25.813 for a Type II or larger passenger 

emergency exit. 

(i) If an integral stair is installed in a passenger entry door that is qualified as a 

passenger emergency exit, the stair must be designed so that under the following 

conditions the effectiveness of passenger emergency egress will not be impaired: 

(1) The door, integral stair, and operating mechanism have been subjected to the 

inertia forces specified in 25.561(b)(3), acting separately relative to the 

surrounding structure. 

61
3/14/2000 AC 25-22 

(2) The airplane is in the normal ground attitude and in each of the attitudes 

corresponding to collapse of one or more legs of the landing gear. 

(j) All lavatory doors must be designed to preclude anyone from becoming 

trapped inside the lavatory, and if a locking mechanism is installed, it be capable 

of being unlocked from the outside without the aid of special tools. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-15, 

32 FR 13262, Sep. 20, 1967; Amdt. 23-23, 35 FR 5676, Apr. 8, 1970; Amdt. 25-

54, 45 FR 60173, Sep. 11, 1980; Amdt. 25-72, 55 FR 29780, Jul. 20, 1990; Amdt. 

25-88, 61 FR 57956, Nov. 8, 1996]

b. Intent of Rule. The intent of the term "reasonably free" is to provide proper dimensional 

clearances around door structure and seals. The requirement is also defined under § 25.809(g): 

"There must be provisions to minimize the probability of jamming of the emergency exits 

resulting from fuselage deformation in a minor crash landing." 

 c. Background. This regulation is the subject of a Federal Aviation Regulations/Joint 

Aviation Requirements (FAR/JAR) harmonization effort under the Aviation Rulemaking 

Advisory Committee (ARAC). The ARAC working group is tasked to recommend revisions to 

§ 25.783 and related crashworthiness regulations and to recommend revisions to AC 25.783-1A. 

d. Policy/ComplianceMethods. For policy and guidance on compliance with this 

requirement, refer to Advisory Circular (AC) 25-783-1, Fuselage Doors, Hatches, and Exits; 

AC-25-21, Certification of Transport Airplane Structure, paragraph 85; and AC 25-17A, 

Transport Airplane Cabin Interiors Crashworthiness Handbook. 

 e. Reference. None. 

27 - 29. [RESERVED] 

62 
3/14/2000 AC 25-22 

Section 4. VENTILATION AND HEATING 

36. SECTION 25.831 - VENTILATION.

a. Rule Text. 

(a) Under normal operating conditions and in the event of any probable failure 

conditions of any system which would adversely affect the ventilating air, the ventilation 

system must be designed to provide a sufficient amount of uncontaminated air to enable 

the crewmembers to perform their duties without undue discomfort or fatigue and to 

provide reasonable passenger comfort. For normal operating conditions, the ventilation 

system must be designed to provide each occupant with an airflow containing at least 

0.55 pounds of fresh air per minute. 

(b) Crew and passenger compartment air must be free from harmful or hazardous 

concentrations of gases or vapors. In meeting this requirement, the following apply: 

(1) Carbon monoxide concentrations in excess of 1 part in 20,000 parts of air are 

considered hazardous. For test purposes, any acceptable carbon monoxide detection 

method may be used. 

(2) Carbon dioxide concentration during flight must be shown not to exceed 0.5 percent 

by volume (sea level equivalent) in compartments normally occupied by passengers or 

crewmembers. 

(c) There must be provisions made to ensure that the conditions prescribed in paragraph 

(b) of this section are met after reasonably probable failures or malfunctioning of the 

ventilating, heating, pressurization, or other systems and equipment. 

(d) If accumulation of hazardous quantities of smoke in the cockpit area is reasonably 

probable, smoke evacuation must be readily accomplished, starting with full 

pressurization and without depressurizing beyond safe limits. 

(e) Except as provided in paragraph (f) of this section, means must be provided to enable 

the occupants of the following compartments and areas to control the temperature and 

quantity of ventilating air supplied to their compartment or area independently of the 

temperature and quantity of air supplied to other compartments and areas: 

(1) The flight crew compartment. 

(2) Crewmember compartments and areas other than the flight crew compartment unless 

the crewmember compartment or area is ventilated by air interchange with other 

compartments or areas under all operating conditions. 

(f) Means to enable the flight crew to control the temperature and quantity of ventilating 

air supplied to the flight crew compartment independently of the temperature and 

quantity of ventilating air supplied to other compartments are not required if all of the 

following conditions are met: 

(1) The total volume of the flight crew and passenger compartments is 800 cubic feet or 

less. 

63
3/14/2000 AC 25-22 

(2) The air inlets and passages for air to flow between flight crew and passenger 

compartments are arranged to provide compartment temperatures within 5 degrees F. of 

each other and adequate ventilation to occupants in both compartments. 

(3) The temperature and ventilation controls are accessible to the flight crew. 

(g) The exposure time at any given temperature must not exceed the values shown in the 

following graph after any improbable failure condition. 

temperature curve from Amdt. 25-87

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 

42 FR 36970, Jul. 18, 1977; Amdt. 25-87, 61 FR 28695, Jun. 5, 1996; Amdt. 25-

89, 61 FR 63956, Dec. 2, 1996]

b. Intent of Rule. The intent of this rule is to supply passengers and crewmembers with 

enough uncontaminated air to provide reasonable comfort during normal operating conditions 

and also after any probable failure of any system that would adversely affect the cockpit or cabin 

ventilation air. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The majority of the 

requirements from section 4b.371 of the CAR were carried over directly to § 25.831 of 14 CFR 

for ventilation. 

 (1) Amendment 25-41 (July 18, 1977) revised the rule to clarify the requirements for 

the means by which crewmembers can control the temperature and quantity of ventilating air in 

the flight crew compartment and compartments occupied by other crew members. 

 (2) Amendment 25-87 (June 5, 1996) revised the rule to codify special conditions that 

had been used to certify certain airplanes for high altitude operation (maximum altitudes up to 

64 
3/14/2000 AC 25-22 

45,000 or 51,000 feet). The amendment added the requirement that each occupant be provided 

with 0.55 pounds (equivalent to 10 cubic feet) per minute of fresh air, reference § 25.831(a). 

The amendment also specifies limits for interior temperatures after specific failures, reference 

§ 25.831(g). Compliance to Amendment 25-87 is required for only those airplanes whose 

certification basis includes this amendment. Most of the airplanes that are currently produced do 

not fall into this category. An accompanying advisory circular, AC 25-20, Pressurization, 

Ventilation, and Oxygen Systems Assessment for Subsonic Flight including High Altitude 

Operation, was also issued to provide one means, but not the only means, of compliance to the 

regulations affected by Amendment 25-87. 

 (3) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group will be tasked to 

recommend revisions to §§ 25.831, 25.841, and related regulations and to recommend revisions 

to AC 25.20. 

 (4) Amendment 25-89 (December 2, 1996) revised § 25.831(b)(2) to change the carbon 

dioxide requirements from three percent by volume to 0.5 percent by volume (sea level 

equivalent). The following information is extracted from an FAA letter dated May 18, 1999, to 

the Society of Automotive Engineers (SAE) Committee AC-9, Aircraft Environmental Systems, 

to acknowledge an error in the preamble to Amendment 25-89, page P-610, 3rd paragraph. It 

states that ". . . sea level equivalent (SLE) for a cabin altitude of 8,000 feet would be 0.5 percent 

multiplied by 0.74 . . ." The word "multiplied" should have read "divided." The FAA does not 

feel that an increase in carbon dioxide concentration from 0.5 percent SLE to 0.68 percent at 

altitude will introduce an appreciable discomfort or hazard to passengers and maintains a 

concentration level well below industry recommendations for carbon dioxide exposure. 

 d. Policy/Compliance Method. For guidance on compliance with this requirement, refer to 

the preambles of this rule and the following information. 

 (1) Ventilation rates are determined by tests and/or analyses. For compliance with 

§ 25.831(b) and (c), carbon monoxide and carbon dioxide levels may be measured with hand 

held samplers or gas measurement devices for the various flight conditions, such as takeoff, 

cruise, and landing. Measurements are made at "head" level in the cockpit, other compartments 

occupied by crewmembers, and in the passenger cabin. 

 (2) Smoke Evacuation. For compliance with § 25.831(d), cockpit smoke evacuation 

test methods are addressed in AC 25-9A, Smoke Detection, Penetration, and Evacuation Tests 

and Related Flight Manual Emergency Procedures. The AC 25-9A also provides guidance 

regarding lavatories, galleys, crew rest areas, equipment cooling bays, and cargo compartments 

smoke detection/penetration/evacuation. 

 (3) Ventilation System Design. The ventilation system design should be reviewed and 

tests or analyses conducted to ensure compliance with § 25.831(e) and (f), which address control 

of temperatures and ventilating air quantity in compartments occupied by crewmembers. 

65
3/14/2000 AC 25-22 

 (4) Airplane Operation with Airconditioning Packs Off. The following policy is 

extracted from an FAA memorandum dated September 3, 1999, in response to questions 

regarding compliance with § 25.831 while operating with the airconditioning packs off. 

(a) Background. The following summarizes some key elements that are necessary 

to understand in determining the certification policy regarding cabin ventilation, smoke detection 

and evacuation, and equipment cooling. 

(b) The provisions of § 25.831(a), as amended by Amendment 25-41, state: “Each 

passenger and crew compartment must be ventilated, and each crew compartment must have 

enough fresh air (but not less than 10 cu. ft. per minute per crewmember) to enable 

crewmembers to perform their duties without undue discomfort or fatigue”. 

(c) A 1983 memo addressing airconditioning packs-off operation provides 

additional guidance to the regulatory provisions of § 25.831(a), as amended by Amendment 

25-41. 

 1 The memo states: “The proposed environmental control system (ECS) 

takeoff procedure for the Model ABC should be processed for certification by an exemption to 

25.831(a).” 

 2 The memo also states: “Analysis and tests have been used to substantiate 

that operating without the required 10 cubic feet per minute per pilot (§ 25.831(a)) for short 

duration will not impair pilot performance or significantly affect equipment reliability. This is 

not to say that the pilots need be comfortable. These ECS ‘off’ approvals were made in 

accordance with adequate criteria, but not all were made in accordance with correct certification 

procedures (i.e., equivalency or exemption).” 

(d) Subsequently, Amendment 25-87, effective April 30, 1998, amended several of 

the airworthiness provisions concerning cabin ventilation. Section 25.831(a) was changed to 

require that “. . . the ventilation system must be designed to provide a sufficient amount of 

uncontaminated air to enable the crewmembers to perform their duties without undue discomfort 

or fatigue and to provide reasonable passenger comfort. For normal operating conditions, the 

ventilation system must be designed to provide each occupant with an airflow containing at least 

0.55 pounds of fresh air per minute.” 

 1 The preamble to Amendment 25-87 adds the following additional 

information. “One commenter recommends allowing the fresh air requirements proposed to be 

required under § 25.831(a) to remain a crewmember requirement only. The FAA does not 

concur with this recommendation. It has been determined that this level of airflow is required 

for several reasons. Members of the flightcrew performing their functions in the passenger cabin 

are not sedentary and must perform their duties without undue discomfort or fatigue. In addition, 

fresh airflow has been determined to be necessary to provide adequate smoke clearance in the 

event of smoke accumulation due to a system failure or fire. However, it is clear that the 

66 
3/14/2000 AC 25-22 

additional airflow is not required at all times and under all operating conditions. Therefore, the 

wording in the final rule has been changed to state that the ventilation system must be designed 

to provide the fresh airflow. This also addresses concerns regarding the low fresh airflow 

capability that occurs during descent at low power levels.” 

(e) Advisory Circular 25-20 was developed concurrent with Amendment 25-87. 

Paragraph 5d states: “Takeoff with airconditioning or bleed air system ‘off’ may be an 

acceptable procedure provided the ventilation system continues to provide an acceptable 

environment in the passenger cabin and cockpit for the brief period when the ventilation system

is not operating normally.” 

(f) The Draft Version of this AC (dated November 30, 1998) was issued for public 

comment as well as coordinated for directorate comments. This AC includes a statement that 

recommends that the exemption process be used to approve the cabin ventilation system for 

operation with airconditioning packs-off operation. However, the Aerospace Industries 

Association (AIA) has questioned the need for an exemption and stated that the no packs takeoff 

procedures already exist; that as the transient for this condition is short, this is an accepted 

practice; and that it does not cause the cabin environment to be unsafe. In addition, the Los 

Angeles Aircraft Certification Office (ACO) commented that an equivalent safety finding was 

more appropriate than an exemption. 

(g) Analysis. The Transport Standards Staff’s (TSS) interpretation of the 

provisions of § 25.831(a), as amended by Amendment 25-41, is that the prescribed airflow for 

the flightcrew is required to be provided during normal operation. However, the certification 

records do not show that any such equivalent safety finding or exemption for airconditioning 

packs-off operation (i.e., no fresh air for crewmembers) was issued for the model ABC or for any 

other transport category airplane. Therefore, the TSS assumes that any transport category 

airplane for which the AFM allows packs-off operation, does not strictly comply with the 

provision of § 25.831(a), as amended by Amendment 25-41. 

 1 Recognizing that the cabin ventilation provisions of Amendment 25-41 were 

overly restrictive, the FAA, in Amendment 25-87, changed the requirement of § 25.831 so that 

the regulatory provision now reads: “ . . . the ventilation system must be designed to provide a 

sufficient amount of uncontaminated air to enable the crewmembers to perform their duties 

without undue discomfort or fatigue . . .” 

 2 The preamble discussion quoted above states that the reason for this change 

is that “. . . it is clear that the additional airflow is not required at all times and under all 

operating conditions.” 

 3 Therefore, provisions of § 25.831, as amended by Amendment 25-87, 

specify a design requirement and allow for limited interruptions of the specified design airflow. 

However, this interruption may not result in “undue” discomfort or fatigue to the crewmembers. 

Advisory Circular (AC) 25-20, which was developed concurrently with Amendment 25-87, 

supports this interpretation (see excerpt above). 

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3/14/2000 AC 25-22 

 4 Furthermore, an airplane accident heightened FAA awareness to the need to 

address smoke penetration and include cargo compartment fire protection (detection and 

suppression) throughout the flight, including taxi, takeoff and climb. When determining 

compliance with the sections of part 25 relating to flight crew compartment air quality, 

ventilation, smoke penetration/evacuation, cargo compartment fire protection (detection and 

suppression) and equipment cooling, all the impacts related to operating with the airconditioning 

packs off for a limited time must be determined. The TSS considers the failure to document any 

consideration of these issues in previous compliance findings an oversight by the FAA. 

 5 While direct compliance to § 25.831(a), as amended by Amendment 25-87, 

for airconditioning packs-off operation is possible without using an equivalent safety finding; an 

equivalent safety finding (§ 21.21(b)(1)) may be used for showing compliance with the 

provisions of § 25.831(a), as amended by Amendment 25-41, for airconditioning packs-off 

operation. 

(h) Conclusion. The TSS has determined that while direct compliance to 

§ 25.831(a), as amended by Amendment 25-87, for airconditioning packs-off operation is 

possible without using an equivalent safety finding, an equivalent safety finding (§ 21.21(b)(1)) 

may be used for showing similar compliance with the provisions of § 25.831(a), as amended by 

Amendment 25-41. The equivalent safety finding must document that the ventilation system

continues to provide an acceptable environment in the passenger cabin and cockpit for the brief 

period when the ventilation system is not operating normally. The degradation of crewmember 

air quality must not reach the level that would cause undue discomfort and fatigue to the point 

that it could affect the performance of their duties. 

(i) Advisory Circular (AC) 25.1581-1, Airplane Flight Manual, issued in 1997, 

states in paragraph 2b(11), Systems and Equipment Limitations, that; “all limitations applicable 

to systems and equipment installations that are considered necessary for safe operation must be 

included.” Airconditioning packs-off operation is intended to be a short duration operation. 

Therefore, the maximum period of operation in this configuration should be defined by the 

applicant and specified in the AFM, along with any related operating procedures necessary to 

maintain compliance with the regulatory issues discussed above. An example of establishing 

"the maximum period of operation (short duration) for takeoff" would be an operational phase 

beginning with turning packs off when cleared into position for takeoff, and ending when packs 

were turned back on after takeoff thrust was reduced to climb thrust or when accomplishing the 

"after-takeoff" checklist.

(j) Packs-On and Packs-Off. In accordance with §§ 25.855 and 25.857, there must 

be a means to exclude hazardous quantities of smoke, flames, or extinguishing agent from any 

compartment occupied by crew or passengers. The FAA has historically found compliance to 

these requirements for packs-on operation but has not recognized the need to assess compliance 

with requirements for limited duration packs-off operation. A change in precedent set by 

historical practice is beyond the scope of this informational memorandum. The Transport 

Standards Staff (TSS) recognizes that service experience to date has not demonstrated that 

68 
3/14/2000 AC 25-22 

packs-off operations have resulted in hazardous quantities of smoke. Therefore, at this time, 

determination that packs-off operation is acceptable can be predicated on an FAA finding that no 

unsafe condition due to limited duration packs-off operation would result should a cargo 

compartment fire occur. Criteria that should be considered include: (1) packs-on operation will 

not allow any smoke from the cargo compartment to penetrate the passenger compartment; and 

(2) during limited duration packs-off operations the cargo compartment smoke detection system

is effective and the aircondition packs can be turned on and returned to the approved packs-on 

configuration to exclude hazardous quantities of smoke. The local ACO may make a 

determination of compliance to the aforementioned criteria based upon available test data, 

analysis and extrapolation of these data, as has been historically accepted to demonstrate 

compliance to these and other sections of part 25. 

(5) Fresh Air and Normal Operating Conditions. The following policy is extracted from

an FAA memorandum dated September 10, 1997, in response to questions regarding the terms 

“fresh air” and “normal operating conditions” as it applies to compliance with § 25.831. 

(a) The term “fresh air” was first documented with the ventilation requirements in 

§ 4b.371 of the CAR. Section 4b.371(a) of the CAR required that all crew compartments be 

ventilated with a sufficient amount of fresh air. It was specified that “an outside air supply of 

approximately 10 cubic feet per minute is considered a minimum for each crewmember.” This 

requirement was later codified into 14 CFR § 25.831(a). 

(b) Notice of Proposed Rulemaking (NPRM) 89-31 first proposed that § 25.831(a) 

be amended to include additional minimum airflow requirements for both crew and passengers. 

NPRM 89-31 and Amendment 25-87 proposed that the minimum fresh airflow of 10 cubic feet 

per minute per crewmember be provided to each occupant during normal operation. Both the 

“Discussion” section of NPRM 89-31 and the “Background” section of Amendment 25-87 state 

that some airplanes “incorporate ventilation systems in which fresh air is augmented with 

conditioned and recirculated air.” It was specified that the rule change would permit a 

ventilation system that uses a mixture of the minimum amount of fresh air and any desired 

quantity of recirculated air that is shown to be uncontaminated by odors, particulates, or gases. It 

is stated that the “minimum amount of fresh air is specified by weight. . . .” Ten cubic feet of 

standard air at a typical cabin altitude of 8,000 feet weighs approximately 0.55 pounds, which is 

how the final rule was stated. This provides a clear distinction that the minimum amount of fresh 

air (outside air) is 0.55 pounds per minute per person, but may be augmented by additional 

conditioned and recirculated air. 

(c) Requests have been made that “normal operating conditions” as used in 

§ 25.831(a) be interpreted to mean normal cruise conditions with this interpretation excluding 

ground, climb, descent, and holding operations. The FAA does not agree. Ground, climb, 

descent, and holding operations are considered normal operating conditions. However, it should 

be noted that the wording in the preamble to the final rule stated that for normal operating 

conditions, the ventilation system must be designed to provide the fresh airflow. This was a 

change from the NPRM that which stated that each occupant must be provided the fresh airflow. 

The preamble to Amendment 25-87 states that it is clear that the additional airflow is not 

69
3/14/2000 AC 25-22 

required at all times and under all operating conditions. The change in wording reflects the need 

to express certification rules in terms of design requirements rather than operating restrictions 

and addresses concerns regarding potential low airflow capability that occurs during descent at 

low power levels. If an applicant proposes not to provide the minimum required fresh airflow 

during phases of flight using low power levels, the applicant must show that the cabin air quality 

is not compromised for that period of time. 

(6) Ventilation During Takeoff with the Environmental Control System Turned Off. 

The following policy is extracted from an FAA memorandum dated May 20, 1983, in response 

to questions regarding compliance with § 25.831 while conducting takeoff with the 

environmental control system (ECS) off. 

(a) Proposals for takeoff procedures with the ECS off should be processed for 

certification by an exemption to § 25.831(a). [See paragraph (h) Conclusion, above for a 1999 

update] Transient ventilation conditions have been required to be analyzed for current transport 

aircraft in the past. Analysis and tests should be used to substantiate that operating without the 

required 10 cubic feet per minute per occupant (§ 25.831(a)) for short durations will not impair 

pilot performance, passenger safety, or significantly affect equipment reliability. 

(b) Both equipment and the pilot environment should be evaluated during taxi and 

takeoff to ensure pilot performance and equipment reliability are not impaired. This evaluation 

should cover the extremes of ambient hot air temperatures in which the airplane is expected to 

operate. The temperatures evaluated should be no less than the maximum ambient hot air 

temperature limits of engine or equipment, whichever is more limiting. Analysis may be used to 

extrapolate test temperatures. 

 e. References. The address for ordering the latest revision of the advisory circulars listed 

below can be found in the Appendix to this AC.

AC 25-9A, Smoke Detection, Penetration, and Evacuation Tests and Related Flight 

Manual Procedures. 

AC 25-20, Pressurization, Ventilation, and Oxygen Systems Assessment for Subsonic 

Flight Including High Altitude Operation.

AC 20-32B, Carbon Monoxide (CO) Contamination in Aircraft-Detection and 

Prevention. 

AC 25-7A, Flight Test Guide for Certification of Transport Category Airplanes. 

31. SECTION 25.832 - CABIN OZONE CONCENTRATION.

 a. Rule Text.

(a) The airplane cabin ozone concentration during flight must be shown not to 

exceed70 
3/14/2000 AC 25-22 

(1) 0.25 parts per million by volume, sea level equivalent, at any time above flight 

level 320; and 

(2) 0.10 parts per million by volume, sea level equivalent, time-weighted average 

during any 3-hour interval above flight level 270. 

(b) For the purpose of this section, ''sea level equivalent'' refers to conditions of 

25o C and 760 millimeters of mercury pressure. 

(c) Compliance with this section must be shown by analysis or tests based on 

airplane operational procedures and performance limitations, that demonstrate 

that either- 

(1) The airplane cannot be operated at an altitude which would result in cabin 

ozone concentrations exceeding the limits prescribed by paragraph (a) of this 

section; or 

(2) The airplane ventilation system, including any ozone control equipment, will 

maintain cabin ozone concentrations at or below the limits prescribed by 

paragraph (a) of this section. 

[Amdt. 25-50, 45 FR 3883, Jan. 1, 1980, as amended by Amdt. 25-56, 47 FR 

58489, Dec. 30, 1982]

b. Intent of Rule. Ozone is an odorless, colorless gas that can be irritating to the 

respiratory tract and eyes when present in high enough concentrations. This rule was formulated 

to limit the ozone concentrations in the occupied areas of transport category airplanes. Because 

the level of discomfort is proportional to the level of activity of the parties exposed, cabin 

attendants are more likely to be adversely affected. The ozone limits embodied in this section 

will prevent crew members and passengers from being exposed to concentrations high enough to 

be hazardous. There is a parallel requirement in the operating rules (§ 121.578). 

 c. Background. Following complaints from crewmembers regarding various adverse 

health effects associated with ozone in the airplane cabins, and a petition for rulemaking, parts 

25 and 121 were amended at Amendments 25-50 (January 1, 1980) and 121-154 (January 1, 

1980), to add standards for maximum ozone concentration in transport category airplanes. 

 (1) Amendment 25-56 (December 30, 1982) raised the flight levels at which 

compliance must be demonstrated, and excluded certain airplanes from the requirements of the 

part 121 rule. 

(2) The Operating Rule, § 121.578, allows route and altitude adjustments to be used, in 

conjunction with global ozone concentration data and airplane ozone retention factors, to avoid 

exceeding safe ozone levels. Part 25 requirements were intended to be addressed through design 

features, e.g., ozone filters or catalytic converters, to ensure that ozone limits would not be 

exceeded regardless of altitude, latitude, and time of year for the operation. However, some

manufacturers have requested that they be allowed to comply with § 25.832 by placing 

limitations in the airplane flight manual that would stipulate altitude limits for operations at 

specific latitudes during certain times of the year. 

71
3/14/2000 AC 25-22 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Sea Level Equivalent. The following guidance is extracted from an FAA letter 

dated May 18, 1999. It was developed in response to inquiries regarding a definition of the term

“sea level equivalent” and to acknowledge an error in the preamble to Amendment 25-89 with 

reference to AC 120-38. The following guidance will demonstrate that for cabin air 

concentration the FAA intended to maintain constant parital pressures; however, the 

concentrations may change with altitude. 

(a) Confusion still exists pertaining to the term “sea-level equivalent.” This term

has been used as far back as 1953 in the Civil Air Regulations (CAR) part 4b, § 4b.371, 

Ventilation, to address carbon dioxide concentrations. This later became Title 14, Code of 

Federal Regulations (14 CFR), part 25, § 25.831(b)(2). Amendments 121-154 (Operations) and 

25-50 (Certification), effective February 20, 1980, introduced §§ 121.578 and 25.832, Cabin 

Ozone Concentration. These regulations defined the term sea level equivalent in §§ 

121.578(a)(2) and 25.832(b) as follows: “Sea level equivalent (SLE) refers to conditions of 250

C and 760 millimeters of mercury pressure." The preamble to these amendments further states 

that ozone concentration is stated in parts per million by volume (ppmv) and expressed as a sea 

level equivalent, i.e., the ratio of ozone to air that would exist at 760 millimeters of mercury (mm

Hg) pressure and 250

 C. 

(b) The NPRM that originally proposed ozone limitations, (NPRM No. 78-15), 

introduced an ozone limit of 0.3 ppmv at any given time. During the comment period, confusion 

was expressed as to what was meant by the proposed cabin ozone concentrations. The ratios 

proposed were those that would be expected at the air pressure which is normally maintained in 

the passenger cabins of the affected aircraft (it was considered to be 6,000 feet). This original 

proposal was modified, however, since concentration limits are typically expressed at standard 

sea level pressure (760 mm Hg) and temperature (250

 C), including those adopted by the 

Occupational Safety and Health Administration (OSHA). 

(c) The FAA adjusted the final rule, Amendment. Nos. 121-154 and 25-50, to be 

expressed based on the sea level criteria. The preamble states, “Since there is more air in a given 

volume at sea level, the proposed 0.3 ppmv limit converts to 0.25 as a ppmv, sea level 

equivalent.” The intent was not to allow a higher ozone concentration at altitude, but to take the 

proposed cabin ozone concentration and express it in standardized terms based on published 

OSHA limits. 

 (d) Advisory Circular (AC) 120-38, provides illustrations for determining a sea 

level equivalent concentration. Per the AC, the terms OZSLE and O3 are concentrations in 

ppmv. OZSLE is the sea level equivalent absolute ozone concentration and O3 is the absolute 

concentration at a given altitude (assuming a cabin temperature of 250

 C). The AC provides the 

following equation to relate O3 and OZSLE based on a pressure ratio (P/Po), which will always 

be less than zero for altitudes higher than sea level. 

72 
3/14/2000 AC 25-22 

 1 OZSLE = (P/Po)(O3) 

 P/Po is the ratio of the total cabin pressure (P) to total sea level pressure (Po). This equation 

is nothing more than a conversion of a concentration value at altitude back to sea level. 

(e) Dalton’s Law states that the total pressure of a gas mixture is the sum of the 

individual pressure (or partial pressures) that each gas would exert if it alone occupied the whole 

volume. Each kind of gas exerts its own pressure depending on the percentage of that gas in the 

mixture; thus, even though the percentage or concentration of oxygen in the atmosphere is nearly 

constant (around 20.9%), its partial pressure will decrease proportionately as atmospheric 

pressure decreases. If c is the concentration of a gas, P is the absolute pressure, and Pp is the 

partial pressure, they are related by the following equation. 

 2 c = (Pp/P) 

Advisory Circular (AC) 120-38 uses SLE to essentially provide a constant partial pressure 

requirement independent from the cabin pressure. Substituting the concentrations of ozone at sea 

level (OZSLE) and of ozone under cabin pressure (O3) into equation (1) you get: 

 3 Pp,OZSLE/P = (P/Po)(Pp,O3/P) 

At sea level, P = Po, therefore, equation 3 becomes 

 4 Pp,OZSLE/ Po = (P/Po)(Pp,O3/P) 

which reduces to Pp,OZSLE = Pp,O3. 

(f) This exercise demonstrates that the FAA intended to maintain constant partial 

pressures; however, the concentrations may change with altitude. It was not the intent of the 

FAA to allow higher partial pressures of a gas at low cabin pressures. In fact, the partial 

pressures are intended to be independent of the cabin altitude. With this in mind, the sea level 

equivalent of a gas can be considered the gas concentration (volume %) at sea level for which the 

partial pressure of the gas matches the ambient partial pressure of the gas at a given cabin 

altitude. 

(g) Based on this interpretation of the preamble to Amendment. Nos. 121-154 and 

25-50, and AC 120-38 material, the preamble to Amendment. No. 25-89 should be corrected to 

read as shown in paragraph 36(c)(4)(a) above. 

 (2) Ground test. Applicants have demonstrated compliance with this rule by 

conducting a ground test to establish the ozone retention of the aircraft. This should be 

accomplished by inducing a known ozone concentration into the engine inlet/bleed air system, 

representative of the worst case concentration predicted for the atmospheric environment, and 

measuring the concentration in the passenger cabin and cockpit. Alternatively, ground tests 

could be conducted utilizing ambient ozone (vs. machine generated ozone) in regions with high 

73
3/14/2000 AC 25-22 

levels of ozone provided the cognizant ACO finds them acceptable. The airplane engines and 

ventilation systems by themselves reduce the concentration without any additional filtering. 

Using the ratio of the ozone introduced to the amount of ozone retained in the cabin environment 

(the retention ratio), the airplane operating characteristics in terms of the ozone concentration to 

be expected in the occupied areas can be determined. If an ozone filter or catalytic converter is 

installed, the manufacturer may perform a similar ground test (or an equivalent laboratory test 

acceptable to the cognizant ACO) to demonstrate that the ozone concentration in the airplane 

will never exceed the ozone limits in service, regardless of the atmospheric concentration. 

Typically, a maintenance manual limit is established for removing and cleaning the units. The 

maintenance interval ensures that the converters will be removed when the efficiency of the 

converter degrades to the point that the airplane would no longer be in compliance with § 25.832. 

(3) Use of Default Retention Ratio. The following guidance is extracted from an FAA 

letter dated February 18, 1999. It was developed in response to an inquiry regarding certification 

of cabin ozone limits to satisfy § 25.832: 

(a) One letter requested FAA concurrence that for future certification programs a 

value of 0.7 be accepted as a default value for the ozone retention ratio when demonstrating 

compliance to 14 CFR § 25.832. 

(b) An applicant for a type certificate under part 25, or a part 121 operator, must 

demonstrate compliance with the maximum ozone concentration requirements of § 25.832 or 

§ 121.578, respectively. These two regulations were introduced per Amendments 25-50 

and 121-154. Amendments No. 25-56 and 121-181 made revisions stipulating these ozone 

concentration limits be applicable at revised flight levels. Section 25.832 stipulates a limit for a 

time weighted average during any three hour interval above flight level 270. Section 121.578 

stipulates a limit for a time weighted average for each flight segment that exceeds four hours and 

includes flight above flight level 270 during that segment. The time difference is referenced in 

the preamble material in Notice No. 78-15 and ensuing Amendments No. 25-50 and 121-154. 

As explained in the notice, approximately one hour of the scheduled flight segment would not be 

at higher altitudes. This was based on conservative times for start, taxi, takeoff, climb, descent, 

approach, and landing. The operating rule considers the entire flight segment involving time at 

lower altitudes. 

(c) Certification for ozone concentration limits under both parts 25 and 121 entails 

demonstrating that the cabin and cockpit ozone concentration will not exceed the limits 

mandated by the regulations. In either case, the testing and/or analysis for showing compliance 

utilizes a retention ratio. This ratio is the fractional value of the cabin interior ozone 

concentration compared with the ambient concentration. 

(d) The FAA published a table of previously measured transport airplane retention 

ratios in Notice No. 81-15. The notice also stipulated that the FAA would accept a retention 

ratio of 0.7 to facilitate compliance to § 121.578 when computing cabin ozone concentrations, 

even when a retention ratio has not been measured for that particular airplane. The final rule 

74 
3/14/2000 AC 25-22 

preamble (Amendment Nos. 25-56 and 121-181) incorporating the proposed changes to both part 

25 and part 121 states that the FAA considers it reasonable to accept the value of 0.7. 

(e) Since that notice was published, retention ratios have been determined for a 

number of airplanes. The retention values that have been demonstrated on turbine-powered 

aircraft indicate that 0.7 continues to be a reasonable value. The FAA accepts a default value of 

0.7 for the ozone retention ratio when demonstrating compliance to both §§ 25.832 and 121.578. 

If you propose to use a retention value of less than 0.7, the FAA will require certification testing 

to validate the proposed retention level. 

(4) Use of Operational Limitations. The following guidance is extracted from an FAA 

memorandum dated March 31, 1997. It was developed in response to inquiries regarding 

certification of cabin ozone limits to satisfy § 25.832: 

(a) The cabin ozone limits specified in § 25.832 were first introduced in 

Amendment 25-50, and were simultaneously placed into part 121. When the part 25 rule went 

into effect, it was expected that many of the existing airplanes in the fleet would not be modified 

with ozone filters; since § 25.832 was not part of their type design, these airplanes did not have 

to comply with this rule. Most operators of these airplanes chose, instead, to comply with the 

ozone limits required by § 121.578 by way of operational considerations, i.e., reduced cruise 

altitudes, amended routes of flight, etc. For the existing airplanes at that time, this solution was 

deemed the most economical method of compliance, and Advisory Circular (AC) 120-38 was 

developed to provide guidance on acceptable methods of operational compliance to § 121.578. 

(b) Compliance with § 25.832 requires a demonstration (by test or analysis) that 

the cabin ozone concentration will not exceed the stated limits at the maximum atmospheric 

ozone concentration expected in service. If an ozone filter or other device is installed, it must be 

shown that this device, plus the natural ozone dissociation which occurs in the engines and in the 

airplane air conditioning system, will meet the requirements of the rule at the airplane’s 

certificated altitude ceiling. If an ozone filter or other device is not used, however, certification 

testing may reveal that the cabin ozone concentration may exceed the limits at the airplane’s 

certificated altitude ceiling, and this would cause the certificated ceiling of the airplane to be 

lowered to whatever altitude would allow compliance with § 25.832. Proposals have been made 

to comply with § 25.832 by applying operational limitations to the Airplane Flight Manual in 

lieu of installing ozone control equipment. These limitations would restrict the airplane to 

certain maximum altitudes based on the latitude and time of year. Such data is currently 

considered acceptable in finding compliance to § 121.578. This concept could be an option for 

compliance to § 25.832, but a part 25 approval is not geographically limited. Therefore, charts 

need to be available for ozone limitations globally for this method of complying with § 25.832. 

This type of compliance would have to be approved by the Manager of the responsible Aircraft 

Certification Office. 

(5) Meaning of the Term Time-Weighted Average in § 25.832. The following policy 

material is extracted from an FAA letter dated October 26, 1987. It was in response to a request 

75
3/14/2000 AC 25-22 

for information on the calculation of time-weighted-average ozone exposure for compliance with 

§ 25.832(a)(2). 

 (a) Time-weighted-average (TWA) ozone exposure, as described in AC 120-38, is 

calculated for flight segments that occur in different atmospheric ozone concentrations by 

averaging the individual exposures and the duration of each exposure over the entire flight. This 

concept can be explained by an example. 

(b) Assume a 2 hour flight in zero parts per million (PPM) atmospheric ozone 

concentration, and then a 4 hour segment in a 1.3 ppm concentration. Assume also a cabin ozone 

retention ratio of 0.7, a cabin altitude of 7,000 feet, and no ozone filter installed. 

(c) The time-weighted-average (TWA) atmospheric ozone concentration is 

calculated as follows: 

(0.0ppm)(2 hrs) + (1.3ppm)(4 hrs) = 0.867 ppm time-weighted-average (TWA) 

 6 hrs atmospheric ozone concentration 

(d) The 7,000-foot cabin altitude results in a ratio of cabin altitude/sea level 

pressure of 0.77 (the allowable cabin ozone concentration is corrected to sea level conditions). 

The ozone retention ratio of 0.7 is a value that the FAA will accept without a test to determine 

the actual value.: 

 1 This is equation 6 from Appendix 1 of AC 120-38: 

0Z16 = (l-E)(R)(P/PO) (OZ )iTi / T F S

i

N

16 =1

∑⎡

⎣⎢ ⎤⎦⎥

 2 Using equation 6 from Appendix 1 of AC 120-38, the calculation of the 

resultant cabin ozone concentration for this sample flight is as follows: 

TWA = 0.77 x 0.7 x 0.867 = 0.467 ppm cabin ozone concentration. 

(e) In this example, the cabin ozone concentration exceeds the 0.1 ppm limit in 

§ 121.578, and the flight could not be made under these conditions. Either the altitude, route of 

flight, or duration of exposure would have to be changed to make the flight acceptable. This 

procedure is explained, with several examples, in AC 120-38. 

(f) The calculation of the TWA ozone exposure for compliance with 

§ 25.832(a)(2) is done in the same way except that the TWA value would be based on the 4-hour 

segment, not the 6-hour segment used in the example. 

(6) Example of Compliance with § 25.832. The following is extracted from an 

applicant's letter in response to the FAA’s request for an example of the calculation methods 

used to certify previous models under § 25.832, Cabin Ozone Concentration. Included below are 

76 
3/14/2000 AC 25-22 

generic examples of the calculations used to show compliance to the maximum and timeweighted average ozone concentration limits. 

(a) For this example, extracted from report no. FAA-EQ-78-03 (reference (b) of 

applicant's letter), the maximum certified flight level of the airplane will be FL430. The highest 

ambient ozone concentrations at flight level 430 are at 80° N latitude, Eastern North America 

during the month of January (atmospheric ozone concentration, confidence level of 84%). The 

minimum ozone converter efficiency required for southern and northern hemisphere flights are 

the same; however, while the worst seasons for high ambient ozone concentrations in the 

southern hemisphere are the summer and fall, the worst seasons for the northern hemisphere are 

the winter and spring. 

(b) In lieu of an experimentally based retention factor, a retention factor of 0.7 

will be used in the following analysis extracted from Docket No. 22438, Notice No. 81-15 

(reference (c) of applicant's letter). The FAA will accept a retention factor of 0.7 when 

computing cabin ozone concentrations if a retention factor has not been measured for the 

airplane being analyzed. 

(c) To comply with § 25.832, an airplane must be designed so that the cabin ozone 

concentration does not exceed a maximum ozone concentration level of 0.25 parts per million by 

volume (ppmv), sea level equivalent (SLE), at any time above flight level 320; and time

weighted average (TWA) limit 0.1 ppmv, SLE, during any three hour time interval above flight 

level 270. Compliance will be shown by statistical analysis as described in Advisory Circular 

120-38 (reference (d) of applicant's letter). 

(d) The following is the calculation of the minimum ozone converter efficiency 

required to meet the maximum cabin ozone concentration limit (OZMAX) of § 25.832 during a 

77
3/14/2000 AC 25-22 

worst case flight. Using equation 1 from Appendix 1 of AC 120-38, the calculation of the 

resultant cabin ozone concentration is as follows: 

 1 Given: 

 Month: January 

 Latitude: 80° N 

 Flight Level: 430 

Flight Time: 3 Hours Cruise; 4 Hours Total 

 *( ) 1 E *OZ16

P

P OZMAX R * o

= − 

where: 

OZMAX = Maximum Cabin Ozone Concentration Limit = 0.25 ppmv (§ 25.832) R = Airplane Ozone Retention Factor = 0.7

Po

P = Ratio of Cabin Pressure to Sea Level Atmospheric Pressure = 0.74 

E = Efficiency of Ozone Converter

OZ16 = Ambient Ozone Concentration = 1.6 PPMv (From report no. FAA-EQ-78-03) 

Solving: 

*OZ16

P

P R *

OZMAX E 1

o

= −

0.70

0.7 * 0.74 *1.6

0.25 E =1− =

E = 70 % CONVERTER EFFICIENCY; minimum converter efficiency required to comply with 

§ 25.832 OZMAX limit. 

(e) The following is the calculation of the minimum ozone converter efficiency 

required to meet the time-weighted average cabin ozone concentration limit (OZTWA) of

§ 25.832 during a worst case flight. Using equation 6 from Appendix 1 of AC 120-38, the 

calculation of the resultant cabin ozone concentration for is as follows: 

78 
3/14/2000 AC 25-22 

 1 Given: 

 Month: January 

 Latitude: 80° N 

 Flight Level: 430 

Flight Time: 3 Hours Cruise; 4 Hours Total 

∑=

= −

N

i 1 FS

i

i

o T

T * (OZ16) * PP OZTWA (1 E)*R * 

where: 

OZTWA = Time-Weighted Average Cabin Ozone Concentration Limit = 0.10 ppmv (§ 25.832) 

E = Efficiency of Ozone Converter

R = Airplane Ozone Retention Factor = 0.7

Po

P = Ratio of Cabin Pressure to Sea Level Atmospheric Pressure =0.74

N = Number of Individual Flight Segments = 3 

OZ16 = Ambient Ozone Concentration = 1.6 ppmv (from report no. FAA-EQ-78-03) 

Ti = Time of Individual Flight Segment = 1 Hour 

TFS = Total Time of Flight = 4 Hours 

Solving: 

i

N

i 1

i

o

FS

* (OZ16) *T

P

P R *

OZTWA*T E 1

∑=

= −

[ ] ( ) ( ) ( ) 0.84

0.70* 0.74* 1.6 *1 1.6 *1 1.6 *1

0.10* 4 E 1 = + + = −

E = 84 % CONVERTER EFFICIENCY; minimum converter efficiency required to comply with 

§ 25.832 OZTWA limit. 

(f) When calculating the minimum ozone converter efficiency required to satisfy 

the 14 CFR time-weighted average limit of 0.1 ppmv, a four-hour flight was assumed, taking a 

half-hour each for takeoff and landing and three hours for cruise. 

(g) Based on the above calculations, the minimum required ozone converter 

efficiency for the airplane to comply with § 25.832 of 14 CFR is 84%. 

 e. References. The addresses for ordering the latest revision of the advisory circulars and 

other referenced documents listed below can be found in the Appendix to this AC.

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3/14/2000 AC 25-22 

AC 120-38, Transport Category Airplanes Cabin Ozone Concentrations. 

FAA Report FAA-AM-79-20, Effects of Ozone on Exercising and Sedentary Adult Men 

and Women Representative of the Flight Attendant Population. 

FAA Report FAA-AEQ-77-13, Ozone Concentration by Latitude, Altitude, and Month, 

Near 80o West.

FAA Report FAA-AM-80-9, Effects of Ozone (0.30 Parts Per Million, ~ 600μg/m3

) on 

Sedentary Men Representative of Airline Passengers and Cockpit Crewmembers. 

FAA Report FAA-AM-80-16, Effects of Long-Term Exposure to Low Levels of 

Ozone: A Review. 

32. SECTION 25.833 - COMBUSTION HEATING SYSTEMS.

 a. Rule Text. 

 Combustion heaters must be approved.

[Amdt. 25-72, 55 FR 297783, Jul 20, 1990]

b. Intent of Rule. The intent of this rule is that combustion heaters installed on transport 

category airplanes must be approved. Technical Standard Order (TSO) C-20 provides standards 

for approval of these devices. 

 c. Background. Combustion heaters were common in airplanes certificated to parts 3 and 

4b of the Civil Air Regulations (CAR). These airplanes, usually powered by reciprocating 

engines, supplemented various engine exhaust heating systems by separate combustion heaters. 

There are few combustion heaters in use in modern part 25 airplanes, but their use is not 

prohibited. The certification requirements for combustion heaters appeared in § 4b.372 of the 

CAR, entitled “Heating Systems,” which stated: "Combustion heaters shall be of an approved 

type and shall comply with the fire protection requirements of § 4b.386. Engine exhaust heaters 

shall comply with the provisions of § 4b.467(c) and (d)." Section 4b.386 contained extensive 

standards for combustion heater fire zones, ventilating air ducts, combustion air ducts, heater 

controls, air intakes, and exhaust systems. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Standards for Certification of Combustion Heaters. The following policy material 

is extracted from an FAA memorandum dated March 14, 1948, in response to questions 

regarding standards for certification of combustion heaters to § 4b.372 of the CAR. 

(a) Heater Approval. The approval of heater units is considered essential to safety 

in all cases. In this regard two separate classes of approval are evidently in order to cover the 

various applications that can be foreseen. 

80 
3/14/2000 AC 25-22 

 1 The first class of approval will cover units which will be considered 

eligible for use only as space heaters. In the case of such units, the criteria for acceptance will 

involve the general airworthiness of the unit under vibration and operating temperatures and 

considerations of safety from the standpoint of fire and carbon monoxide contamination only. 

Heater performance, i.e., the ability of the heater to deliver a given output between normal 

overhaul periods, ability to start under all expected conditions of altitude and temperature, etc., 

will not enter into such approvals as this does not normally constitute a safety item in the case of

space heaters. 

 2 The second class of approval will involve heater performance in addition 

to the criteria for the first class of approval. This type of approval will cover units that are to be 

used for de-icing purposes since heater performance obviously becomes a safety consideration 

when the heater is to be used for windshield, surface, or propeller de-icing. 

 3 Technical Standard Order TSO-C-20 provides standards for combustion 

heaters and should be used for new installations. 

(b) Heater Isolation. Heaters must be isolated from the remainder of the airplane. 

However, this need not necessarily mean a complete shield around the entire heater unit 

(although this would be satisfactory) since in many heater designs, the air jacket largely 

surrounds the flame chamber. Thus, the heater design itself practically provides a steel shield 

between the combustion unit and the remainder of the airplane. In such cases, it is acceptable to 

provide isolation for the fuel system components mounted on the heater, and for the heater 

exhaust system and combustion chamber drains. The following schematic sketch shows an 

example of an installation which should be satisfactory: 

 1 The shut-off valve shown in the sketch should be provided if there are fuel 

system components within the ventilating air shroud which may be subject to leakage or failure. 

In such cases, that portion of the ventilating air duct up to the valve, as well as the valve itself,

should be of fire resistant construction and the valve should provide as flame tight a seal as 

possible. If the fuel system is so arranged that there are no fittings or connections within the 

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3/14/2000 AC 25-22 

ventilating air shroud, the downstream air shut-off valve and fire resistant duct between the 

heater and the valve may be dispensed with. 

 2 For shrouds for the combustion chamber drain lines, the necessity for 

these will generally depend upon the location of the drain in the heater. If the drain outlet from

the combustion chamber is so located that products of combustion can issue through the drain 

line, it will no doubt become hot and require isolation. However, drains are sometimes 

connected in such a manner that they do not carry exhaust gases and remain relatively cool. In 

such cases, shrouds are not necessary. 

(c) Fire Detector and Extinguisher Equipment. 

 1 For transport category airplanes, extinguishers and detectors will normally 

be required, in addition to isolation provisions. 

 2 Detectors and extinguishers should be provided wherever potential sources 

of fuel leakage and sources of ignition are in close proximity. In the foregoing sketch, the space 

within the shield would require such protection. In addition, detectors and extinguisher nozzles 

should be installed in the ventilating air passages of the heater if this chamber contains fuel 

system fittings or connections that may be subject to leakage. Hand fire extinguishers should be 

considered equivalent to a fixed fire extinguisher installation only when the heater is located in 

such a manner that it is readily accessible to the crew and when all fire zones in the installation 

can easily be reached with a hand extinguisher. All extinguishers may also be dispensed with 

when the heater is so shielded and located that a fire could be permitted to burn itself out without 

danger of damage to any important structural members or otherwise endangering the safety of 

the airplane.

 3 Detectors may be dispensed with only when the heater is so located that 

the occurrence of fire would immediately be noted by the crew. 

 (d) Heater Fuel System. The heater fuel system should comply with airworthiness 

standards for the engine fuel system as regards fuel lines, fittings and accessories. While in 

flight, valves should be provided for shutting off the flow of fuel at its source, unless equivalent 

provisions in the form of a separate heater fuel pump are available. In the case of transport 

category airplanes, fuel lines located in fire zones should, in addition, be of fire resistant 

construction. In the foregoing sketch, this would apply to that portion of the fuel line within the 

shield and in the heater ventilating air compartment. All pressure lines should comply with the 

provisions of §§ 03.4211 and 04.4211 of the CAR, regarding pressure cross feed arrangements. 

 e. Reference. The address for ordering the latest revision of the referenced document listed 

below can be found in the Appendix to this AC.

Technical Standard Order No. TSO-C-20, Combustion Heaters. 

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3/14/2000 AC 25-22 

33. SECTION 25.841 - PRESSURIZED CABINS.

 a. Rule Text.

(a) Pressurized cabins and compartments to be occupied must be equipped to 

provide a cabin pressure altitude of not more than 8,000 feet at the maximum 

operating altitude of the airplane under normal operating conditions. 

(1) If certification for operation above 25,000 feet is requested, the airplane must 

be designed so that occupants will not be exposed to cabin pressure altitudes in 

excess of 15,000 feet after any probable failure condition in the pressurization 

system. 

(2) The airplane must be designed so that occupants will not be exposed to a 

cabin pressure altitude that exceeds the following after decompression from any 

failure condition not shown to be extremely improbable: 

(i) Twenty-five thousand (25,000) feet for more than 2 minutes; or 

(ii) Forty thousand (40,000) feet for any duration. 

(3) Fuselage structure, engine and system failures are to be considered in 

evaluating the cabin decompression. 

(b) Pressurized cabins must have at least the following valves, controls, and 

indicators for controlling cabin pressure: 

(1) Two pressure relief valves to automatically limit the positive pressure 

differential to a predetermined value at the maximum rate of flow delivered by the 

pressure source. The combined capacity of the relief valves must be large enough 

so that the failure of any one valve would not cause an appreciable rise in the 

pressure differential. The pressure differential is positive when the internal 

pressure is greater than the external. 

(2) Two reverse pressure differential relief valves (or their equivalents) to 

automatically prevent a negative pressure differential that would damage the 

structure. One valve is enough, however, if it is of a design that reasonably 

precludes its malfunctioning. 

(3) A means by which the pressure differential can be rapidly equalized. 

(4) An automatic or manual regulator for controlling the intake or exhaust 

airflow, or both, for maintaining the required internal pressures and airflow 

rates. 

(5) Instruments at the pilot or flight engineer station to show the pressure 

differential, the cabin pressure altitude, and the rate of change of the cabin 

pressure altitude. 

(6) Warning indication at the pilot or flight engineer station to indicate when the 

safe or preset pressure differential and cabin pressure altitude limits are 

exceeded. Appropriate warning markings on the cabin pressure differential 

indicator meet the warning requirement for pressure differential limits and an 

aural or visual signal (in addition to cabin altitude indicating means) meets the 

warning requirement for cabin pressure altitude limits if it warns the flight crew 

when the cabin pressure altitude exceeds 10,000 feet. 

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3/14/2000 AC 25-22 

(7) A warning placard at the pilot or flight engineer station if the structure is not 

designed for pressure differentials up to the maximum relief valve setting in 

combination with landing loads. 

(8) The pressure sensors necessary to meet the requirements of paragraphs (b)(5)

and (b)(6) of this section and 25.1447(c), must be located and the sensing system 

designed so that, in the event of loss of cabin pressure in any passenger or crew 

compartment (including upper and lower lobe galleys), the warning and 

automatic presentation devices, required by those provisions, will be actuated 

without any delay that would significantly increase the hazards resulting from 

decompression. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 

41 FR 55466, Dec. 20, 1976; Amdt. 25-87, 61 FR 28696, Jun. 5, 1996]

b. Intent of Rule. This rule provides standards for pressurized compartments in transport 

category airplanes. The requirements for various controls and pressure relief valves are 

addressed. Testing required for demonstrating compliance with many of the requirements of this 

section are addressed in § 25.843. 

 c. Background. The rules in this section originated in sections 4b.373 through 4b.376 of 

the Civil Air Regulations, and were carried over essentially unchanged when part 25 was 

codified. 

 (1) Amendment 25-38 (December 20, 1976) added § 25.841(b)(8). This amendment 

requires pressure sensors and systems designed so the indicators (and automatic oxygen mask 

presentation required in § 25.1447(c)) will be actuated without any delay that would significantly

increase the hazards resulting from decompression. 

 (2) Amendment 25-87 (June 5, 1996) changed section 25.841(a) based largely on 

special conditions that had been applied to aircraft operating at high altitudes. The objective of 

the amended § 25.841(a) (when applied in conjunction with amended § 25.1447(c)) was to 

provide airworthiness standards that allow subsonic airplanes to operate at their maximum

achievable altitudes. This is the highest altitude for which an applicant chooses to demonstrate 

that, after decompression caused by a single failure or combination of failures that are not shown 

to be extremely improbable: (1) the flightcrew will remain alert and be able to fly the airplane; 

(2) the cabin occupants will be protected from the effects of hypoxia; and (3) in the event that 

some occupants do not receive supplemental oxygen, they nevertheless will be protected against 

permanent physiological damage. Compliance to Amendment 25-87 is required for only those 

airplanes whose certification basis includes this amendment. Most of the currently produced 

airplanes do not fall into this category.

 (3) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group will be tasked to 

recommend revisions to §§ 25.831, 25.841 and related regulations as well as a revised AC 

25.20. 

84 
3/14/2000 AC 25-22 

 d. Policy/Compliance Methods. Flight tests for demonstrating compliance with the 

requirements of the section are addressed in § 25.843. For guidance on compliance with this 

requirement, refer to the preamble of this rule and the following information. Advisory Circular 

(AC) 25-7A, Flight Test Guide for Certification of Transport Category Airplanes, provides 

additional guidance on tests to verify the 8,000 feet and 15,000 feet cabin altitude requirements 

as well as tests for high altitude (i.e., greater than 8,000 feet) airports. 

(1) Rapid Equalization of Pressure. The following policy is extracted from an FAA 

memorandum dated June 21, 1982, in response to an inquiry about the intent of, and need for, 

§ 25.841(b)(3). 

(a) Section 25.841(b)(3) ensures that a means exists to provide rapid equalization 

of pressure in the event that (1) a plug type door must be opened for an emergency, and (2) a 

windshield cracks (§ 25.775(d)). Part 25 manufacturers typically have complied with this 

regulation by providing a manual position indication and control for the outflow valves. Later 

certification documents reference, in addition to the manual control, the ability to shut off cabin 

air inflow as a means of meeting this requirement. 

(b) The FAA does not recommend a specific depressurization rate other than the 

pilot being able to control the rate. In other words, in the manual pressure control mode, he 

should be able to set the rate as fast or as slow as the pilot judges the emergency requires or as 

the flight manual specifies. Typical maximum rates for large aircraft will open the outflow valve 

in 30 seconds, and the depressurization that follows will occur in two minutes or less (with the 

air conditioning packs on). 

(c) From a design and certification standpoint, the need to establish a rate of 

equalization of pressure can be assessed from failure analyses, abnormal procedures, and the 

need to prevent a failure from becoming a disaster. 

(2) Access to a Pressurized Cargo Compartment in the Engine Burst Zone During High 

Altitude Operation. The following policy is extracted from an FAA memorandum dated June 25, 

1986, in response to an inquiry regarding access to a cargo compartment located in the engine 

burst zone during operation above 10,000 feet. 

(a) Section 25.841(a)(2) requires that the cabin pressure altitude not exceed 40,000 

feet for any duration, or 25,000 feet for more than two minutes after any failure not shown to be 

extremely improbable. Section 25.841(a)(3) states that engine failures are to be considered in 

evaluating the cabin decompression, this includes rotor burst. Prior to Amendment 25-87, these 

stipulations were made through special conditions on several aircraft. Compliance has been 

shown for many airplanes by locating the pressure bulkhead forward of the aft mounted engines 

for aft fuselage mounted engines. 

(b) In some cases, the primary pressure bulkhead of the airplane is located aft of 

the engines; however, an additional pressure bulkhead installed forward of the cargo 

85
3/14/2000 AC 25-22 

compartment, forward of the engine, and forward of the aft pressure bulkhead was proposed to 

meet these requirements. The additional pressure bulkhead is likely to have a door that is used 

for access to the cargo compartment from the main passenger compartment. Should a 

depressurization of the cargo compartment occur, the integrity of this pressure bulkhead is 

dependent on the door being closed. This arrangement is considered satisfactory for compliance 

with the above requirements, provided there is a means of ensuring that the door is not opened 

during high altitude flight. 

(c) Allowing the door to the aft cargo compartment of the airplane to be opened at 

high altitudes would not ensure the level of safety established by the high altitude special 

conditions and subsequent rule change. Preventing access to the cargo compartment to ensure 

that the door is always closed and latched at the higher altitudes is an inconvenience. Most 

present day airplane designs provide overhead storage bins or limited underseat storage in the 

main cabin. With some forethought, the limited baggage needed during flight can be stored in 

these locations prior to operation at high altitudes without exposing the airplane and occupants to 

possible catastrophic decompression. For these reasons, the FAA has determined that these 

conditions do not impose an undue burden on the applicant.

(d) In view of the above, the applicant must provide a positive means of ensuring 

that: (1) The cargo compartment is inaccessible and unoccupied for high altitude flight; (2) There 

is an indication in the cockpit to warn the flight crew when the cargo compartment door is not 

closed and latched; and (3) There is a means to automatically close the door. A flight manual 

limitation should specify that the cargo compartment door must be closed and latched during any 

operation above 10,000 feet. 

(3) Cabin Altitude Limit; § 25.841(a). The following policy is extracted from an FAA 

memorandum dated January 12, 1994, in response to the question: Is the 8,000 foot cabin 

altitude limit in § 25.841(a) an absolute limit?

(a) Background. The pressurization systems on the airplanes of interest allow 

cabin altitudes to exceed the 8,000 feet maximum allowed by § 25.841(a). The manufacturer 

states that the systems currently used on these airplanes allow 8,000 feet plus or minus 200 feet. 

It is also noted that the cabin altitude display employs a tolerance of an additional plus or minus 

250 feet. The applicant states that this design does not result in an unsafe condition, and while 

the system does not literally comply with the 8,000 feet limit, the intent of § 25.841(a) is met. 

(b) The preamble to both the Notice of Proposed Rulemaking (NPRM) and the 

final rule associated with part 4b of the Civil Air Regulations (CAR) have been reviewed. 

Amendment 4b-6, effective August 12, 1957, changed the requirements in section 4b.374 of the 

CAR to limit cabin altitude to 8,000 feet. Prior to that amendment, the limit was 10,000 feet. 

This limit was carried over into § 25.841 of 14 CFR. The NPRM states, ". . . it is proposed to 

change § 4b.374 to require on airplanes intended for operation over 25,000 feet to maintain 

under normal conditions a pressure altitude of no more than 8,000 feet. . ." The wording adopted 

was similar to that found in § 25.841(a). When deciding on the interpretation of a rule, the 

policy of the FAA has been to determine the intent of the parties preparing the rule. The use of 

86 
3/14/2000 AC 25-22 

the words "no more than 8,000 feet" indicates that the rule was intended to mandate a cabin 

altitude limit of 8,000 feet. There is nothing in either preamble to indicate that 8,000 feet plus 

some equipment tolerance was the goal. Further, the pressurization systems on many other 

airplanes are designed to limit cabin altitude to 7,950 feet including equipment tolerances. 

(c) It is recognized that the difference in the partial pressure of oxygen at 8,000 

versus 8,450 feet is not significant. There is, of course, a difference, and if a passenger has a 

breathing impairment, there could be some effect on their health. There is no adverse effect 

associated with the cabin altitude indicator tolerance because oxygen partial pressure is a 

function of actual cabin altitude. 

(d) Based on the above review, the FAA has determined that issuing a new or 

amended Type Certificate (TC) for an airplane which does not meet the criteria as written is not 

in the public interest. When a maximum (or minimum) value of a parameter is provided in a 

section of 14 CFR, the airplane should be shown to meet that limit in normal operation, including 

expected equipment and sensor manufacturing tolerances. This applies to the actual measured 

value as determined in FAA flight or ground tests, but does not apply to tolerance in the display 

device, such as the cabin altitude indicator. A displayed value of 50 or 100 feet above the 8,000 

feet limit would not be considered unacceptable, as long as the actual pressure altitude in the 

cabin does not exceed 8,000 feet. 

(e) For airplanes that are already certificated, it is not necessary to require a 

modification to revise the pressurization control system. That action would be required only if 

an unsafe condition warranting mandatory action is identified.

 e. References. The addresses for ordering the latest revision of the advisory circular (AC) 

listed below can be found in the Appendix to this AC. 

AC 25-20, Pressurization, Ventilation and Oxygen Systems Assessment for Subsonic 

Flight Including High Altitude Operation. 

AC 25-7A, Flight Test Guide for Certification of Transport Category Airplanes. 

34. SECTION 25.843 - TESTS FOR PRESSURIZED CABINS.

 a. Rule Text:

(a) Strength test. The complete pressurized cabin, including doors, windows, and 

valves, must be tested as a pressure vessel for the pressure differential specified in

25.365(d). 

(b) Functional tests. The following functional tests must be performed: 

(1) Tests of the functioning and capacity of the positive and negative pressure 

differential valves, and of the emergency release valve, to simulate the effects of 

closed regulator valves.

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3/14/2000 AC 25-22 

(2) Tests of the pressurization system to show proper functioning under each 

possible condition of pressure, temperature, and moisture, up to the maximum 

altitude for which certification is requested. 

(3) Flight tests, to show the performance of the pressure supply, pressure and 

flow regulators, indicators, and warning signals, in steady and stepped climbs 

and descents at rates corresponding to the maximum attainable within the 

operating limitations of the airplane, up to the maximum altitude for which 

certification is requested. 

(4) Tests of each door and emergency exit, to show that they operate properly 

after being subjected to the flight tests prescribed in paragraph (b)(3) of this 

section. 

b. Intent of Rule. This rule provides specific tests that must be conducted in demonstrating 

compliance with the cabin pressurization requirements contained in § 25.841. 

 c. Background. The standards in this section originated in section 4b.376 of the Civil Air 

Regulations (CAR), and were carried over essentially unchanged when part 25 was codified. 

 (1) Amendment 25-87 (June 5, 1996) added another factor to § 25.365(d), referred to in 

§ 25.843(a). Airplanes operating up to 45,000 feet use a factor of 1.33; airplanes operating over 

45,000 feet use a factor of 1.67. Compliance to Amendment 25-87 is required for only those 

airplanes whose certification basis includes this amendment. Most of the currently produced 

airplanes do not fall into this category. 

 (2) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group will be tasked to 

recommend revisions to §§ 25.365, 25.831, 25.841, and related regulations as well as revisions to 

AC 25.20. 

88 
3/14/2000 AC 25-22 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. While flight tests to demonstrate 

compliance are not specifically required in this section, the functional tests identified are usually 

conducted during flight tests because the pressurization system, including the positive and 

negative pressure relief valves, can be exercised during other flight test activities, and many of 

the tests are more meaningful when conducted in flight. The functional tests involving 

environmental factors (temperature and humidity) are generally conducted on the ground under 

laboratory conditions. 

(1) Pressure Tests For Compliance with § 25.843(a): The following policy is extracted 

from an FAA memorandum dated February 21, 1990, in response to concerns regarding cabin 

test pressures for demonstrating compliance with § 25.843(a). 

(a) Some manufacturers conduct cabin pressure vessel tests to a pressure 

corresponding to the maximum relief valve setting, while others test to 1.33 times the maximum 

relief valve setting. Effective June 5, 1996, if the airplane certification basis includes 

Amendment 25-87, the factor 1.33 (ref. § 25.365(d)) applies for operation to 45,000 feet while a 

factor of 1.67 applies for operation above 45,000 feet. While there is not a regulatory 

requirement to test each and every production airplane to the 1.33 times maximum relief valve 

setting, many major manufacturers test to this pressure to enhance the fatigue characteristics of 

the structure. The FAA recommends that modifiers follow the lead of these major 

manufacturers. If the airframe was originally tested to the 1.33 times relief valve setting, then 

the modified airplane should also be tested to the same pressure. 

 e. Reference. The address for ordering the latest revision of the advisory circular listed 

below can be found in the Appendix to this AC. 

AC 25-7A, Flight Test Guide for Certification of Transport Category Airplanes.

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3/14/2000 AC 25-22 

Section 5. FIRE PROTECTION 

35. SECTION 25.851 - FIRE EXTINGUISHERS.

 a. Rule Text.

(a) Hand fire extinguishers. 

(1) The following minimum number of hand fire extinguishers must be 

conveniently located and evenly distributed in passenger compartments: 

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

Passenger Capacity Number of extinguishers 

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

7 through 30 1 

31 through 60 2 

61 through 200 3 

201 through 300 4 

301 through 400 5 

401 through 500 6 

501 through 600 7 

601 through 700 8 

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

(2) At least one hand fire extinguisher must be conveniently located in the pilot 

compartment. 

(3) A least one readily accessible hand fire extinguisher must be available for use 

in each Class A or Class B cargo compartment and in each Class E cargo or 

baggage compartment that is accessible to crewmembers in flight. 

(4) At least one hand fire extinguisher must be located in, or readily accessible 

for use in, each galley located above or below the passenger compartment. 

(5) Each hand fire extinguisher must be approved. 

(6) At least one of the required fire extinguishers located in the passenger 

compartment of an airplane with a passenger capacity of at least 31 and not more 

than 60, and at least two of the fire extinguishers located in the passenger 

compartment of an airplane with a passenger capacity of 61 or more must contain 

Halon 1211 (bromochlorodifluoromethane CBrC1F2), or equivalent, as the 

extinguishing agent. The type of extinguishing agent used in any other 

extinguisher required by this section must be appropriate for the kinds of fires 

likely to occur where used. 

(7) The quantity of extinguishing agent used in each extinguisher required by this 

section must be appropriate for the kinds of fires likely to occur where used. 

(8) Each extinguisher intended for use in a personnel compartment must be 

designed to minimize the hazard of toxic gas concentration. 

(b) Built-in fire extinguishers. If a built-in fire extinguisher is provided- 

(1) Each built-in fire extinguisher system must be installed so that-- 

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(i) No extinguishing agent likely to enter personnel compartments will be 

hazardous to the occupants; and 

(ii) No discharge of the extinguisher can cause structural damage. 

(2) The capacity of each required built-in fire extinguishing system must be 

adequate for any fire likely to occur in the compartment where used, considering 

the volume of the compartment and the ventilation rate. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-54, 

45 FR 60173, Sep. 11, 1980; Amdt. 25-72, 55 FR 29783, Jul. 20, 1990; Amdt. 25-

74, 56 FR 15456, Apr. 16, 1991]

b. Intent of Rule. This rule prescribes the standards for both hand held and built-in fire 

extinguishers. Section 25.851(a), which addresses hand fire extinguishers, is addressed in 

Advisory Circular (AC) 25-17. Section 25.851(b) is intended to ensure that the built-in fire 

extinguishing system does not introduce a hazard to occupants or the airplane structure, and that 

the system is adequate to control any fire likely to occur. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the of the Civil Air Regulations (CAR). The 

standards in section 4b.380(b) of the CAR were carried over essentially unchanged to § 25.851 

of 14 CFR. While § 25.851 has been amended several times, the standards for built-in fire 

extinguishers have not been not changed.

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. Flight tests are conducted, in which 

the fire extinguishing system is activated and the extinguishent concentration is measured. The 

common method is to utilize a gas spectrum analyzer to sample the extinguishent concentration 

in real time at several locations in the compartment. An advisory circular is being written that 

will provide guidance on certification of cargo compartment fire extinguishing or suppression 

systems. Guidance related to compliance with § 25.851(b)(2) is included under §§ 25.855 and 

25.857 of this AC. 

 e. References. None. 

36. SECTION 25.854 - LAVATORY FIRE PROTECTION.

 a. Rule Text.

For airplanes with a passenger capacity of 20 or more: 

(a) Each lavatory must be equipped with a smoke detector system or equivalent 

that provides a warning light in the cockpit, or provides a warning light or 

audible warning in the passenger cabin that would be readily detected by a flight

attendant; and 

(b) Each lavatory must be equipped with a built-in fire extinguisher for each 

disposal receptacle for towels, paper or waste, located within the lavatory. The 

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3/14/2000 AC 25-22 

extinguisher must be designed to discharge automatically into each disposal 

receptacle upon occurrence of a fire in that receptacle. 

[Amdt. 25-74, 56 FR 15456, Apr. 16, 1991]

b. Intent of Rule. This section provides standards for certification of lavatory smoke 

detection and waste chute fire extinguishing systems. 

 c. Background. Prior to the adoption of § 25.854, part 121 was amended in March 1985, 

by the addition of § 121.308, to require that each lavatory be equipped with smoke detectors, and 

that each lavatory trash receptacle be equipped with a fire extinguisher that discharged 

automatically when a fire occurs in the receptacle. Amendment 25-74, April 1991, added 

§ 25.854 to require the same provisions for newly certificated airplanes.

 d. Policy/Compliance Methods. For guidance concerning lavatory smoke detection testing 

plus an acceptable smoke generator, see AC 25-9A. For compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Approval of Lavatory Smoke Detectors. The following policy is extracted from an 

FAA memorandum dated October 28, 1991, in response to an inquiry regarding certification of 

smoke detectors in transport category airplane lavatories. 

(a) The following guidance provides policy for determining approval criteria for 

the lavatory smoke detectors that are required by Amendment 25-74 to 14 CFR. This 

amendment added § 25.854, which requires the installation of lavatory smoke detectors on all 

transport category airplanes with a passenger capacity of 20 or more. The preamble to the 

amendment contains a statement that "a commercially available smoke detector, such as the type 

commonly used in residential buildings, which is demonstrated to serve its intended function as 

installed, could be considered adequate under the proposal of Notice 89-1." When the economic 

evaluation was performed in support of the rule change, the cost impact was based in part on 

costs associated with inexpensive smoke detectors such as those noted above. 

(b) When § 25.854 was added by Amendment 25-74, the justification used 

followed that contained in the preamble to the relevant part 121 amendment that was already in 

force. Section 121.308, effective March 29, 1985, requires lavatory smoke detectors for all 

airplanes operating under part 121. The original intent of the part 121 rule was to discourage 

passengers from smoking in the lavatories, and the smoke detectors were perceived as a costeffective way to provide that deterrent. This concern has become even greater now that virtually 

all US, and many international, commercial flights are "smoke free." The incentive to smoke in 

the lavatory is even greater because passengers cannot smoke at their seats. 

(c) The primary FAA concern is to detect smoke from a lavatory fire in a timely 

manner, and the most likely location for a fire in the lavatory is in the waste bin containing paper 

towels. For that reason, the smoke detectors currently in use in lavatories are tested using 

burning paper towels, and are not necessarily optimized for the particulate size contained in 

tobacco smoke. The detectors continue to be effective as a deterrent to smokers, however, 

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3/14/2000 AC 25-22 

because tobacco smoke is also detected, as evidenced by continuing reports of lavatory smoke 

alarms due to smoking in lavatories. To the extent that smoke detectors do not detect cigarette 

smoke, there is a perception that they are not properly designed and/or functioning. Therefore, it 

is the FAA position that industry be encouraged to do all they can to ensure that lavatory smoke 

detectors not only function well to detect smoke from lavatory fires, but also function to 

effectively deter lavatory smoking by efficiently detecting cigarette smoke as well. 

(d) When the airframe manufacturers were approached by their customers to

install smoke detectors as part of the type design, the FAA was requested to provide certification 

requirements covering the installation. The guidance provided was the following: 

 1 The detectors had to meet §§ 25.1301(a) and (d) and 25.1309(a). This 

meant that environmental, as well as performance, standards had to be considered. 

 2 Testing was to be performed in flight, and the combustible material used 

for testing was to be representative of what would be expected to burn in a lavatory waste bin, 

e.g., paper towels. 

 3 As a design goal, the detector was to provide a warning within one minute 

after a fire started. 

 4 If unpressurized flight was to be allowed, testing under those conditions 

had to be performed. 

(e) It has been suggested that Technical Standard Order (TSO) TSO-C1c might be 

appropriate for use in approving lavatory smoke detectors. TSO-C1c provides standards for 

approval of detectors to be used in cargo compartment smoke detection systems. The TSO 

requirements are not considered appropriate for lavatory smoke detection, as the environment 

and products of combustion to be detected are both different from that to be expected in a cargo 

compartment fire. This is not to imply, however, that a TSO-C1c detector would be 

unacceptable. With appropriate testing, and recognizing that the environment in a lavatory is 

different than a cargo compartment, a detector authorized under the TSO might be an acceptable 

choice. 

(f) TSO-C1c authorization should not be a requirement for lavatory smoke 

detector certification. The certification basis for the lavatory smoke detector installation must 

include §§ 25.1301 and 25.1309, as discussed above, plus any other appropriate sections such as 

those addressing structural or electrical supply requirements. If application for certification 

occurs after the effective date of Amendment 25-74 (April 15, 1991), § 25.854 would also be 

applicable. For these installations, the intended function is to detect smoke resulting from a fire 

occurring in a lavatory. 

(2) Approval of Lavatory Fire Extinguishers Containing Agents Other Than Halon. 

The following policy is extracted from an FAA memorandum dated March 31, 1997, to 

distribute a minimum performance standard for use in certifying lavatory fire extinguishers. 

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3/14/2000 AC 25-22 

(a) Sections 25.854(b) and 121.308(b) require that each lavatory on passengercarrying transport category airplanes with a passenger capacity of 20 or more, be equipped with a 

built-in fire extinguisher for each disposal receptacle for towels, paper, or waste located within 

the lavatory. The fire extinguisher must be designed to discharge automatically into each 

disposal receptacle upon occurrence of a fire in the receptacle. Currently, although not required 

by airworthiness regulations, the typical aircraft lavatory disposal receptacle fire extinguisher 

uses Halon as the fire extinguishing agent. 

(b) Halon production was banned as of January 1, 1994, under the provisions of 

the Montreal Protocol for those subscribing nations, due to its identification as an ozone 

destroying compound. The Environmental Protection Agency has exempted the aviation 

industry from the ban on the use of Halon. However, the FAA established the International 

Halon Replacement Working Group to help identify acceptable replacements for halons. A key 

aspect of this work is to define minimum performance standards which can be used to assess the 

performance of candidate replacement agents to assure that they will provide protection 

equivalent to Halon. Standards are being developed for fire extinguishers used in: lavatory trash 

receptacles; cargo compartments; engines and auxiliary power units; and hand held 

extinguishers. The first of these minimum performance standards is published in report number 

DOT/FAA/AR-96/122, titled “Development of a Minimum Performance Standard for Lavatory 

Trash Receptacle Automatic Fire Extinguishers,” dated February 1997. 

(c) The minimum performance standard in report DOT/FAA/AR-96/122 provides 

guidance on acceptable methods of compliance to § 25.854. The performance of an alternative 

agent needs to be measured against a standard test method. This document establishes fire load, 

trash disposal receptacle test article, test procedures, and pass/fail criteria for built-in 

extinguishers for lavatory disposal receptacles. 

 e. References. The addresses for ordering the latest revision of advisory circulars (AC), 

technical standard orders (TSO), and other referenced documents listed below can be found in 

the Appendix to this AC.

AC 25-9A, Smoke Detection, Penetration, and Evacuation Tests and Related Flight 

Manual Emergency Procedures. 

Technical Standard Order TSO-C1c, Cargo and Baggage Compartment Smoke 

Detection Instruments. 

FAA Report DOT/FAA/AR-96/122, Development of a Minimum Performance Standard 

for Lavatory Trash Receptacle Automatic Fire Extinguishers.

37. SECTION 25.855 - CARGO OR BAGGAGE COMPARTMENTS.

 a. Rule Text.

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For each cargo and baggage compartment not occupied by crew or passengers, 

the following apply: 

(a) The compartment must meet one of the class requirements of 25.857. 

(b) Class B through Class E cargo or baggage compartments, as defined in 

25.857, must have a liner, and the liner must be separate from (but may be 

attached to) the airplane structure. 

(c) Ceiling and sidewall liner panels of Class C compartments must meet the test 

requirements of Part III of appendix F of this Part or other approved equivalent 

methods. 

(d) All other materials used in the construction of the cargo or baggage 

compartment must meet the applicable test criteria prescribed in Part I of

appendix F of this Part or other approved equivalent methods. 

(e) No compartment may contain any controls, wiring, lines, equipment, or 

accessories whose damage or failure would affect safe operation, unless those 

items are protected so that- 

(1) They cannot be damaged by the movement of cargo in the compartment; and 

(2) Their breakage or failure will not create a fire hazard. 

(f) There must be means to prevent cargo or baggage from interfering with the 

functioning of the fire protective features of the compartment. 

(g) Sources of heat within the compartment must be shielded and insulated to 

prevent igniting the cargo or baggage. 

(h) Flight tests must be conducted to show compliance with the provisions of 

25.857 concerning- 

(1) Compartment accessibility, 

(2) The entries of hazardous quantities of smoke or extinguishing agent into 

compartments occupied by the crew or passengers, and 

(3) The dissipation of the extinguishing agent in Class C compartments. 

(i) During the above tests, it must be shown that no inadvertent operation of 

smoke or fire detectors in any compartment would occur as a result of fire 

contained in any other compartment, either during or after extinguishment, unless 

the extinguishing system floods each such compartment simultaneously. 

[Doc. No. 5066, 29 FR 18291. Dec. 24, 1964, as amended by Amdt. 25-15, 

32FR 13266, Sep. 20, 1967; Amdt. 25-32, 37 FR 3972, Feb. 24, 1972; Amdt. 25-

60, 51 FR 18242, May 16, 1986; Amdt. 25-72, 55 FR 29784, Jul. 20, 1990; Amdt. 

25-93, 63 FR 8048, Feb. 17, 1998]

b. Intent of Rule. This rule contains the material standards and design considerations for 

cargo compartment interiors; the statement that each cargo compartment must meet one of the 

class requirements of § 25.857; and the flight tests which must be conducted for certification. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The standards in 

section 4b.382 of the CAR, cargo and baggage compartments, and section 4b.384, proof of 

compliance, of the Civil Air Regulations (CAR), and were carried over essentially unchanged to 

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§ 25.855 of 14 CFR. This section contains the standards related to those requirements that must 

be demonstrated by flight test. 

 (1) Amendment 25-15 (September 20, 1967) added reference to the requirements of 

§ 25.853 for materials, rather than stating the materials must be at least flame resistant. 

 (2) Amendment 25-32 (February 24, 1972) specifically addressed insulation blankets 

and cargo covers meeting the requirements of § 25.853(b-3). 

 (3) Amendment 25-60 (May 16, 1986) added more stringent flammability standards for 

cargo compartment materials. 

 (4) Amendment 25-72 (July 20, 1990) made editorial changes and relocated the 

material flammability requirements to § 25.853 and Appendix F of part 25. 

 (5) Amendment 25-93 (February 17, 1998) revised §§ 25.855(c) and 25.857(c)(2) to 

eliminate reference to Class D cargo compartments. . These amendments upgraded the fire 

safety standards for cargo or baggage compartments in certain transport category airplanes by 

eliminating Class D compartments as an option for future type certification. Compartments that 

can no longer be designated as Class D must meet the standards for Class C or Class E 

compartments, as applicable. The Class D compartments in certain transport category airplanes 

manufactured under existing type certificates and used in passenger service must meet the fire or 

smoke detection and fire suppression standards for Class C compartments by early 2001 for use 

in air carrier, or most other commercial service. The Class D compartments in certain transport 

category airplanes manufactured under existing type certificates and used only for the carriage of 

cargo must also meet such standards or the corresponding standards for Class E compartments by 

that date for such service. These improved standards are adopted to increase protection from

possible in-flight fires. 

 (6) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group is recommending 

revisions to §§ 25.855 and 25.857, and a new advisory circular may evolve from that process. 

As a result of this process, Class B compartments may be revised and a new Class F 

compartment may be proposed. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preambles of this rule and the following information. In accordance with § 25.855(h), 

flight tests must be conducted to show compliance with the standards concerning: compartment 

accessibility, smoke and extinguishing agent penetration into occupied areas, and the dissipation 

of extinguishing agent. Cargo compartment smoke detection testing should be conducted in 

flight, and is usually accomplished concurrently with the smoke penetration tests. Advisory 

Circular 25-9A contains guidance regarding smoke detection and penetration testing. Where 

suitable data is available from previous certification flight test programs of the same airplane 

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model, the cognizant ACO may review an applicant’s proposal on a case by case basis and, if 

satisfactory, may accept certification ground tests for smoke detection purposes. 

 (1) Halon Concentration. The following policy is extracted from an FAA 

memorandum dated March 9, 1998, regarding the adequacy of Halon concentration levels 

needed for certification of cargo compartments in loaded vs. empty configurations. 

(a) Traditionally cargo fire extinguishing systems have been certified by 

demonstrating five percent initial Halon concentration levels and subsequent concentration levels 

of three percent, when the cargo compartment has been in the empty configuration. The majority 

of applicants in previous certification programs have not tested or analyzed a loaded cargo 

compartment configuration. 

(b) Recent tests and analyses have raised concern regarding the adequacy of Halon 

concentrations in a loaded cargo compartment. Regardless of this concern, the Transport 

Airplane Directorate does not currently have any written guidance, stating that applicants must 

test or analyze the loaded configuration. Requiring testing or analysis to show Halon 

concentration levels remain at or above three percent in the loaded cargo compartment may be 

considered beyond the scope of what the FAA can require without public comment. Although 

some applicants have previously considered loaded cargo compartments, we do not require 

testing or analysis for the loaded cargo condition at this time. 

NOTE: This may be considered a safety issue that the Transport Airplane Directorate may 

address in future policy. 

(2) Class D to Class C Conversion. The following guidance is extracted from an FAA 

memorandum dated October 29, 1997, regarding criteria for certification of cargo compartments 

to be converted from Class D to Class C. 

(a) On June 9, 1997, the FAA issued a Notice of Proposed Rulemaking, Docket 

No. 28937, Notice 97-10 (62 FR 32412, dated June 13, 1997), that proposes to upgrade the fire 

safety standards for cargo or baggage compartments in certain transport category airplanes by 

eliminating Class D compartments as an option for future certification. This notice also proposes 

that Class D compartments in certain transport category airplanes manufactured under existing 

type certificates and used in passenger service would have to meet the fire detection and 

suppression standards for Class C compartments by early 2001 for use in air carrier, commuter, 

on demand, or most other commercial service. 

(b) A policy memorandum, dated August 18, 1997, was issued in response to 

questions concerning what guidance was available regarding the certification of smoke

detection/penetration and fire suppression system evaluations in anticipation of the final rule 

associated with Notice 97-10. After considering industry objections to our memorandum of 

August 18, 1997, the information in our memorandum is revised to read: Fire Suppression Tests. 

Several companies have objected to the guidance provided under “Fire Suppression Tests” in the 

memorandum dated August 18, 1997, even though some recent certification projects have been 

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3/14/2000 AC 25-22 

approved using that guidance. Some manufacturers argued that they used a number of 

techniques to establish a minimum Halon concentration, including a volumetric averaging 

technique. The certification criteria contained under “Fire Suppression Tests” in our 

memorandum of August 18, 1997, is considered new policy for Class D to C conversions. 

Therefore, the manufacturers may continue to use the certification techniques they have been 

using, including volumetric averaging, to establish the minimum concentration of Halon for fire 

suppression. 

 (aa) At the FAA/Industry workshop of April 22-24, 1997, the FAA’s 

Technical Center expressed concern that the current Halon measuring technique using 

volumetric averaging may allow a concentration of Halon insufficient to suppress a fire. The 

FAA Technical Center tests have shown that Halon, having a higher density than air, settled in 

the cargo compartment. Further, the tests showed that fires may reignite at the higher water lines 

in the cargo compartment due to insufficient Halon concentration even though the average 

volumetric concentration of Halon was considered adequate. There was no subsequent measured 

increase in Halon concentration near the fire due to convective stirring. This information was 

presented to and discussed with Industry at the April 22-24,1997, workshop and is the basis of 

the “Fire Suppression Tests” guidance provided in our memorandum of August 18, 1997. 

(bb) The use of the technique of volumetric averaging to determine the 

minimum Halon concentration is questionable in light of the testing accomplished by the FAA’s 

Technical Center. Therefore, establishing minimum Halon concentrations near the ceiling 

should be considered. Applicants may elect to take advantage of this information in measuring 

the Halon concentration in their tests even though the FAA will not require this technique at this 

time. Furthermore, the Transport Standards Staff will develop an Advisory Circular (AC) that 

addresses measuring the minimum acceptable level of Halon in all cargo compartments. 

(c) The rest of the memorandum, dated August 18, 1997, was revised as follows, 

and may be used for guidance: 

(d) System Reliability. Use advisory material appropriate to the certification basis. 

When applying AC 25.1309-1A, the following is suggested: Detection and suppression systems 

are considered complex in terms of paragraph 6d of the AC. A failed detection system and/or a 

failed suppression system in conjunction with a fire should be considered a catastrophic event. 

Therefore, utilizing figure 2 of AC 25.1309-1A, knowing the system is complex and the failure 

event is a catastrophic event, the depth of analysis should include both a qualitative and 

quantitative assessment (reference paragraphs 8d, 9, and 10 of the AC). 

(e) Dispatch. For dispatch relief, the systems should be tested in the proposed 

Master Minimum Equipment List (MMEL) configurations. Dispatch may be allowed with 

detection or suppression systems inoperative in a cargo compartment provided the AFM 

prohibits the carriage of cargo in the affected compartment. 

 (f) Smoke Detection/Penetration/Evacuation Tests. Use AC 25- 9A, “Smoke

Detection, Penetration, and Evacuation Tests and Related Flight Manual Emergency 

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3/14/2000 AC 25-22 

Procedures,” supplemented by ANM-100 memorandum, dated June 18, 1997, and the referenced 

video. The supplemental memorandum elaborates on a small smoldering fire for use in detection 

tests. 

(g) If the applicant disagrees with any of the guidance above, then an issue paper 

can be created identifying the disagreement as an issue and coordinate it with the FAA. 

(3) Protection of Critical Equipment in Class E Cargo Compartments. The following 

guidance is extracted from an FAA Generic Issue Paper regarding protection of critical 

equipment in Class E cargo compartments. 

(a) Statement Of Issue. 

 1 A specific requirement to protect critical systems and equipment in 

Class E cargo compartments from fire damage does not exist. Both Class B and Class E cargo 

compartments are similar in that they do not have fire suppression systems. On Class B 

compartments, it is assumed that the fire can be controlled manually. On Class E compartments, 

the fire is controlled by shutting off ventilating airflow. In both types of cargo compartments 

fires can quickly reach dangerous proportions because no fast-acting suppression system is 

installed. In a Class E compartment, an uncontrolled fire could damage critical systems and 

equipment to compromise flight safety before ventilating airflow could be effectively shut off. 

For this reason, protection of these critical systems and equipment must be ensured. 

 (b) Background. 

 1 The FAA has issued Airworthiness Directives (AD) 91-10-02, which 

requires certain design, equipment, and operational changes to maximize fire protection on 

"Combi" Class B cargo compartments. One of these requirements is to ensure appropriate 

protection of cockpit voice and flight data recorders, windows, primary flight control systems 

(unless it can be shown that a fire could not cause jamming or loss of control), and other 

equipment within the compartment that is required for safe flight and landing. If protective 

covers are used, they must be constructed of materials that meet the flame penetration resistance 

requirements of § 25, 14 CFR, appendix F, part III [Amendment 25-60]. 

 2 The generic model airplane main deck Class E cargo compartment is 

similar to the main deck Class B cargo compartment on generic model "Combis." Therefore, 

similar protection must be afforded to the affected systems on the generic model aircraft. 

 

(c) FAA Position. 

 1 Section 21.21(b)(2) of 14 CFR requires that no feature or characteristic of 

an airplane make it unsafe for the category in which certification is requested. Based on the 

similarities of the generic model aircraft Class E cargo compartment to the generic model aircraft 

"Combi" Class B cargo compartment, a concern exists that the unsafe conditions identified on 

"Combis" will also exist on the generic model aircraft. 

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 2 Prior to issuance of the Generic Model Airplane Type Certificate, it must 

be shown by design or analysis that adequate fire protection has been provided to the cockpit 

voice and flight data recorders, windows, primary flight controls (unless it can be shown that a 

fire could not cause jamming or loss of control), and other systems and equipment within the 

compartment that is required for safe flight and landing. If protective covers are used, they must 

be constructed of materials that meet the flame penetration resistance requirements of 14 CFR 

part 25, appendix F, part III [Amendment 25-60]. 

 e. Reference. The address for ordering the latest revision of advisory circulars and other 

referenced documents listed below can be found in the Appendix to this AC.

AC 25-9A; Smoke Detection, Penetration, and Evacuation Tests and Related Flight 

Manual Emergency Procedures. 

38. SECTION 25.857 - CARGO COMPARTMENT CLASSIFICATION.

 a. Rule Text.

(a) Class A. A Class A cargo or baggage compartment is one in which- 

(1) The presence of a fire would be easily discovered by a crewmember while at 

his station; and 

(2) Each part of the compartment is easily accessible in flight. 

(b) Class B. A Class B cargo or baggage compartment is one in which- 

(1) There is sufficient access in flight to enable a crewmember to effectively reach 

any part of the compartment with the contents of a hand fire extinguisher; 

(2) When the access provisions are being used, no hazardous quantity of smoke, 

flames, or extinguishing agent, will enter any compartment occupied by the crew 

or passengers; 

(3) There is a separate approved smoke detector or fire detector system to give 

warning at the pilot or flight engineer station. 

(c) Class C. A Class C cargo or baggage compartment is one not meeting the 

requirements for either a Class A or B compartment but in which- 

(1) There is a separate approved smoke detector or fire detector system to give 

warning at the pilot or flight engineer station; 

(2) There is an approved built-in fire-extinguishing or suppression system 

controllable from the cockpit; 

(3) There are means to exclude hazardous quantities of smoke, flames, or 

extinguishing agent, from any compartment occupied by the crew or passengers; 

(4) There are means to control ventilation and drafts within the compartment so 

that the extinguishing agent used can control any fire that may start within the 

compartment. 

(d) [Reserved]

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(e) Class E. A Class E cargo compartment is one on airplanes used only for the 

carriage of cargo and in which- 

(1) (Reserved)

(2) There is a separate approved smoke or fire detector system to give warning at 

the pilot or flight engineer station; 

(3) There are means to shut off the ventilating airflow to, or within, the 

compartment, and the controls for these means are accessible to the flight crew in 

the crew compartment; 

(4) There are means to exclude hazardous quantities of smoke, flames, or noxious 

gases, from the flight crew compartment; and 

(5) The required crew emergency exits are accessible under any cargo loading 

condition. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-32, 

37 FR 3972, Feb. 24, 1972; Amdt. 25-60, 51 FR 18243, May 16, 1986; Amdt. 25-

93, 63 FR 8048, Feb. 17, 1998]

b. Intent of Rule. This rule provides the standards for the various classes of transport 

category airplane cargo compartments. It does not cover closets and stowage compartments (See 

§ 25.787). Cargo compartment liner standards and required flight tests to demonstrate 

compliance are addressed in § 25.855. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in section 4b.383 of the CAR were carried over essentially unchanged to § 25.857 of 14 CFR. 

 (1) Amendment 25-32 (February 24, 1972) moved the liner requirements for Classes B 

through E to § 25.855. 

 (2) Amendment 25-60 (May 16, 1086) limited the volume of Class D compartments to 

a maximum of 1,000 cubic feet. 

 (3) Amendment 25-93 (February 17, 1998) revised paragraph (c)(2) to add reference to 

suppression systems as well as extinguishing systems, and eliminated reference to Class D 

compartments and reserved paragraph (d). Also see the explanation under § 25.855, Amendment

25-93. 

 (4) Harmonization. This regulation is the subject of a Federal Aviation 

Regulation/Joint Aviation Requirement (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group has provided the FAA 

with a fully harmonized draft notice of proposed rulemaking (NPRM) and an accompanying 

draft proposed advisory circular (AC) for regulatory evaluation. The NPRM and the AC provide 

the criteria for corrective action(s) that, if implemented, would be acceptable to the FAA for 

showing compliance with the airworthiness directive AD 93-07-15, or any new "combi" designs 

for Class B or the new proposed Class F cargo compartments. The NPRM proposes to revise 

§§ 25.855 and 25.857, and the proposed AC provides one means but not the only means of 

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compliance. Should an applicant not desire to comply with the requirements for Class B or class 

F compartments as stipulated above, then the main deck cargo compartments should comply with 

requirements for a Class C compartment as defined in current §§ 25.855 and 25.857. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preambles of this rule and the following information. 

(1) Main Deck Cargo Compartment Fire Protection Certification Procedures. The 

following policy is extracted from an FAA letter dated June 6, 1997, and addresses requirements 

for “combi” airplanes. The letter is in response to a request from an applicant for information 

regarding certification of “combi” airplanes. The FAA refers to airplanes that can carry both 

passengers and cargo on the main deck as "combi's." The guidance is based on proposed 

rulemaking and advisory material being developed through the ARAC process. 

(a) Information was requested concerning Federal Aviation Administration (FAA) 

requirements for certification of “combi” airplanes. The minimum design requirements for 

issuance of a Type Certificate for large transport airplanes are identified in 14 CFR part 25. The 

FAA was specifically asked to identify the criteria against which the FAA would evaluate the 

main deck fire protection features for FAA certification. 

(b) The passenger and cargo compartments are separated by a partition that is

either fixed or movable to accommodate various passenger/cargo combinations. Some

“combi’s” can also accommodate all passenger or all cargo configurations. In most cases, the 

main deck cargo compartments on these airplanes are identified as Class B cargo compartments, 

which meet the fire protection requirements specified in § 25.857(b). Class B cargo 

compartments rely on manual firefighting as the primary method of controlling fires. 

(c) In 1987, a South African Airways 747-200 “combi” crashed in the Indian 

Ocean following a fire in it’s main deck Class B cargo compartment. Following the 

investigation, the FAA concluded that manual firefighting in large Class B cargo compartments 

was not effective, and that the certification requirements identified in § 25.857(b) were 

inadequate. As a result, the FAA issued Airworthiness Directive (AD) 93-07-15, which required 

operators of existing large “combi” airplanes manufactured by Boeing or McDonnell Douglas to 

make significant modifications to their airplanes and operations. Operators were offered four 

options: (1) convert the Class B cargo compartment to Class C, (2) carry all cargo in the Class B 

cargo compartment in containers that meet the requirements for Class C cargo compartments, 

(3) use fire containment covers or containers for all main deck cargo, along with other airplane 

and operational changes, or (4) incorporate a 90 minute duration fire suppression system, along 

with other airplane and operational changes. 

(d) To address new “combi” designs, the FAA, through the Aviation Rulemaking 

Advisory Committee (ARAC) process, is proposing a revision to §§ 25.855 and 25.857 which 

would require Class B cargo compartments to be small enough that a crewmember with a hand 

held fire extinguisher can extinguish a fire anywhere within the compartment without entering it. 

The FAA is also proposing a new cargo compartment classification, Class F. The proposed 

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Class F compartment would not be size-limited, and would require incorporation of a means to 

extinguish or control the fire without the need for a crewmember to enter the compartment. This 

includes not reaching in to move something in order to have full access to the fire extinguisher. 

(e) The FAA has identified the criteria against which the FAA would evaluate 

main deck fire protection features on new “combi’s.” Although AD 93-07-15 provides an 

adequate level of safety for existing airplanes, the FAA recommends that all new “combi’s” meet 

the intent of the proposed changes to §§ 25.855 and 25.857. In addition, the effectiveness of the 

third and fourth options of AD 93-07-15 depends heavily on the FAA-monitored procedures, 

maintenance, and training at individual airlines, whereas the proposed part 25 changes rely on 

the airplane design, rather than airline operations. The FAA recognizes that the limited size of 

the proposed Class B cargo compartment would not meet operational needs. The FAA 

recommends that the main deck cargo compartment on the “combi” meet either the requirements 

for Class C cargo compartments identified in §§ 25.855 and 25.857, or the proposed 

requirements for Class F cargo compartments. The FAA and the ARAC are proposing that Class 

F cargo compartments meet the general cargo compartment requirements of § 25.855, and the 

following: 

 1 A Class F cargo compartment must have a liner that is separate from, but 

may be attached to, the airplane structure, unless it can be shown that other means for containing 

the fire and protecting critical systems and structure are provided. 

 2 The ceiling and sidewall panels of the liner in the Class F cargo 

compartment, if required, must meet the test requirements of part 25, Appendix F, part III, or

other FAA-approved equivalent methods. This includes windows and window shades/covers in 

combi airplanes. 

 3 A means to extinguish or control the fire in the Class F cargo 

compartment, without requiring a crewmember to enter the compartment, must be provided. 

 4 A means to prevent hazardous amounts of smoke, flames, or suppression 

agent in the Class F cargo compartment from entering any occupied compartment must be 

provided. 

 5 A separate, FAA-approved smoke or fire detector system to alert the pilot 

or flight engineer station in the event of a fire must be provided. 

(2) Class A Cargo Compartments: The following policy is extracted from an FAA 

memorandum dated February 8, 1996, to provide FAA policy regarding certification of 

passenger airplanes converted to all-cargo Class A or unclassified cargo compartment 

configurations. 

(a) The FAA Transport Airplane Directorate has become aware of a number of 

transport category airplanes that have been reconfigured as all-cargo airplanes with cargo 

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compartments classified as Class A, or with the cargo compartment classification not addressed 

at all. 

(b) An FAA memorandum , dated September 13, 1988, states that " . . . the agency 

policy has always been, and still is, that cargo compartments larger than that stated in Order 

8110.27A(5)(b) are to be classified as Class B through E, whichever is applicable, and that 

operators comply with all requirements relative to each class." The aforementioned order 

describes the Class A compartment as " . . . small open compartments used for storage of crew 

luggage and located in the cockpit area where a fire can be easily discovered by a crewmember." 

The order goes on to state that: "During the 74/75 Airworthiness Review, it was mentioned that 

full cabins or other large cargo compartments were presented for approval under Class A 

category, and that these compartments were consistently rejected on the basis that their volume

was outside the intent of the Class A category where a fire must be rapidly detected and 

extinguished. Since the Class A compartment has no liner, large cargo areas have been 

considered to be outside the intent of the Class A category. It was recommended to limit the 

volume to 200 cubic feet." 

(c) Advisory Circular (AC) 25-18, issued January 6, 1994, Transport Category 

Airplanes Modified for Cargo Service, addresses this subject. In the AC, Class A compartments 

are defined as follows: 

 1 "A Class A compartment is one that is located so close to the station of a 

crewmember that the crewmember would discover the presence of a fire immediately. In 

addition, each part of the compartment is easily accessible so that the crewmember could quickly 

extinguish a fire with a portable fire extinguisher. A Class A compartment is not required to 

have a liner. 

 2 "Typically, a Class A compartment is a small open compartment in the 

cockpit area used for storage of crew luggage. A Class A compartment is not, however, limited 

to such use; it may be located in the passenger cabin and used for other purposes provided it is 

close to a crewmember's station. Typically, the crewmember would be a member of the 

flightcrew; however, the compartment could be located adjacent to the station of any other 

crewmember. 

 3 "Because a Class A compartment does not have a liner, it is absolutely 

essential that the compartment be small and located close enough to a crewmember that any fire 

that might occur could be discovered and extinguished immediately. Without a liner to contain 

it, an undetected or uncontrolled fire could quickly become catastrophic by burning out of the 

compartment and spreading throughout the airplane. There is no specific limit on the volume; 

however, all portions of the compartment must be virtually within arms length of the 

crewmember in order for any fire to be detected immediately and extinguished in a timely 

manner. Although there may be some exceptions, such as a 'U-Shaped' compartment for 

example, a Class A compartment greater than 50 cubic feet in volume would not typically have 

the accessibility required by § 25.857(a)(2) for fighting a fire." 

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(d) Advisory Circular 25-17, Transport Airplane Cabin Interiors Crashworthiness 

Handbook, dated July 15, 1991, or latest revision, notes that a Class A compartment was 

envisioned as a small, open compartment located in the cockpit area. 

(e) It is clear from reviewing the above policy, and noting the chronology, that the 

intent of Class A compartments is not consistent with the conversion of an entire passenger cabin 

to a cargo compartment and identifying it as Class A. A number of issues support this view. In 

order to exit the airplane in the event of an uncontrollable fire, the only path may be through the 

fire zone. Access to emergency exits often does not exist with the compartment fully loaded. 

Further, there are no means to protect the flightcrew from the effects of toxic gasses that can be 

generated in a smoldering fire. In cargo compartments classified Class B through Class E, the 

cargo and crew areas are separated and certification tests are conducted to ensure that no 

hazardous quantities of smoke or fumes will penetrate occupied areas, including the cockpit. In 

addition, other factors that are frequently not addressed in these modifications are: lack of a 

cargo restraint system, compartment liners, and a second emergency exit. 

(f) The FAA has determined, for the reasons discussed above, that it is 

inappropriate to classify full cabin or other large areas as Class A compartments. For future 

certification projects (whether new, amended, or supplemental), the above guidance must be 

considered. 

(3) Class E Cargo Compartments. The following policy is extracted from an FAA 

memorandum dated March 21, 1991, in response to a request for guidance on compliance 

relating to the installation of a Class E cargo compartment. 

(a) Smoke detection capability must be demonstrated in any flight condition 

contemplated by the operator. This usually means a flight test must be performed at the 

maximum contemplated cabin altitude. 

(b) Smoke penetration from Class E compartment into cockpit tests must be 

conducted starting in a cruise configuration and continuing into a descent to sea level landing. If 

any other procedures, such as attempting to suppress the fire by going to higher altitudes, involve 

other operating altitudes, these must also be addressed. 

(c) An Airplane Flight Manual (AFM) supplement is needed which addresses crew 

action following smoke annunciation, i.e., crew shuts off ventilating air to the cargo 

compartment, dons oxygen equipment, and proceeds to the nearest adequate airport. 

(d) The protective breathing equipment requirement in § 25.1439(a) must be 

addressed. 

(4) Access to Class E Cargo Compartments. The following policy is extracted from an 

FAA letter dated July 3, 1990, in response to a request for guidance on compliance on the 

carriage of dangerous goods, and the need for access, in Class E compartments. 

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(a) The International Civil Aviation Organization regulations for hazardous goods 

carried in Class E compartments require adequate access to these goods. Section 25.857 of 14 

CFR does not preclude access to Class E compartments in flight. Sections 25.851(a)(3) and 

121.309(c)(2) require that one hand fire extinguisher must be provided for use in each Class E 

cargo compartment that is accessible to crewmembers in flight. 

(b) Although part 25 does not require in-flight access to Class E cargo 

compartments, in-flight access to the compartment can be accommodated with proper design if 

operational requirements result in the carriage of hazardous goods in a Class E compartment. 

(5) Freighter Aircraft with Accessible Class E Compartment. The following is 

extracted from an issue paper dated March 10, 1995, titled “Accessible Class E Cargo 

Compartment.” 

(a) The following applies to certification of a freighter aircraft with an accessible 

Class E Cargo Compartment, which would allow the operator to carry hazardous material. The 

following would allow the crew and "persons" access to the entire cargo compartment. 

(b) Access by one cargo handler will be allowed during flight to a Class E cargo 

compartment carrying hazardous material provided that one Portable Breathing Equipment 

(PBE) meeting TSO C-116 is available in the flight deck. The cargo handler in the Class E cargo 

compartment must carry the portable oxygen at all times while in the cargo compartment. A 

placard must state that the portable oxygen must be carried at all times. In addition, there must 

be a readily detectable means to alert the cargo handler to don oxygen equipment and/or return to 

the flight deck. Further, two Halon fire extinguishers of adequate size must be installed near the 

cargo compartment entrance with procedures for their use. Finally, there must be a means to 

prohibit smoking. Compliance with the above will meet §§ 25.855(e)(1), (e)(2), and 

25.857(e)(4), taking into consideration the opening caused by a person exiting the cargo 

compartment. 

(c) The applicant should show by test or by analysis that with the access door open 

in flight the positive ventilation with only one pack operating is capable of excluding smoke, 

flames, or noxious gasses in compliance with § 25.857(e)(4). In addition, the additional fire 

extinguisher that was kept on the flight deck was required to be a 16-pound Halon extinguisher 

to comply with § 121.309(c)(2). 

 e. References. The addresses for ordering the latest revision of referenced documents 

listed below can be found in the Appendix to this AC.

Advisory Circular, 25-XX, Cargo Compartment Fire Extinguishing or Suppression 

Systems, date TBD. 

Advisory Circular, 25-XX, Class B and F Cargo Compartments, date TBD. 

Two methods of determining the leakage rate for Class D compartments that have been 

approved for certification are contained in: 

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3/14/2000 AC 25-22 

Appendix I of CAA Technical Development Report No. 146, Evaluation of Flight 

Fire Protection Means for Inaccessible Aircraft Baggage Compartments, dated June 1951, and 

FAA Technical Center Report DOT/FAA/CT-83/1, Analysis of Dissipation of 

Gaseous Extinguisher Agents in Ventilated Compartments, dated May 1983. 

39. SECTION 25.858 - CARGO OR BAGGAGE COMPARTMENT SMOKE OR FIRE 

DETECTION SYSTEMS.

 a. Rule Text.

If certification with cargo or baggage compartment smoke or fire detection 

provisions is requested, the following must be met for each cargo or baggage 

compartment with those provisions: 

(a) The detection system must provide a visual indication to the flight crew within 

one minute after the start of a fire. 

(b) The system must be capable of detecting a fire at a temperature significantly 

below that at which the structural integrity of the airplane is substantially 

decreased. 

(c) There must be means to allow the crew to check in flight, the functioning of 

each fire detector circuit. 

(d) The effectiveness of the detection system must be shown for all approved 

operating configurations and conditions. 

[Amdt. 25-54, 45 FR 60173, Sep. 11, 1980; Amdt. 25-93 63 FR 8032, Feb. 17, 

1998]

b. Intent of Rule. This rule provides minimum design and certification requirements for 

cargo or baggage compartment fire or smoke detection systems. This rule is intended to provide 

a standard that a fire be detected and indicated to the crew in less than one minute after 

inception. 

 c. Background. This section was added to part 25 of the Code of Federal Regulations (14 

CFR) with Amendment 25-54. Prior to Amendment 25-54, a detection limit of no more than five 

minutes was established by policy since there was no regulatory reference to detection time. 

Section 25.858(d) requires detection testing in any configuration proposed by the applicant, 

including depressurized flight. The cognizant ACO may certify minor changes by analysis of 

similarities to previously certified configurations, laboratory test, or ground test. 

 (1) Amendment 25-93 (February 17, 1998) revised the section heading and 

introductory paragraph to add reference to baggage compartments and smoke detection systems. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

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3/14/2000 AC 25-22 

(1) Smoke Detection Certification Testing. The following policy is extracted from an 

FAA memorandum dated June 18, 1997, which was developed in response to inquiries regarding 

smoke quantities to certify smoke detection systems in cargo compartments. The memo was 

used to distribute video guidance to help clarify smoke quantities specified in Advisory Circular

(AC) 25-9A. 

(a) Section 25.858(a) requires that a cargo compartment fire detection system provide a 

visual indication to the flightcrew within one minute after the start of a fire. AC 25-9A, section 10, 

provides guidelines for the conduct of certification tests related to smoke detection. The AC states 

that systems which provide a warning within one minute from the start of smoke generation are 

considered to be in compliance with the requirements of § 25.858 for cargo compartments. 

Paragraph 10.b. states that the objective of the smoke detection test is “to demonstrate that the smoke 

detection system installation will detect a smoldering fire producing a small amount of smoke.” 

(b) In February 1997, FAA/Industry meetings were held to discuss how manufacturers 

and operators intended to support the part 25 and 121 rule changes requiring fire detection and 

suppression systems to be installed in existing Class D cargo compartments. During the meeting, 

Industry expressed concern that a large number of conversions were going to be made and several 

Aircraft Certification Offices (ACO’s) would be impacted. Industry was concerned that the smoke 

detection standards used by the ACO’s varied and requested the FAA’s effort to standardize the test 

criteria used to satisfy § 25.858. 

(c) In response to this request, the FAA held a standardization workshop at the FAA 

Technical Center in Atlantic City, NJ, during April 1997. ACO personnel and designated 

engineering representatives (DER’s) attended to discuss smoke generation and quantities, and other 

certification requirements relevant to cargo compartments requiring both detection and suppression 

systems. 

(d) During the standardization workshop, the Technical Center provided a 

demonstration of a typical smoke quantity from a smoldering fire by igniting rags inside a suitcase. 

There was a consensus agreement among the ACO members present that this suitcase demonstration 

generated an appropriate quantity of smoke for a one minute detection system. It was agreed that a 

video of the burning suitcase should be distributed to the ACO’s as an aid to visually demonstrate the 

amount of smoke that is representative of smoke from a smoldering fire as is mentioned in AC 

25-9A. This video is considered visual guidance demonstrating a typical amount of smoke from a

smoldering fire. The amount of smoke demonstrated in this video helps to clarify the objective of 

paragraph 10.b. in AC 25-9A. 

(e) Copies of the video mentioned above are available through the FAA William J. 

Hughes Technical Center. Refer to the Appendix of this AC for the mailing address.

(f) Additional items of discussion at the workshop included the following: 

 1 The tests conducted at the FAA Technical Center have shown that some

materials that are routinely carried in cargo compartments burn very quickly with open flames 

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3/14/2000 AC 25-22 

and little smoke. It was shown that a fire internal to a suitcase can rupture an aerosol can 

(containing butane, propane, or isobutane) contained therein and cause the liner to be breached. 

Once the liner is breached, ventilation control and ability to maintain adequate Halon 

concentration will be lost. It is inappropriate to rely on detection systems that depend on copious 

quantities of smoke prior to detection. A fire does not know what volume compartment it is in. 

A smoldering fire will generate the same levels of smoke regardless of the compartment size and 

should be detected within the required time to detection (one minute or five minutes as 

applicable) in any compartment. The solution to larger compartments may be an increased 

number of detectors, increased sensitivity of detectors, different type of detectors (fanned or 

ionic), or infrared detectors. 

 2 All of the current smoke detection limits and guidance have been 

subjective and based on previous certification practices. The FAA Technical Center's planned 

test program includes defining fire hazards, the amount of heat generated, and the typical 

quantity (in more objective terms) of smoke generated, associated with typical cargo fire 

hazards. The suitcase video is intended to provide clearer guidance until the Technical Center 

test program is complete. It is a continuing step toward standardizing the approach to smoke 

detection testing until the FAA determines usable, objective criteria. 

(2) Certification Testing for Smoke Detection. The following policy is extracted from

an FAA memorandum dated February 11, 1993, in response to an inquiry regarding certification 

of smoke detection systems in the cargo compartments of transport category airplanes, and 

provides guidance for determining detection times for these systems. 

(a) Advisory Circular (AC) 25-9 and AC 25-17, differ in the guidance provided for 

time to detection for certification testing. AC 25-9 states in paragraph 7.e.(3) that "smoke 

detection should occur within one minute after the start of smoke generation." AC 25-17 states 

in paragraph 671.b.(6) that "An acceptable detection time for smoke detectors is 5 minutes. Use 

the smoke quantity and location criteria of AC 25-9, 'Smoke Detection, Penetration, and 

Evacuation Tests, and Related Flight Manual Emergency Procedures,' for showing that the 

smoke detection system detects a fire in satisfactory time. The time for fire detection systems 

was changed to one minute by Amendment 25-54 in § 25.858." 

(b) The guidance contained in AC 25-9 reflects the detection times required in 

Amendment 25-54, which added § 25.858. This section established a higher standard of safety 

than previously existed for transport category airplanes, and applies to airplanes which have 

Amendment 25-54 in their certification basis. When an applicant applies for an amended or 

supplemental type certificate to convert a passenger configuration to a cargo configuration on an 

older airplane whose certification basis predates Amendment 25-54, the allowable maximum 

detection time is not specifically stated in 14 CFR. The five minute detection time, which is not 

mentioned in AC 25-9, was established in an FAA letter in 1965, and has been the accepted 

maximum detection time requirement until § 25.858 became effective on September 11, 1980. 

(c) If an applicant applies for an amended or supplemental type certificate and the 

certification basis for the airplane predates Amendment 25-54, there is no regulatory basis for 

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3/14/2000 AC 25-22 

requiring a detection time of one minute. The applicant should be encouraged to meet the later 

amendment, or to make detection within one minute a design goal, but it would be inappropriate 

to require that the applicant comply with § 25.858. 

(d) Two additional points should be made regarding wording in regulations that 

address fire detection systems. When the terms "fire detector" or "fire detection system" are 

used, this also encompasses smoke detectors and smoke detection systems. A review of the 

preamble to Amendment 25-54 reveals that the rule is intended to address systems which detect

fires, and smoke detection systems were included. Also, questions are frequently asked 

regarding the expression that detection must occur "within one minute after the start of a fire." 

As noted in AC 25-9, time to detection is measured "after the start of smoke generation." 

Therefore, detection time is defined as the elapsed time from when the smoke generator is 

activated (switched to the mode which actively generates smoke) to when detection occurs 

(annunciation of smoke detection to the flight crew). 

(3) Smoke Detection Certification Testing. The following policy is extracted from two 

FAA letters dated December 3 and December 15, 1992. The letters were in response to a request 

for approval of five items discussed in a meeting between the FAA and an applicant on

November 6, 1992, all related to "Smoke Detection Testing Methodology/Procedures For 

Compliance To § 25.858." 

(a) The five items below were discussed individually. It was noted prior to 

responding that the positions provided in the letter related to demonstration of compliance with 

the provisions of § 25.858, as promulgated by Amendment 25-54, and were not intended to 

provide guidance as to methods to be used for certification of an airplane which does not have 

Amendment 25-54 in its certification basis. In addition, the responses provided were not 

intended to be all inclusive. Other methods of demonstrating compliance could also be 

acceptable. 

(b) The first proposal related to a specific fuel load (using tobacco as the fuel),

tamping, and lighting procedure to be used for generating smoke to show compliance with the 

FAA requirements. The procedure and the tobacco weights were proposed to meet the 

requirements in a large compartment such as in a large transport category all cargo airplane. 

There were two charges (basket sizes) proposed. These two charges should be acceptable for all 

large transport category airplane cargo compartments. The FAA reserves the option to stipulate 

smaller fuel loads when finding compliance in small compartments. 

(c) The second proposal discussed the use of a smoke deflector hat, which is 

required to provide a more realistic smoke propagation pattern than a direct smoke exhaust 

directed toward the crown of the compartment. The FAA had seen no evidence that a direct 

smoke stack exhaust provided a representative smoke pattern. It was stated that the deflector hat 

would continue to be required in future tests using a tobacco generator. 

(d) The third proposal requested approval of continuous smoke generation until 

detection, and allowed "averaging" the detection times to give a one minute overall average for 

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3/14/2000 AC 25-22 

all positions tested, with maximum time not to exceed 110 seconds at any location. 

Section 25.858 states that detection must occur within one minute after the start of a fire. When 

§ 25.858 was proposed, several commenters stated that the one minute detection time was too 

stringent, especially in light of the existing policy (5 minutes to detection). It was the opinion of 

the commenters that the technology was not available to enable the airplane manufacturers to 

meet the one minute requirement. The preamble to the final rule, issued as part of Amendment 

25-54, noted that the technology was available or would be available when new airplanes 

affected by the rule were certificated. Subsequently, a number of airplanes have shown 

compliance with the one minute rule, indicating that the promulgators of the rule were correct in 

their assessment. The technology is available, and the FAA could not arbitrarily change that 

requirement to allow a fire to go undetected for 110 seconds in some locations in the cargo 

compartment. Such a finding would not be in accordance with the rule as issued. A rule change 

to allow longer detection times in special cases, such as for very large cargo compartments on 

all-cargo airplanes, would be required. 

(e) The fourth proposal dealt with a means to reduce the number of locations at 

which smoke would be generated within the compartment to diminish the costs associated with

flight testing new cargo compartment configurations. The proposal stated that smoke detection 

locations would be representative of high, low, and normal detection locations as documented in 

the applicant's tests. The certification locations were to be randomly selected from a population 

of representative locations. The number of representative test locations must be at least twice the

number of required certification test locations and they must be available for FAA review. 

Typically, laboratory and ground testing determine the worst locations in each part of the 

compartment. The random selection of flight test locations should be acceptable provided the 

whole population of ground test times is already compliant with the rule. The FAA determined 

that this approach was acceptable for demonstrating compliance with § 25.858, and also for 

testing on the applicant's airplane. 

(f) The fifth proposal discussed the "Statistical Procedure and Methodology" for 

demonstrating compliance with § 25.858. The proposal to use an average detection time for all 

locations, with no tested location having a detection time exceeding 110 seconds, was not 

acceptable to the FAA. However, it was recognized that smoke testing is not a perfectly 

repeatable process, and testing at any one location may yield varying detection times with 

multiple tests. Previous certification tests have proceeded as follows: If the first test at a given 

location results in a detection time of greater than 60 seconds, the applicant must repeat the test 

two more times; if the resulting average of the three tests at the given location is 60 seconds or 

less, the given location is a “Pass;” if the resulting average of the three tests at the given location 

is greater than 60 seconds, the given location and the fire detection system are both “Failed” and 

require a redesign of the system. 

(g) Two additional points were raised by the applicant. In the FAA response to the 

second proposal, paragraph (c) above, regarding the use of a smoke deflector hat on the smoke 

generator, the FAA noted that "...we will continue to require the deflector hat on future tests." 

The applicant was concerned that this statement could lead to a misunderstanding regarding the 

use of the deflector hat on previous tests. The FAA had commented in previous correspondence 

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3/14/2000 AC 25-22 

that the smoke generator tends to direct smoke toward the top of the compartment, which could 

result in earlier detection times than would be the case if a cargo fire actually occurred (the 

objective of the generated smoke is to simulate natural bouyancy, not necessarily the upward 

velocity). It was noted that there had been no certification tests conducted using the applicant's 

smoke generator with a deflector hat in place. However, a deflector hat was informally tested 

during a certification flight test, and the FAA decided to require the use of a smoke deflector for 

the tobacco generator for any future certification flight tests that involved demonstrating 

compliance with the provisions of § 25.858. 

(h) The FAA noted in our response to the third proposal, paragraph (d) above, 

regarding approval of continuous smoke generation until detection and allowing "averaging" the 

detection times to give a one minute overall average for all positions tested, that " Such a finding

would not be in accordance with the rule as issued." The applicant expressed the opinion that the 

§ 25.858 wording "...within one minute after the start of a fire" was not adequately defined in the 

rule. Demonstration of compliance for smoke detection systems, whether the requirement is the 

previous 5 minute standard per policy or the 1 minute provision in § 25.858, has always been 

accomplished by measuring the time from smoke generator "ON" until detection. We did not 

anticipate any change in this policy. If this were not the case, the period of time during which 

the tobacco charge is being ignited with a torch would be included in the time for detection. For 

the tobacco generator, the "ON" time is when the tobacco basket is placed in the generator and 

the fan turned on. For a theatrical smoke generator, the "ON" time is when the switch is turned 

on. It is recognized that a theatrical smoke generator may have a built in time delay (by design) 

between the "switch on" event and when the smoke begins to exit; the cognizant ACO engineer 

needs to make that determination. Such a time differential may be negligible or not measurable. 

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3/14/2000 AC 25-22 

(4) William J. Hughes Technical Center Research and Development. The FAA 

Technical Center is currently conducting a research and development (R&D) program to help 

standardize fire detection certification methods. When this R&D effort is complete, it is 

intended to revise AC 25-9A to provide more objective criteria for smoke/fire detection 

certification. The above Policy/Compliance Methods will be subject to change and/or

incorporation into the revision to AC 25-9A in the future. 

 e. References. The addresses for ordering the latest revision of advisory circulars and 

other referenced documents listed below can be found in the Appendix to this AC.

AC 25-9A - Smoke Detection, Penetration, and Evacuation tests and related flight 

manual Emergency Procedures. 

AC 25-17 - Transport Airplane Cabin Interiors Crashworthiness Handbook

FAA Video, Smoke Quantities to Certify Smoke Detection Systems in Cargo Areas. 

40. SECTION 25.863 - FLAMMABLE FLUID FIRE PROTECTION.

 a. Rule Text.

(a) In each area where flammable fluids or vapors might escape by leakage of a 

fluid system, there must be means to minimize the probability of ignition of the 

fluids and vapors, and the resultant hazards if ignition does occur. 

(b) Compliance with paragraph (a) of this section must be shown by analysis or 

tests, and the following factors must be considered: 

(1) Possible sources and paths of fluid leakage, and means of detecting leakage. 

(2) Flammability characteristics of fluids, including effects of any combustible or 

absorbing materials. 

(3) Possible ignition sources, including electrical faults, overheating of 

equipment, and malfunctioning of protective devices. 

(4) Means available for controlling or extinguishing a fire, such as stopping flow 

of fluids, shutting down equipment, fireproof containment, or use of extinguishing 

agents. 

(5) Ability of airplane components that are critical to safety of flight to withstand 

fire and heat. 

(c) If action by the flight crew is required to prevent or counteract a fluid fire 

(e.g., equipment shutdown or actuation of a fire extinguisher) quick acting means 

must be provided to alert the crew. 

(d) Each area where flammable fluids or vapors might escape by leakage of a 

fluid system must be identified and defined. 

[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970, as amended by Amdt. 25-46, 

43 FR 50597, Oct. 30, 1978] 

b. Intent of Rule. The intent of this rule is to minimize the probability of ignition of 

flammable fluid or vapor leakage in any area where they may exist and, in case of ignition, to 

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3/14/2000 AC 25-22 

minimize the resultant hazard to the aircraft. The definition in 14 CFR 1.1 says "Flammable, 

with respect to a fluid or gas, means susceptible to igniting readily or to exploding. Examples 

are aviation jet fuel, hydraulic fluid, and oxygen. Hydraulic equipment, e.g., pumps are qualified 

per airplane and engine manufacturers' specifications, (SAE) industry standards and/or military 

standards addressing issues of bonding and grounding, flame and arc resistance, explosion 

proofing, dielectric strength and case drain temperature limits. 

 c. Background. For guidance on compliance with this requirement, refer to Advisory 

Circular (AC) 25-XX, Propulsion Systems Handbook, chapter 3, section 1. 

 d. Policy/Compliance Methods. For guidance with this requirement, refer to Advisory 

Circular (AC) 25-XX, Propulsion Systems Handbook, chapter 3, section 1. Section 25.1435(c), 

Hydraulic systems-fire protection, requires compliance with § 25.863. 

 e. Reference. The address for ordering the latest revision of the advisory circular listed 

below can be found in the Appendix to this AC. 

AC 25-XX, Propulsion Systems Handbook. 

41. SECTION 25.869 - FIRE PROTECTION SYSTEMS.

 a. Rule Text.

(a) Electrical system components:

(1) Components of the electrical system must meet the applicable fire and smoke 

protection requirements of 25.831(c) and 25.863. 

(2) Electrical cables, terminals, and equipment in designated fire zones, that are 

used during emergency procedures, must be at least fire resistant. 

(3) Main power cables (including generator cables) in the fuselage must be

designed to allow a reasonable degree of deformation and stretching without 

failure and must be- 

(i) Isolated from flammable fluid lines; or 

(ii) Shrouded by means of electrically insulated, flexible conduit, or equivalent, 

which is in addition to the normal cable insulation. 

(4) Insulation on electrical wire and electrical cable installed in any area of the 

fuselage must be self-extinguishing when tested in accordance with the applicable 

portions of Part I, appendix F of this Part. 

(b) Each vacuum air system line and fitting on the discharge side of the pump that 

might contain flammable vapors or fluids must meet the requirements 25.1183 if 

the line or fitting is in a designated fire zone. Other vacuum air systems 

components in designated fire zones must be at least fire resistant. 

(c) Oxygen equipment and lines must- 

(1) Not be located in any designated fire zone, 

(2) Be protected from heat that may be generated in, or escape from, any 

designated fire zone, and 

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3/14/2000 AC 25-22 

(3) Be installed so that escaping oxygen cannot cause ignition of grease, fluid, or 

vapor accumulations that are present in normal operation or as a result of failure 

or malfunction of any system. 

[Amdt. 25-72, 55 FR 29784, Jul. 20, 1990]

b. Intent of Rule. The intent of this rule is to minimize the probability of fire damage and 

to assure critical components of the electrical system, vacuum air system, and oxygen equipment 

and lines maintain the capability needed to complete safe flight and landing following a fire. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Section 4b.626 of 

the CAR became § 25.1359 of 14 CFR for electrical system fire and smoke protection; section 

4b.658 of the CAR became §§ 25.1433(b) and 25.1433(c) of 14 CFR for vacuum air system line 

and fittings fire protection; and § 4b.651(f) of the CAR became § 25.1451 of 14 CFR for oxygen 

equipment and lines fire protection.. 

 (1) Amendment 25-72 (July 20, 1990) consolidated all fire protection requirements for 

electrical, vacuum, and oxygen systems and transferred to § 25.869 “Fire protection: systems.” 

Section 25.1359 became § 25.869(a) except for test acceptance criteria which was transferred to 

Appendix F. Sections 25.1433(b) and (c) were combined as §§ 25.869(b) and 25.1451 became

§ 25.869(c). Minor editorial changes were also made as necessary. 

 d. Policy/Compliance Methods: For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. For policy and guidance on 

§ 25.869(a), electrical system components, refer to Advisory Circular (AC) 25-16, AC 25-17 

(section 25.1359, paragraph 781), and Appendix F to part 25. For policy and guidance on 

§ 25.869(b), vacuum air systems, refer to the Propulsion Systems Handbook, section 25.1183. 

For policy and guidance on § 25.869(c), oxygen equipment and lines, see paragraph 115. d.(3)(e) 

of this document and AC 25-17, section 25.1451, par 1011. 

 e. References. The address for ordering the latest revision of the advisory circulars listed 

below can be found in the Appendix to this AC.

AC 25-16, Electrical Fault and Fire Prevention and Protection. 

AC 25-17, Transport Airplane Cabin Interiors Crashworthiness Handbook. 

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3/14/2000 AC 25-22 

Chapter 2. - POWERPLANT 

Section 1. GENERAL 

42. SECTION 25.943 - NEGATIVE ACCELERATION.

 a. Rule Text.

No hazardous malfunction of an engine, an auxiliary power unit approved for use 

in flight, or any component or system associated with the powerplant or auxiliary 

power unit may occur when the airplane is operated at the negative accelerations 

within the flight envelopes prescribed in 25.333. This must be shown for the 

greatest duration expected for the acceleration. 

[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]

NOTE: This regulation is the subject of a Federal Aviation Regulations/Joint 

Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group will be 

tasked to recommend revisions to § 25.943, paragraphs 25.943 and 25X1315 of 

the JAR, as well as develop any advisory information if neccessary. 

NOTE: For policy and guidance on compliance with this requirement, see 

Advisory Circular (AC) 25-XX, Propulsion Systems Handbook, chapter 4. This 

regulation applies to hydraulic and pneumatic system components as well. See 

AC 25-7A, Flight Test Guide for Certification of Transport Category Airplanes, 

paragraph 101, for flight test procedures. Also see paragraph 25X1315 of the 

JAR and ACJ 25X1315 for additional information. 

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3/14/2000 AC 25-22 

Section 2. FUEL SYSTEM COMPONENTS 

43. SECTION 25.1001 - FUEL JETTISONING SYSTEM.

 a. Rule Text.

(a) A fuel jettisoning system must be installed on each airplane unless it is shown 

that the airplane meets the climb requirements of §§ 25.119 and 25.121(d) at 

maximum takeoff weight, less the actual or computed weight of fuel necessary for 

a 15-minute flight comprised of a takeoff, go-around, and landing at the airport of

departure with the airplane configuration, speed, power, and thrust the same as 

that used in meeting the applicable takeoff, approach, and landing climb 

performance requirements of this part. 

(b) If a fuel jettisoning system is required it must be capable of jettisoning enough 

fuel within 15 minutes, starting with the weight given in paragraph (a) of this 

section, to enable the airplane to meet the climb requirements of §§ 25.119 and 

25.121(d), assuming that the fuel is jettisoned under the conditions, except weight, 

found least favorable during the flight tests prescribed in paragraph (c) of this 

section. 

(c) Fuel jettisoning must be demonstrated beginning at maximum takeoff weight 

with flaps and landing gear up and in- 

(1) A power-off glide at 1.4 Vs1; 

(2) A climb at the one-engine inoperative best rate-of-climb speed, with the 

critical engine inoperative and the remaining engines at maximum continuous 

power; and 

(3) Level flight at 1.4 Vs1 ; if the results of the tests in the conditions specified in 

paragraphs (c) (1) and (2) of this section show that this condition could be 

critical. 

(d) During the flight tests prescribed in paragraph (c) of this section, it must be 

shown that- 

(1) The fuel jettisoning system and its operation are free from fire hazard; 

(2) The fuel discharges clear of any part of the airplane; 

(3) Fuel or fumes do not enter any parts of the airplane; and 

(4) The jettisoning operation does not adversely affect the controllability of the 

airplane. 

(e) For reciprocating engine powered airplanes, means must be provided to 

prevent jettisoning the fuel in the tanks used for takeoff and landing below the 

level allowing 45 minutes flight at 75 percent maximum continuous power. 

However, if there is an auxiliary control independent of the main jettisoning 

control, the system may be designed to jettison the remaining fuel by means 

of the auxiliary jettisoning control. 

(f) For turbine engine powered airplanes, means must be provided to prevent 

jettisoning the fuel in the tanks used for takeoff and landing below the level 

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allowing climb from sea level to 10,000 feet and thereafter allowing 45 minutes 

cruise at a speed for maximum range. However, if there is an auxiliary control 

independent of the main jettisoning control, the system may be designed to 

jettison the remaining fuel by means of the auxiliary jettisoning control. 

(g) The fuel jettisoning valve must be designed to allow flight personnel to close 

the valve during any part of the jettisoning operation. 

(h) Unless it is shown that using any means (including flaps, slots, and slats) for 

changing the airflow across or around the wings does not adversely affect fuel 

jettisoning, there must be a placard, adjacent to the jettisoning control, to warn 

flight crewmembers against jettisoning fuel while the means that change the 

airflow are being used. 

(i) The fuel jettisoning system must be designed so that any reasonably probable 

single malfunction in the system will not result in a hazardous condition due to 

unsymmetrical jettisoning of, or inability to jettison, fuel. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 

33 FR 12226, Aug. 30, 1968; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]

NOTE: This regulation will be the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under 

the Aviation Rulemaking Advisory Committee (ARAC). The ARAC working 

group may recommend revisions to the regulation and any associated advisory 

material.

NOTE: For policy and guidance on compliance with this requirement, see 

Advisory Circular (AC) 25-XX, Propulsion Systems Handbook. This regulation 

may require special consideration for certain equipment where the airplane is not 

capable of a return landing without exceeding equipment ratings/capabilities such 

as brakes and tires. Brake maximum kinetic energy rating(s) and tire maximum 

speed ratings may be exceeded for an immediate return/turnback, or a flapless 

landing, especially for large two engine airplanes. 

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Section 3. POWERPLANT FIRE PROTECTION 

44. SECTION 25.1183 - FLAMMABLE FLUID-CARRYING COMPONENTS.

 a. Rule Text.

(a) Except as provided in paragraph (b) of this section, each line, fitting, and 

other component carrying flammable fluid in any area subject to engine fire 

conditions, and each component which conveys or contains flammable fluid in a 

designated fire zone must be fire resistant, except that flammable fluid tanks and 

supports in a designated fire zone must be fireproof or be enclosed by a fireproof 

shield unless damage by fire to any non-fireproof part will not cause leakage or 

spillage of flammable fluid. Components must be shielded or located to safeguard 

against the ignition of leaking flammable fluid. An integral oil sump of less than 

25-quart capacity on a reciprocating engine need not be fireproof nor be enclosed 

by a fireproof shield. 

(b) Paragraph (a) of this section does not apply to- 

(1) Lines, fittings, and components which are already approved as part of a type 

certificated engine; and 

(2) Vent and drain lines, and their fittings, whose failure will not result in, or add 

to, a fire hazard. 

(Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR 

6913, May 5, 1967; Amdt. 25-36, 39 FR 35461, Oct. 1, 1974; Amdt. 25-57, 49 FR 

6849, Feb. 23, 1984) 

NOTE: For policy and guidance on compliance with this requirement, see 

Advisory Circular (AC) 25-XX, Propulsion Systems Handbook, chapter 4, 

section 9. Section 25.1435(c) - Hydraulic systems - Fire protection, requires 

compliance with § 25.1183. 

NOTE: Under planned harmonization effort, paragraph 25.1183(c) of the JAR 

will be adopted by FAA as § 25.1183(c); the text is as follows: 

"(c) All components, including ducts, within a designated fire zone must be 

fireproof if, when exposed to or damaged by fire, they could - 

(1) Result in fire spreading to other regions of the airplane, or 

(2) Cause unintentional operation of, or inability to operate, essential 

services or equipment." 

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45. SECTION 25.1185 - FLAMMABLE FLUIDS.

 a. Rule Text.

(a) Except for the integral oil sumps specified in 25.1013 (a), no tank or reservoir 

that is a part of a system containing flammable fluids or gases may be in a 

designated fire zone unless the fluid contained, the design of the system, the 

materials used in the tank, the shut-off means, and all connections, lines, and 

control provide a degree of safety equal to that which would exist if the tank or 

reservoir were outside such a zone. 

(b) There must be at least one-half inch of clear airspace between each tank or 

reservoir and each firewall or shroud isolating a designated fire zone. 

(c) Absorbent materials close to flammable fluid system components that might 

leak must be covered or treated to prevent the absorption of hazardous quantities 

of fluids. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964 as amended by Amdt. 25-19, 

33 FR15410, Oct. 17, 1968]

NOTE: For policy and guidance on compliance with this requirement, see 

Advisory Circular (AC) 25-XX, Propulsion Systems Handbook, chapter 4, 

section 9. Section 25.1435(c) - Hydraulic systems - Fire protection, requires 

compliance with § 25.1185. 

46. SECTION 25.1189 - SHUTOFF MEANS.

 a. Rule Text.

(a) Each engine installation and each fire zone specified in § 25.1181(a) (4) and 

(5) must have a means to shut off or otherwise prevent hazardous quantities of 

fuel, oil, deicer, and other flammable fluids, from flowing into, within, or through 

any designated fire zone, except that shutoff means are not required for-

(1) Lines, fittings, and components forming an integral part of an engine; and 

(2) Oil systems for turbine engine installations in which all components of the 

system in a designated fire zone, including oil tanks, are fireproof or located in 

areas not subject to engine fire conditions. 

(b) The closing of any fuel shutoff valve for any engine may not make fuel 

unavailable to the remaining engines. 

(c) Operation of any shutoff may not interfere with the later emergency operation 

of other equipment, such as the means for feathering the propeller. 

(d) Each flammable fluid shutoff means and control must be fireproof or must be 

located and protected so that any fire in a fire zone will not affect its operation. 

(e) No hazardous quantity of flammable fluid may drain into any designated fire 

zone after shutoff. 

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3/14/2000 AC 25-22 

(f) There must be means to guard against inadvertent operation of the shutoff 

means and to make it possible for the crew to reopen the shutoff means in flight

after it has been closed. 

(g) Each tank-to-engine shutoff valve must be located so that the operation of the 

valve will not be affected by powerplant or engine mount structural failure. 

(h) Each shutoff valve must have a means to relieve excessive pressure 

accumulation unless a means for pressure relief is otherwise provided in the 

system. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 

35FR 5677, Apr. 8, 1970; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984]

NOTE: This regulation is the subject of a Federal Aviation Regulations/Joint Aivation 

Requirements (FAR/JAR) harmonization effort under the Aviation Rulemaking Advisory 

Committee (ARAC). The ARAC working group is planning to recommend criteria for 

compliance as advisory material. 

NOTE: For policy and guidance on compliance with this requirement, see 

Advisory Circular (AC) 25-XX, Propulsion Systems Handbook, chapter 4, 

section 9. Section 25.1435(c) - Hydraulic systems - Fire protection, requires 

compliance with § 25.1189. 

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3/14/2000 AC 25-22 

Chapter 3. EQUIPMENT

Section 1. GENERAL 

47. SECTION 25.1301 - FUNCTION AND INSTALLATION.

 a. Rule Text.

Each item of installed equipment must- 

(a) Be of a kind and design appropriate to its intended function; 

(b) Be labeled as to its identification, function, or operating limitations, or any applicable 

combination of these factors; 

(c) Be installed according to limitations specified for that equipment; and 

(d) Function properly when installed. 

b. Intent of Rule. The intent of this requirement is to define general conditions for 

certification of all installed equipment regardless if it's covered under another requirement. This 

requirement applies to all equipment/systems whether required as minimum equipment for 

installation on airplanes or non-required/non-essential equipment (e.g., advisory systems, pilot 

aids, or passenger comfort related amenities). 

 c. Background. This requirement originated under § 4b.682 of the Civil Air Regulations 

(CAR), October 1, 1949, with reference to items of equipment for which type certification was 

required outlined in part 15 of the CAR. Effective July 20, 1950, it was recodified as § 4b.601 

by deleting reference to part 15 per Amendment 04b-2 (effective February 6, 1950), and 

allowing approval of materials, parts, and appliances per § 04b.05 and the new TSO system. 

Effective February 1, 1965, part 25 was added to Title 14, Code of Federal Regulations (14 CFR) 

to replace part 4b of the CAR. Section 4b.601 of the CAR became § 25.1301 of 14 CFR. 

 (1) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group is recommending 

revisions to §§ 25.1301, 25.1309, a new 25.1310, and a revised Advisory Circular AC 25.1309-

1B. Section 25.1301(d) will be deleted as being redundant with the proposed § 25.1309(a). 

 d. Policy/Compliance Methods. Information regarding installation limitations and proper 

functioning is normally available from the equipment manufacturers in their installation and 

operations manuals. For example, TSO Rating Limits apply under § 25.1309(c). Under 

§ 25.1301(d), TSO is not an installation approval. The user/installer is responsible for form, fit, 

and function. 

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3/14/2000 AC 25-22 

(1) Joint Aviation Requirements (JAR) ACJ 25.1301(b). The following guidance is 

extracted from JAR ACJ 25.1301(b): 

(a) An adequate means of identification should be provided for all cables, 

connectors, and terminals. The means employed should be such as to ensure that the 

identification does not deteriorate under service conditions. 

(b) When pipelines (hydraulic/pneumatic, etc.) are marked for the purpose of 

distinguishing their functions, the markings should be such that the risk of confusion by 

maintenance or servicing personnel will be minimized. Distinction by means of color markings 

alone is not acceptable. The use of alphabetic or numerical symbols will be acceptable if 

recognition depends upon reference to a master key and any relation between symbol and 

function is carefully avoided. Specification ISO.12 gives acceptable graphical markings. 

 e. Reference. None. 

48. SECTION 25.1309 - EQUIPMENT, SYSTEMS, AND INSTALLATION.

 a. Rule Text.

(a) The equipment, systems, and installations whose functioning is required by 

this subchapter, must be designed to ensure that they perform their intended 

functions under any foreseeable operating condition. 

(b) The airplane systems and associated components, considered separately and 

in relation to other systems, must be designed so that- 

(1) The occurrence of any failure condition which would prevent the continued 

safe flight and landing of the airplane is extremely improbable, and 

(2) The occurrence of any other failure conditions which would reduce the 

capability of the airplane or the ability of the crew to cope with adverse operating 

conditions is improbable. 

(c) Warning information must be provided to alert the crew to unsafe system 

operating conditions, and to enable them to take appropriate corrective action. 

Systems, controls, and associated monitoring and warning means must be 

designed to minimize crew errors which could create additional hazards. 

(d) Compliance with the requirements of paragraph (b) of this section must be 

shown by analysis, and where necessary, by appropriate ground, flight, or

simulator tests. The analysis must consider- 

(1) Possible modes of failure, including malfunctions and damage from external 

sources. 

(2) The probability of multiple failures and undetected failures. 

(3) The resulting effects on the airplane and occupants, considering the stage of 

flight and operating conditions, and 

(4) The crew warning cues, corrective action required, and the capability of 

detecting faults. 

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3/14/2000 AC 25-22 

(e) Each installation whose functioning is required by this subchapter, and that 

requires a power supply, is an ''essential load'' on the power supply. The power 

sources and the system must be able to supply the following power loads in 

probable operating combinations and for probable durations: 

(1) Loads connected to the system with the system functioning normally. 

(2) Essential loads, after failure of any one prime mover, power converter, or 

energy storage device. 

(3) Essential loads after failure of- 

(i) Any one engine on two-engine airplanes; and 

(ii) Any two engines on three-or-more-engine airplanes. 

(4) Essential loads for which an alternate source of power is required by this 

chapter, after any failure or malfunction in any one power supply system, 

distribution system, or other utilization system. 

(f) In determining compliance with paragraphs (e) (2) and (3) of this section, the 

power loads may be assumed to be reduced under a monitoring procedure 

consistent with safety in the kinds of operation authorized. Loads not required in 

controlled flight need not be considered for the two-engine-inoperative condition 

on airplanes with three or more engines. 

(g) In showing compliance with paragraphs (a) and (b) of this section with regard 

to the electrical system and equipment design and installation, critical 

environmental conditions must be considered. For electrical generation, 

distribution, and utilization equipment required by or used in complying with this 

chapter, except equipment covered by Technical Standard Orders containing 

environmental test procedures, the ability to provide continuous, safe service 

under foreseeable environmental conditions may be shown by environmental 

tests, design analysis, or reference to previous comparable service experience on 

other aircraft. 

[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-38, 

41 FR 55467, Dec. 20, 1976; Amdt. 25-41, 42 FR 36970, Jul. 18, 1977]

b. Intent of Rule. This regulation covers, but is not limited to, mechanical, electrical, 

pneumatic, and hydraulic power sources, associated distribution, and corresponding utilization 

systems. 

 c. Background. This regulation originated under Amendment 4b-6 of the Civil Air 

Regulations (CAR), effective March 5, 1952, as § 4b.606 and was codified on December 31, 

1953. It was amended under Amendment 4b-1, effective May 18, 1954, by clarifying the 

allowed assumptions in the determination of probable operating combinations of essential loads 

for the power failure conditions. Effective February 1, 1965, part 25 was added to Title 14, Code 

of Federal Regulations (14 CFR) to replace part 4b of the CAR. Section 4b.606 of the CAR 

became § 25.1309 of 14 CFR. 

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3/14/2000 AC 25-22 

 (1) Amendment 25-23 (April 8, 1970) added these requirements: 

(i) Where the consequences of complete loss of system function would be 

catastrophic, combinations of failures and multiple failures due to common causes must be 

considered, and sufficient reliability, redundancy and isolation provided to make catastrophic 

systems failure extremely improbable. 

(ii) Occurrence of failures that result in serious degradation of flight characteristics, 

a large increase in crew work-load, or difficult emergency procedures be improbable. 

(iii) Warning information must be provided to alert the crew to unsafe system

operating conditions and to enable them to take corrective action. 

(iv) Compliance to (i), (ii), and (iii) be shown by analysis and 

ground/flight/simulator tests. 

 (2) Amendment 25-38 (February 1, 1977) amended the title by inserting a comma 

between the words "equipment" and "Systems" and between the words "Systems" and "and." 

The change was strictly editorial. 

 (3) Amendment 25-41 (September 1, 1977(revised §§ 25.1309(b)(2), (c), lead-in of (d), 

(e)(3), and (f). The reference to occupant injury in (b)(2) was deleted because the matter of 

preventing injuries to occupants is covered elsewhere in the regulations; deleted reference to (c) 

in (d) because it could have been unreasonably burdensome for compliance; added provisions in 

(c) to minimize crew errors; revised (e)(3) and (f) related to electrical systems. 

(4) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group is recommending 

revisions to §§ 25.1301, 25.1309, a new 25.1310, and a revised AC 25.1309-1B. Section 

25.1301(d) will be deleted as being redundant with the proposed § 25.1309(a). 

 d. Policy/Compliance Methods. The term "subchapter" in § 25.1309(a) is intended to 

cover not only subchapter "C," but also the equipment, systems, and installations not specifically 

required by subchapter "C" but installed in airplanes in order to engage in operations covered by 

other subchapters. Dependence for safety of flight might be placed on installations not otherwise 

mentioned in the rules. 

(1) Probability of a Fire. The following is extracted from an FAA memorandum, dated 

March 9, 1995. This is in response to a request for guidance regarding the use of the probability 

of a fire in a failure analysis when showing compliance with § 25.1309. The focus of the request 

was directed at the three issues addressed below: 

(a) The first issue concerns whether the occurrence of a cargo compartment fire 

should be considered to have a probability of one, or something less than one. Advisory Circular 

(AC) 25.1309-1A, paragraph 8e, "Operational and Environmental Conditions," allows that 

random conditions may be considered to have a probability of occurrence less than one, and may 

usually be included in a safety analysis. The AC indicates that a random condition is a condition 

128 
3/14/2000 AC 25-22 

for which the airplane is not designed, and in which the airplane is not normally approved to 

operate. It also indicates that a random condition may be used in the analysis even when the 

system under analysis is designed to protect against the occurrence of the random condition. The 

AC provides, as an example, that it would be reasonable and rational to use a probability of less 

than one of encountering hazardous turbulence or gust levels after the failure of a structural load 

alleviation system. A probability less than one of fire occurring after a failure of the fire 

protection system is similar to this example, and may be used in the analysis. Advisory Circular 

25.1309-1A also provides counter-examples for conditions which should not be considered 

random, and for which a probability of one should be used. These are conditions for which the 

airplane is designed. The examples provided, instrument meteorological conditions and 

Category III weather operations, are conditions in which the airplane would be expected to 

operate as a matter of course, and for which operational approval would be granted. 

(b) The second issue concerns what value less than one should be assigned to the 

probability of the occurrence of a fire. Advisory Circular 25.1309-1A, paragraph 8e, provides 

guidance that the statistically-derived probability used in the analysis should be based on an 

applicable supporting database and a valid statistical distribution. When requesting approval for 

the use of the probability of a random condition in a safety analysis, it is incumbent on the 

applicant to supply the data, show the applicability of the database from which the data is 

supplied, and derive a valid statistical conclusion. The ACO should then evaluate the applicant's 

statistical information and determine if the value used in the analysis is supported by the data 

presented. This value must be reassessed for subsequent programs, as service experience gained 

in the future may require the probability of the occurrence to be re-evaluated. 

(c) The third issue is whether the required safety level reached through the use of 

the Maintenance Steering Group, Revision 3 (MSG-3) procedures is higher than that reached 

using a § 25.1309 related analysis. We contend that this is not so, and that a like-comparison of 

the two sets of procedures is not valid. The § 25.1309 analysis process is directed at assessing 

the contributions of failures to given failure conditions, and to directing the airplane or system

design appropriately. The MSG-3 process is directed at determining appropriate maintenance 

actions given the contribution of a failure to a reduction in safety. Findings from the § 25.1309 

analysis process are used as starting points for certain MSG-3 processes to determine 

maintenance activities and intervals, or whether the maintenance process can adequately 

minimize the risk elements assigned to it. The MSG-3 procedures must assume the combination 

of a system failure with the occurrence of the condition, which the system was designed to 

protect against, in order to account for hidden failures when determining maintenance actions. 

Expected or allowable probabilities related to system failures are not derived through the MSG-3 

process. In applying the above guidance, note that § 25.1309 is a rule of general applicability, 

and should not be used to replace a more specific and stringent requirement. Cargo 

compartments, and compartment fire detection and protection systems must meet requirements 

specified in §§ 25.855, 25.857, and 25.858, regardless of probabilities as determined in a 

§ 25.1309 related analysis. 

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 (2) Examples. For an example of compliance with § 25.1309(c), see paragraph d(1) 

under § 25.1441, Oxygen equipment and supply, of this AC. Another example of compliance is 

included under paragraph d(2), § 25.855, Cargo or baggage compartments, of this AC. 

(3) Safe and Reliable. The following is extracted from an FAA memorandum dated 

October 26, 1990, and addresses an inquiry regarding (i) The meaning of "safe and reliable" used 

in §§ 25.109(b) and 25.125(b)(3) in terms of AC 25.1309-1A probability, and (ii) The acceptable 

failure probability for deceleration devices like anti-skid and ground spoilers that have relatively 

large and small effects on landing distance. The FAA response follows: 

(a) Safe and reliable is generally used to mean that a failure condition is 

improbable. In terms of AC 25.1309-1A quantitative probability terms, improbable failure 

conditions are those having a probability on the order of 1 X 10-5 or less. 

(b) Each deceleration device, such as anti-skid and ground spoilers, would be 

expected to have a failure condition that is at least improbable (i.e. 1 X 10-5 or less) regardless of 

its effect on stopping distance. 

 e. References. For policy and guidance on compliance with this requirement, see 

AC 25.1309-1A, System Design Analysis, issued June 21, 1988, or latest revision, and 

AC 25-19, Certification Maintenance Requirements, issued November 28, 1994, or latest 

revision. 

49 - 51. [RESERVED] 

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Section 2. INSTRUMENTS: INSTALLATION 

52. SECTION 25.1325 - STATIC PRESSURE SYSTEMS.

a. Rule Text. 

(a) Each instrument with static air case connections must be vented to the outside 

atmosphere through an appropriate piping system. 

(b) Each static port must be designed and located in such manner that the static 

pressure system performance is least affected by airflow variation, or by moisture 

or other foreign matter, and that the correlation between air pressure in the static 

pressure system and true ambient atmospheric static pressure is not changed 

when the airplane is exposed to the continuous and intermittent maximum icing 

conditions defined in Appendix C of this Part. 

(c) The design and installation of the static pressure system must be such 

that- 

(1) Positive drainage of moisture is provided; chafing of the tubing and excessive 

distortion or restriction at bends in the tubing is avoided; and the materials used 

are durable, suitable for the purpose intended, and protected against corrosion; 

and 

(2) It is airtight except for the port into the atmosphere. A proof test must be

conducted to demonstrate the integrity of the static pressure system in the 

following manner: 

(i) Unpressurized airplanes. Evacuate the static pressure system to a pressure 

differential of approximately 1 inch of mercury or to a reading on the altimeter, 

1,000 feet above the airplane elevation at the time of the test. Without additional 

pumping for a period of 1 minute, the loss of indicated altitude must not exceed 

100 feet on the altimeter. 

(ii) Pressurized airplanes. Evacuate the static pressure system until a pressure 

differential equivalent to the maximum cabin pressure differential for which the 

airplane is type certificated is achieved. Without additional pumping for a period 

of 1 minute, the loss of indicated altitude must not exceed 2 percent of the 

equivalent altitude of the maximum cabin differential pressure or 100 feet, 

whichever is greater. 

(d) Each pressure altimeter must be approved and must be calibrated to indicate 

pressure altitude in a standard atmosphere, with a minimum practicable 

calibration error when the corresponding static pressures are applied. 

(e) Each system must be designed and installed so that the error in indicated 

pressure altitude, at sea level, with a standard atmosphere, excluding instrument 

calibration error, does not result in an error of more than +/- 30 feet per 100 

knots speed for the appropriate configuration in the speed range between 1.3 Vso 

with flaps extended and 1.8 Vs1 with flaps retracted. However, the error need not 

be less than +/- 30 feet. 

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(f) If an altimeter system is fitted with a device that provides corrections to the 

altimeter indication, the device must be designed and installed in such manner 

that it can be bypassed when it malfunctions, unless an alternate altimeter system 

is provided. Each correction device must be fitted with a means for indicating the 

occurrence of reasonably probable malfunctions, including power failure, to the 

flight crew. The indicating means must be effective for any cockpit lighting 

condition likely to occur. 

(g) Except as provided in paragraph (h) of this section, if the static pressure 

system incorporates both a primary and an alternate static pressure source, the 

means for selecting one or the other source must be designed so that- 

(1) When either source is selected, the other is blocked off; and 

(2) Both sources cannot be blocked off simultaneously. 

(h) For unpressurized airplanes, paragraph (g)(1) of this section does not apply if 

it can be demonstrated that the static pressure system calibration, when either 

static pressure source is selected, is not changed by the other static pressure 

source being open or blocked. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-5, 

30 FR 8261, Jun. 29, 1965; Amdt. 25-12, 32 FR 7587, May 24, 1967; Amdt. 25-

41, 42 FR 36970, Jul. 18, 1977]

b. Intent of Rule. This rule provides minimum design and certification requirements for 

static pressure systems to ensure proper static system operation in varying operating conditions.

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Many of the 

requirements in § 4b.612(b) of the CAR were carried over essentially unchanged to § 25.1325 of 

14 CFR. 

 (1) Amendment 25-5 (June 29, 1965) added requirements for (1) operation in icing 

conditions and (2) devices that provide corrections to the altimeter indication. 

 (2) Amendment 25-12 (May 24, 1967) added requirements for static system proof 

testing for pressurized and unpressurized airplanes. 

 (3) Amendment 25-41 (July 18, 1977) added standards for static pressure systems, 

which incorporate the ability to select one or the other of primary and alternate static systems.

 d. Policy/Compliance Methods. For policy and guidance on compliance with this 

requirement, refer to Advisory Circular (AC) 25-XX, Electrical Systems Handbook. 

e. References. None. 

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Section 3. SAFETY EQUIPMENT 

53. SECTION 25.1419 ICE PROTECTION.

a. Rule Text.

If certification with ice protection provisions is desired, the airplane must be able 

to safely operate in the continuous maximum and intermittent maximum icing 

conditions of appendix C. To establish that the airplane can operate within the 

continuous maximum and intermittent maximum conditions of appendix C: 

(a) An analysis must be performed to establish that the ice protection for the 

various components of the airplane is adequate, taking into account the various 

airplane operational configurations; and 

(b) To verify the ice protection analysis, to check for icing anomalies, and to 

demonstrate that the ice protection system and components are effective, the 

airplane or its components must be flight tested in the various operational 

configurations, in measured natural atmospheric icing conditions and, as found 

necessary, by one or more of the following means: 

(1) Laboratory dry air or simulated icing tests, or a combination of both, of the 

components or models of the components. 

(2) Flight dry air tests of the ice protection system as a whole, or of its individual 

components. 

(3) Flight tests of the airplane or its components in measured simulated icing 

conditions. 

(c) Caution information, such as an amber caution light or equivalent, must be 

provided to alert the flightcrew when the anti-ice or de-ice system is not 

functioning normally. 

(d) For turbine engine powered airplanes, the ice protection provisions of this 

section are considered to be applicable primarily to the airframe. For the 

powerplant installation, certain additional provisions of Subpart E of this Part 

may be found applicable. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 

35 FR 5680, Apr. 8, 1970; Amdt. 25-72, 55 FR 29785, Jul. 20, 1990]

b. Intent of Rule. This rule provides standards for certification of an airplane with ice 

protection provisions. Certification for flight in icing conditions is optional at the discretion of 

the applicant. If the airplane is not certificated for flight in icing condition, appropriate 

limitations shall be placed in the type certificate data sheet and the airplane flight manual. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Many of the 

provisions in section 4b.640 of the CAR were carried over with minor word changes to 

§ 25.1419 of 14 CFR. However, the intent of the CAR 4b.640 was unchanged. 

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3/14/2000 AC 25-22 

Amendment 25-23 revised paragraph (c) to require flight tests in measured natural atmospheric 

icing conditions. Prior to Amendment 25-23 flight tests in measured natural atmospheric icing 

conditions were optional. 

 (1) Amendment 25-23 (April 8, 1970) added the requirements for flight testing in 

natural icing conditions contained in § 25.1419(c). 

 (2) Amendment 25-72 (July 20, 1990) rearranged the sub-paragraphs and inserted a 

new paragraph (c), which requires caution information when the anti-ice or de-ice system is not 

functioning normally. Prior to this amendment, a means to determine that a pneumatic de-icing 

boot system was working normally was required by § 25.1416. Section 25.1416 was removed to 

allow for the concept of the "dark cockpit.” (The term “dark cockpit” is descriptive language that

refers to the absence of annunciation lights in the cockpit unless a system is operating 

abnormally.) 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and AC 20-73 and AC 25.1419-1. 

 e. References. The addresses for ordering the latest revision of advisory circulars, and 

other referenced documents listed below can be found in the Appendix to this AC. 

AC 20-73, Aircraft Ice Protection. 

AC 23.1419-2, Certification of Part 23 Airplanes for Flight in Icing Conditions. 

AC 25.1419-1, Certification of Transport Category Airplanes for Flight in Icing 

Conditions. 

Report No. DOT/FAA/CT-88/8-1 - Aircraft Icing Handbook (three volumes) 

FAA Technical Report ADS-4 - Engineering summary of Airframe Icing Technical 

Data, December 1963. (Although most of the information contained in this report is still valid, 

some is outdated, and more usable information is now available through recent research and 

experience and is included in the Aircraft Icing Handbook.) 

54 - 58. [RESERVED] 

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Section 4. MISCELLANEOUS EQUIPMENT 

59. SECTION 25.1433 - VACUUM SYSTEMS.

 a. Rule Text.

There must be means, in addition to the normal pressure relief, to automatically 

relieve the pressure in the discharge lines from the vacuum air pump when the 

delivery temperature of the air becomes unsafe. 

[Amdt. 25-72, 55 FR 29785, Jul. 20, 1990]

b. Intent of Rule. This rule provides minimum design and certification requirements for 

vacuum systems, specifically to require automatic pressure relief in the presence of high delivery 

air temperatures. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in section 4b.658 of the CAR were carried over essentially unchanged to § 25.1433 of 14 CFR. 

 (1) Amendment 25-72 (July 20, 1990) moved §§ 25.1433(b) and 25.1443(c), which 

dealt with fire protection for vacuum air systems, to § 25.869, Fire protection: Systems. 

 d. Policy/Compliance Methods. There is no existing written policy or guidance related to 

this subject in our files.

 e. References. None.

60. SECTION 25.1435 - HYDRAULIC SYSTEMS.

 a. Rule Text.

(a) Design. 

(1) Each element of the hydraulic system must be designed to withstand, without 

deformation that would prevent it from performing its intended function, the 

design operating pressure loads in combination with limit structural loads which 

may be imposed. 

(2) Each element of the hydraulic system must be able to withstand, without 

rupture, the design operating pressure loads multiplied by a factor of 1.5 in 

combination with ultimate structural loads that can reasonably occur 

simultaneously. Design operating pressure is maximum normal operating 

pressure, excluding transient pressure. 

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3/14/2000 AC 25-22 

(b) Tests and analysis. 

(1) A complete hydraulic system must be static tested to show that it can withstand 

1.5 times the design operating pressure without a deformation of any part of the 

system that would prevent it from performing its intended function. Clearance 

between structural members and hydraulic system elements must be adequate and 

there must be no permanent detrimental deformation. For the purpose of this test, 

the pressure relief valve may be made inoperable to permit application of the 

required pressure. 

(2) Compliance with 25.1309 for hydraulic systems must be shown by functional 

tests, endurance tests, and analyses. The entire system, or appropriate 

subsystems, must be tested in an airplane or in a mockup installation to determine 

proper performance and proper relation to other aircraft systems. The functional 

tests must include simulation of hydraulic system failure conditions. Endurance 

tests must simulate the repeated complete flights that could be expected to occur 

in service. Elements which fail during the tests must be modified in order to have 

the design deficiency corrected and, where necessary, must be sufficiently 

retested. Simulation of operating and environmental conditions must be 

completed on elements and appropriate portions of the hydraulic system to the 

extent necessary to evaluate the environmental effects. Compliance with 25.1309 

must take into account the following: 

(i) Static and dynamic loads including flight, ground, pilot, hydrostatic, inertial 

and thermally induced loads, and combinations thereof. 

(ii) Motion, vibration, pressure transients, and fatigue. 

(iii) Abrasion, corrosion, and erosion. 

(iv) Fluid and material compatibility. 

(v) Leakage and wear. 

(c) Fire protection. Each hydraulic system using flammable hydraulic fluid must 

meet the applicable requirements of 25.863, 25.1183, 25.1185, and 25.1189. 

[Amdt. 25-13, 32 FR 9154, Jun. 28, 1967, as amended by Amdt. 25-41, 

42 FR 36971, Jul. 18, 1977; Amdt. 25-72, 55 FR 29786, Jul. 20, 1990]

b. Intent of Rule. The intent of this rule is to provide minimum design, performance, and 

safety requirements for transport category airplane hydraulic systems addressing the following 

aspects of component design and qualification, subsystem/system design, and system integration: 

(1) Design operating loads in combination with limit and ultimate structural loads. 

(2) Pressure transients and cyclic pressures. 

 (3) Environmental conditions. 

(4) Fatigue and endurance life. 

(5) Single and multiple failure modes. 

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3/14/2000 AC 25-22 

(6) Indication and warning. 

 (7) Installation and support. 

 (8) Pump/engine interface. 

 (9) Fire protection. 

(10) Continued safe flight and landing . 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Sections 4b.653, 

4b.654, and 4b.655 respectively became §§ 25.1435(a), 25.1435(b), and 25.1435(c) of 14 CFR 

for hydraulic systems. Since then, § 25.1435 has been revised under Amendment 25-13, 

Amendment 25-41, and Amendment 25-72 to make the regulations more comprehensive and to 

delete redundancies. 

 (1) Amendment 25-13 (June 28, 1967) prescribed more comprehensive design and test 

requirements by requiring (1) cockpit indication of system fluid quantity, (2) a means to prevent 

harmful or hazardous concentrations of the fluid or vapors in the crew or passenger 

compartments during flight due to hydraulic fluid leakage/release, and (3) compliance with 

§ 25.1309 by functional tests, endurance tests, and analysis of the entire system or appropriate 

subsystems tested in an airplane or in a mockup installation to determine proper performance and 

proper relation to other aircraft systems. 

(a) Regarding compliance with § 25.1435(b)(2), the preamble to the final rule 

states in part that, "the FAA agrees that the proposed environmental testing of the assembled 

system is not necessary in order to achieve a reasonable and effective testing program for 

improving hydraulic system reliability." The final rule was changed to allow applicants to meet 

endurance test requirements by performing endurance testing on components as long as the test 

program was representative of the airplane installation and environment. 

(b) Applicants have shown compliance to the endurance testing requirements of 

§ 25.1435 by conducting endurance tests on components or subassemblies using fixtures that 

represent the airplane installation and environmental conditions. The FAA normally accepts 

endurance test programs that vary hydraulic fluid temperature during endurance testing. The 

other environmental conditions are usually tested independent of the endurance test program. 

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3/14/2000 AC 25-22 

The regulation does not specify the number of repeated cycles to meet endurance test 

requirements. As a minimum, the applicant must show that the reliability of components will 

meet the airplane system reliability requirements of § 25.1309. 

 (2) Amendment 25-23 (April 8, 1970) expanded § 25.1435(a)(4) to require means to 

ensure that no pressure will exceed a safe limit above design operating pressure by specifying 

pressure variation tolerances of +/- 10% on the pump discharge pressures and an upper limit of 

125% of the design operating pressure for transients. These tolerances were based on in-service 

experience with such systems. 

 (3) Amendment 25-41 (July 18, 1977) clarified under § 25.1435(a)(2) that the pressure 

and quantity indication requirement was applicable to systems performing a function that is 

essential for continued safe flight and landing, or requiring corrective crew action when a system

malfunction has occurred. It also recognized that means other than gages (e.g., warning lights) 

were an acceptable means of compliance. Sections 25.1435(a)(7) and 25.1435(a)(8) were added 

to allow certain transient pressures to exceed prescribed limits under (a)(4)(ii), provided the 

resulting fatigue strength was accounted for, and to require pump design such that loss of fluid 

condition could not create a hazard preventing continued safe flight and landing. 

 (4) Amendment 25-72 (July 20, 1990). Prior to this amendment, § 25.1435(b) was 

labeled "Tests," and contained references to a hydraulic system proof test, which was needed in 

order to comply with § 25.1309. Under Amendment 25-72, the general requirements of the old 

§§ 25.1435(a)(2) through 25.1435(a)(8), relative to indication, system pressures, transients, 

volumetric changes, pump discharge pressure limits, ripple damping devices, vibration, abrasion, 

corrosion, mechanical damage, inertia loads, hazardous vapors, relative motion, differential 

vibration, transient pressures, fatigue strength and loss of fluid to the pumps, were consolidated 

under a general listing as "Tests and Analyses" in the new § 25.1435(b)(2). 

 (5) Exemptions. In October 1993, pursuant to the authority contained in §§ 313(a) and 

601(c) of the Federal Aviation Act of 1958, an applicant was granted an exemption (refer to 

exemption 5758A) from § 25.1435(b)(1) of Title 14, Code of Federal Regulations (14 CFR) to 

the extent necessary to permit type certification of a new airplane by testing of the complete 

hydraulic system at the system relief pressure, but not less than 3400 psig, in lieu of 1.5 times the 

design operating pressure (4500 psig). Additional similar exemptions have been granted (Refer 

to Exemptions 6086, 6504, and 6577). 

138 
3/14/2000 AC 25-22 

 (6) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group has recommended 

revisions to § 25.1435 and a new Advisory Circular (AC) 25.1435-1. The harmonized rule will 

reflect the exemptions granted, see (5) above. The notice of proposed rulemaking (NPRM 96-6) 

and accompanying proposed AC 25.1435-1, were published in the Federal Register on July 3, 

1996 (61 FR 35056) and public comments were invited. The final rule is at the FAA 

headquarters for a regulatory evaluation. The final rule and AC are expected to be adopted 

in 2000. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Hydraulic Fluid Contamination. The following was extracted from an FAA letter 

to the National Transportation Safety Board (NTSB), dated June 29, 1998, and related 

correspondence. The NTSB accepted the FAA-Industry (Society of Automotive Engineers 

Committee A-6) task force resolution concerning the NTSB safety recommendation, A-96-116. 

(a) Background. Following the crash of the USAir flight 427, a Boeing 737-300, 

on September 8, 1994, an NTSB investigation team found hydraulic fluid with a high particulate 

count in the main rudder power control unit (PCU). While the effect of contaminated hydraulic 

fluid on the different aircraft systems was not known at that time, it was examined as a potential 

factor. In October 1996, the NTSB made safety recommendation, A-96-116, which 

recommended that the FAA define and implement standards for in-service hydraulic fluid 

cleanliness and sampling intervals for all transport-category aircraft. 

(b) Discussion. An April 1995, FAA study concluded that the existing standard, 

NAS 1638, was adequate for classifying the particulate contamination levels for aircraft 

hydraulic fluid. At the FAA's request, an industry task force formed by the Society of 

Automotive Engineers (SAE) Committee A-6 studied the fluid contamination issues. The 

industry task force concluded that (i) flight control servoactuators had demonstrated operation at 

contamination levels up to NAS 1638 class 17 and higher and, (ii) an in-service limit of NAS 

1638 class 9 prescribed by the majority of current airframe manufacturers was conservative, 

adequate, and the maximum recommended limit. The task force also studied chemical 

contamination effects and provided in-service limits for fluid properties such as specific gravity, 

moisture content, viscosity, and chlorine content. 

(c) Chemical Sensitivity. Actuation systems are insensitive to chemical 

contamination within normal in-service operational limits (see tabulation below). Gross 

contamination with other fluids is prohibited. 

139
3/14/2000 AC 25-22 

Chemical Quality Recommendations for In-Service Hydraulic Fluid 

Analysis Limit Reason

Appearance No cloudiness or phase 

separation or

precipitation 

Particulate and/or chemical 

contamination 

Moisture 1.0 percent maximum Corrosion, fluid stability (acidity), 

low temperature pumpability 

Neutralization number 1.5 mg KOH/gm

maximum 

Corrosion, deposit formation 

Kinematic viscosity @ 

100oF/38oC

6.0 - 12.5cs Lubricity

Chlorine contamination* 200 ppm maximum Fluid purity, erosion 

* This limitation is in addition to the amount already contained in the base stock.

(d) Sensitivity Methods. The task force recognized that the inconsistencies in the 

existing hydraulic fluid sampling procedures gave wide variations in the measured levels of 

particulate contamination. As a result, an Aerospace Recommended Practice ARP-5376, 

"Methods, Locations and Criteria for System Sampling and Measuring the Solid Particle 

Contamination of Hydraulic Fluids," was developed and published in September 1998. 

(e) Sampling. The fluid sampling interval requirement for particulate 

contamination check is the same as the current one for checking chemical contamination. 

Typical fluid sampling intervals used by airlines are on-condition, C-check, and 2C-check. In 

addition, manufacturers recommend fluid sampling when a suspected hydraulic contamination 

may exist (for example, due to or indicated by excessive component wear or deterioration). 

(f) In summary. The FAA has identified an existing industry standard, NAS 1638, 

that defines fluid cleanliness levels as Classes 00 to 12; has defined NAS 1638 class 9, as the inservice limit, verified that the manufacturers already recommend these limits in their 

maintenance manuals including a sampling interval, and helped develop an SAE ARP document 

for sampling and testing techniques. The intent of the safety recommendation has been met and 

no regulatory corrective action is deemed necessary. 

(2) Certification of Hydraulic Lines for Temporary Repairs. The following was 

extracted from an FAA letter dated April 23, 1992, which addresses certification of hydraulic 

lines for use as temporary repairs. 

(a) It is not clear from the applicant's letter whether the hydraulic tubing 

assemblies being discussed are rigid or flexible, but it is assumed that the discussion involves 

various lengths of flexible hydraulic hose assemblies which can be joined together to replace 

140 
3/14/2000 AC 25-22 

rigid or flexible hydraulic lines that have become unserviceable. There is an FAA Technical 

Standard Order (TSO) which provides minimum airworthiness requirements for hydraulic hose 

assemblies intended for use on transport category airplanes. Technical Standard Order TSOC75, Hydraulic Hose Assemblies, and the accompanying Federal Aviation Standard should be 

considered. It should be noted that while the TSO provides an authorization for manufacturing a 

part that meets certain standards, it does not automatically approve the part for a specific 

installation.

(b) Section 43.13(a) requires that repairs must be accomplished in accordance with 

the manufacturers' maintenance manuals. However, air carriers are supposed to follow their Air 

Carrier Manual approved under part 121. 

(c) The only mechanism available for certification of these repair systems by a 

manufacturer other than the type certificate holder would be a Supplemental Type Certificate 

(STC). An STC, issued for each installation on every model of airplane on which the applicant 

desired approval, would be a daunting task. Major airframe manufacturers, because they have all 

the pertinent data regarding design parameters for the hydraulic systems on their airplanes, can 

issue appropriate instructions in their maintenance manuals for installation of temporary repairs. 

Airlines should use the manufacturer's manuals per § 43.13(a). 

(d) One of the ways a manufacturer, other than the type certificate holder who 

wishes to manufacture parts for a certificated airplane, can apply for a Parts Manufacturing 

Authorization (PMA) is to show that the parts it makes are identical to those produced by the 

original manufacturer. Unless the applicant has original drawings complete with all necessary 

standards and specifications, it is difficult to show identicality. The cognizant FAA 

Manufacturing Inspection District Office can provide insight into the appropriate PMA 

application procedures, reference Order 8110.42A. 

(3) Use of Hydraulic Lines for Temporary Repairs. The following was extracted from

an FAA memorandum dated January 25, 1988, which addresses the use of flexible hydraulic 

hoses as temporary repairs for unserviceable hydraulic tubing.

(a) The use of flexible hydraulic hoses as temporary repairs for unserviceable 

hydraulic tubing has been commonplace on transport category airplanes for a number of years. 

The lack of adverse service history indicates that this practice can be a viable method of repair, 

when used with the proper cautions. Strictly speaking, such substitution of hoses for rigid tubing 

would cause the airplane to be not in compliance with its type design; however, the provisions of 

Part 43 allow repairs and alterations to be made, provided the work is done in accordance with 

the methods, techniques, and practices in the current manufacturer's maintenance manual. With 

respect to multiple repairs, the FAA has concluded that they would be acceptable, provided that

each one is individually performed in accordance with an approved procedure, and, collectively, 

the repairs will not interfere with each other or with any structure, or with the operation or 

performance of any system on the airplane. 

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3/14/2000 AC 25-22 

(b) The key issue is the length of time that these temporary repairs may be allowed 

to exist on the airplane. Only by limiting the exposure can the repair be considered to be a safe 

alternative to the original design, much in the same manner that the Minimum Equipment List 

allows operation of the airplane for a short time with certain items of equipment inoperative. 

Since a temporary repair would not have been evaluated and certificated in the same manner as 

the original system, it would not be proper to allow the repair to become, in any way, a 

permanent part of the type design. It has been suggested by one manufacturer that these repairs 

be replaced at the next "C" check, which in some cases may be a year in the future. The FAA 

has concluded that this much time is unwarranted, and may seriously compromise the safety of 

the repair. An interval of 300 to 350 hours would, in most cases, allow for the ordering and 

delivery of any needed parts and the scheduling of the airplane for restoration of the system to 

the original configuration. Also, limiting the time in this manner will reduce the problem of 

multiple repairs on the system. Any time limit would be largely arbitrary, but this figure, which 

represents approximately a month of operations, appears to be a good compromise between the 

economic interests of the operators, and the need for a short exposure to the temporary repair. 

(4) Hydraulic System Certification Philosophy. The following was extracted from an 

FAA letter dated March 5, 1982, which addresses guidance concerning compliance with 

§ 25.1435(b)(2), 14 CFR. 

(a) The FAA has not been requested to accept endurance and reliability data based 

solely on flight test. This is because it is impractical, within a given certification time frame, to 

conduct the number of test cycles required to comply with § 25.1435(b)(2). The FAA would 

have no objection to using flight test data in addition to simulated functional endurance and 

reliability data, provided the manufacturer establishes that the flights are representative of

expected service cycles.

(b) It is normal practice, for certification purposes, to comply with the 

requirements of § 25.1435(b)(2) by the use of simulators such as an "iron bird." The simulators 

are intended to be representative of the aircraft systems. Mounting brackets, hydraulic and 

electrical system flight controls, and control surfaces are all the same as the actual airplane. For 

practical purposes, simulation may be broken into functional subsystems, i.e., wing, rudder, gear, 

etc. Deviations from actual simulation have been allowed where the deviation will not interfere 

with the intent of establishing endurance and reliability. Deviation cases have to be ruled on in a 

case by case basis. 

(c) The FAA would not accept compliance with § 25.1435(b)(2) based on only 

vendor component tests and flight tests. Functional tests and endurance tests, either using an 

airplane or mock-up (iron bird, simulator) at the system or subsystem level, are required. 

Functional tests are required prior to first flight test. 

(d) The manufacturer will develop load data through analysis, wind tunnel data, 

flight test experience, and/or any combination of these (spoilers and slots require flight test data, 

§ 25.459). This load data is submitted to the FAA in the form of a loads document. The loads 

document is usually approved by FAA designated engineering representatives (DER) and 

142 
3/14/2000 AC 25-22 

submitted to the FAA for a spot check review when and where it is considered necessary. Flight 

test load verification is not required when the methods used in determining those loading 

conditions are shown to be reliable (§ 25.301(b)). Most manufacturers go beyond FAA 

requirements and check the loads with strain gauges during flight tests. 

(e) The airframe manufacturer (type design applicant) identifies the loading 

spectrum (amplitude and cycles) for the hydraulic systems based on the loads document and 

expected service. The FAA will be involved only indirectly in most cases, through the loads 

DER’s. 

(f) The manufacturer subjects components and systems or the aircraft to the 

selected load/cycle spectrum in demonstrating compliance with the endurance requirements of 

§ 25.1435(b)(2). Normally, vendor testing will be used for verification of the components for 

endurance as well as other requirements, such as environmental qualifications, fatigue, etc. The 

airframe manufacturer's simulator (iron bird) or simulators will be used for endurance tests and 

functional tests. The qualification reports documenting these tests are reviewed by FAA DER’s 

and approved or submitted to the FAA with a recommendation for approval. This option is 

designated by the FAA. 

(5) Interpretation of § 25.1435(a)(4) of 14 CFR. The following was extracted from an 

FAA memorandum dated September 14, 1978, which addresses interpretation of 

§ 25.1435(a)(4), Amendment 25-41. 

 (a) Section 25.1435(a)(4)(i) states: "There must be a means to keep hydraulic 

system pressures within +/- 10 percent of the pressure at the discharge of the pump outlet or the 

transient pressure dampening device, if provided." 

(b) This rule was originally established to cover pump ripple pressure variations. 

It can be seen that, under extreme system loading (high flow) conditions, the pump average 

discharge pressure can decrease as much as 25 percent. Apparently, the system pressure 

tolerance of +/- 10 percent mentioned above would then apply to this reduced pump average 

discharge pressure, decreased due to high flow conditions. 

 (c) Section 25.1435(a)(4)(ii) states: "Except as provided in Paragraph (a)(7) of this 

section, system pressures will not exceed 125 percent of the design operating pressure, excluding 

pressure at the pump outlet or dampening device. Design operating pressure is defined as the 

maximum steady operating pressure." 

(d) The FAA has determined that Paragraph (4)(i) covers pressure variations due 

to pump ripple and high flow conditions while Paragraph (4)(ii) covers variations due to pressure 

transients. Note the use of "pump average discharge pressure" in (4)(i) and "design operating 

pressure" in (4)(ii). 

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3/14/2000 AC 25-22 

 e. References. The addresses for ordering the latest revision of advisory circulars, 

technical standard orders, and other referenced documents listed below can be found in the 

Appendix to this AC. 

AC 25.1309-1A - System Design and Analysis. 

AC 120-42A - Extended Range Operation with Two Engine Airplanes. 

AC 20-128 - Design Considerations for Minimizing Hazards Caused by Uncontained

Turbine Engine and Auxiliary Power Unit Rotor and Fan Blade Failures.

TSO-C47 - Pressure Instruments-Fuel, Oil, and Hydraulic. 

TSO-C75 - Hydraulic Hose Assemblies. 

SAE AS-595B - Civil Type Aircraft and Variable Delivery Hydraulic Pump. 

SAE ARP-763 - Accumulators, Ground, Hydropneumatic Pressure. 

SAE AIR-4150 - Inspection of In-Service Airborne Accumulators. 

SAE AIR-1047C - A Guide for Selection of Quick Disconnect Couplings for Aerospace 

Fluid Systems. 

SAE ARP-1709 - Coupling Assembly, Hydraulic Self Sealing, Quick Disconnect. 

SAE AIR-786A - Elastomer Compatibility Considerations Relative to O-Ring and 

Sealant Selection. 

SAE AS-1241B - Fire Resistant Phosphate Ester Hydraulic Fluid for Aircraft. 

SAE AIR-1116 - Fluid Properties. 

SAE ARP-1084 - Hydraulic External Leakage for In-service Components. 

SAE AIR-1362 - Physical Properties of Hydraulic Fluids. 

SAE ARP-1832 - Color Identification for O-Ring Seals. 

SAE AIR-737E - Hydraulic and Pneumatic Specifications and Standards. 

SAE ARP-994 - Design of Tubing Installations for Aerospace Fluid Power Systems. 

SAE AIR-4003 - Turbine Engine Containment. 

SAE AIR-1083B - Airborne Hydraulic and Control System Survivability for Military 

Aircraft. 

SAE AIR-1899 - Aircraft Hydraulic System Characteristics. 

SAE AIR-1918 - Comparison of Hydraulic System Cleanliness Procedures and 

Requirements for Ten Aerospace Companies. 

SAE ARP-24B - Determination of Hydraulic Pressure Drop. 

SAE ARP-490E - Electro-Hydraulic Servovalves. 

SAE ARP-1280A - Aerospace Application Guide for Hydraulic Power Transfer Units.

SAE ARP-1281B - Actuators: Aircraft Flight Controls, Power Operated, Hydraulic, 

General Specification. 

SAE ARP-4379 - Accumulator, Hydraulic, Cylindrical Aircraft. 

 SAE ARP-4752 - Aerospace - Design and Installation of Commercial Transport Aircraft 

Hydraulic Systems. 

FAA Order 8000.40D - Maintenance of Pressure Cylinders in Use as Aircraft 

Equipment. 

61. SECTION 25.1438 PRESSURIZATION AND PNEUMATIC SYSTEMS.

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3/14/2000 AC 25-22 

 a. Rule Text. 

(a) Pressurization system elements must be burst pressure tested to 2.0 times, and 

proof pressure tested to 1.5 times, the maximum normal operating pressure. 

(b) Pneumatic system elements must be burst pressure tested to 3.0 times, and 

proof pressure tested to 1.5 times, the maximum normal operating pressure. 

(c) An analysis, or a combination of analysis and test, may be substituted for any 

test required by paragraph (a) or (b) of this section if the Administrator finds it 

equivalent to the required test. 

[Amdt. 25-41, 42 FR 36971, Jul. 18, 1977]

b. Intent of Rule. This rule provides standards for the tests that must be conducted to 

demonstrate that pneumatic and pressurization ducting and components will not fail in normal 

operation. The rule was originated to address the pneumatic system, from the engine bleed 

port(s) to the pressure regulating and/or shut-off valve, and the pressurization system, which 

comprises all ducting and components of the air distribution system downstream of the abovementioned valve. 

 c. Background. This subject was not addressed in either part 4b of the Civil Air 

Regulations (CAR) or part 25 of Title 14, Code of Federal Regulations (14 CFR) when it was 

originally codified. The reason for proposing the new standards was: "Components (such as 

ducts and couplings) of pressurization and pneumatic systems have failed at an unacceptable rate 

in service. The proposed standards for these components have been effective in preventing 

design deficiencies in the past." 

 (1) Harmonization. This regulation is the subject of a Federal Aviation 

Regulations/Joint Aviation Requirements (FAR/JAR) harmonization effort under the Aviation 

Rulemaking Advisory Committee (ARAC). The ARAC working group may recommend a 

revised § 25.1438/JAR 25.1438 and JAR 25X1436, and may develop an accompanying new 

advisory circular. 

d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Pneumatic Versus Pressurization System Identification: The following policy was 

extracted from an FAA letter dated October 19, 1993, in response to questions regarding 

§ 24.1438 and the difference between pressurization system elements and pneumatic system

elements. 

(a) Section 25.1438(a) states: "Pressurization system elements must be burst 

pressure tested to 2.0 times, and proof pressure tested to 1.5 times, the maximum normal 

operating pressure," and § 25.1438(b) states: "Pneumatic system elements must be burst 

pressure tested to 3.0 times, and proof pressure tested to 1.5 times, the maximum normal 

operating pressure." 

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3/14/2000 AC 25-22 

(b) The FAA considers the pneumatic system to include all air supply elements 

from the bleed port on the engine to the pressure regulating and shutoff valve. Pressurization 

system elements are all other elements in the air distribution system. 

(c) The burst pressure test requirement for pneumatic and pressure system

elements has been used to demonstrate that high pressure air will be contained within the system

or element. The element is not required to function during or after the burst pressure test, but 

pressurized air must be contained within the element or system. The proof pressure test is used 

to demonstrate that pneumatic and pressure system elements can function after a higher than 

normal pressure (1.5 times the maximum normal operating pressure) is introduced into the 

element or system. 

(d) Section 25.1438(c) allows the applicant to use an analysis as a method of 

showing compliance to the burst and proof pressure test requirement in §§ 25.1438(a) & (b). 

 e. References. None. 

62. SECTION 25.1439 PROTECTIVE BREATHING EQUIPMENT.

a. Rule Text.

(a) If there is a class A, B, or E cargo compartment, protective breathing 

equipment must be installed for the use of appropriate crewmembers. In addition, 

protective breathing equipment must be installed in each isolated separate 

compartment in the airplane, including upper and lower lobe galleys, in which 

crewmember occupancy is permitted during flight for the maximum number of 

crewmembers expected to be in the area during any operation. 

(b) For protective breathing equipment required by paragraph (a) of this section 

or by any operating rule of this chapter, the following apply: 

(1) The equipment must be designed to protect the flight crew from smoke, carbon 

dioxide, and other harmful gases while on flight deck duty and while combating 

fires in cargo compartments. 

(2) The equipment must include- 

(i) Masks covering the eyes, nose, and mouth; or 

(ii) Masks covering the nose and mouth, plus accessory equipment to cover the 

eyes. 

(3) The equipment, while in use, must allow the flight crew to use the radio 

equipment and to communicate with each other, while at their assigned duty 

stations. 

(4) The part of the equipment protecting the eyes may not cause any appreciable 

adverse effect on vision and must allow corrective glasses to be worn. 

(5) The equipment must supply protective oxygen of 15 minutes duration per 

crewmember at a pressure altitude of 8,000 feet with a respiratory minute volume 

of 30 liters per minute BTPD. If a demand oxygen system is used, a supply of 300 

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liters of free oxygen at 70 o F. and 760 mm. Hg. pressure is considered to be of 

15-minute duration at the prescribed altitude and minute volume. If a continuous 

flow protective breathing system is used (including a mask with a standard 

rebreather bag) a flow rate of 60 liters per minute at 8,000 feet (45 liters per 

minute at sea level) and a supply of 600 liters of free oxygen at 70 o F. and 760 

mm. Hg. pressure is considered to be of 15-minute duration at the prescribed 

altitude and minute volume. BTPD refers to body temperature conditions (that is, 

37o C., at ambient pressure, dry). 

(6) The equipment must meet the requirements of paragraphs (b) and (c) of 

§ 25.1441. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 

41 FR 55468, Dec. 20, 1976]

b. Intent of Rule. The intent of this rule is for the protective breathing equipment (PBE) 

provisions to protect crewmembers from the effects of hazardous gasses and smoke, either while 

on flight deck duty or combating a fire, if there is a class A, B, or E cargo compartment or an 

isolated compartment in which crewmembers are permitted access during flight. This rule does 

not address passenger protective breathing equipment. Section 25.1439(b) contains wording that 

requires PBE if required "by any operating rule of this chapter. . . " There are two technical 

standard orders (TSO) providing standards for approval of PBE. Technical Standard Order, 

TSO-C99, provides standards for use by flight crewmembers while on flight deck duty. Per the 

TSO, the PBE must provide eye protection in addition to preventing inspiration of smoke/fumes. 

The flightcrew supplemental oxygen system can provide this via a full face mask with a setting 

that provides a positive pressure differential between the mask and ambient pressure in the 

cockpit to prevent smoke/fumes from entering the mask. If the crew supplemental oxygen 

system is equipped with an oro-nasal type mask (covers nose and mouth only) rather than a full 

mask, the eye protection can be achieved via a pair of goggles located within easy reach. This 

type of equipment is intended for use by flight crewmembers at their stations for a prolonged 

length of time. It would not generally be acceptable for firefighting since the crewmember's

ability to reach the fire would be limited by the amount of extra oxygen tubing provided by the 

stationary system. TSO-C116 provides standards for PBE to be used by crewmembers while 

locating and fighting a fire. Per the TSO, it also provides head and shoulder protection from

"drippings," and typically provides protective oxygen for 15 minutes. This type of PBE would 

not be suitable for use by the flightcrew at their stations. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Many of the 

requirements in section 4b.651 of the CAR were carried over essentially unchanged to § 25.1439 

of 14 CFR. Part 25 standards state that PBE must be installed “if there is a class A, B, or E cargo 

compartment.” 

 (1) Amendment 25-38 (December 20, 1976) added the requirement that PBE be 

installed in each isolated separate compartment, including upper and lower lobe galleys, in which 

crewmember occupancy is permitted during flight for the maximum number of crewmembers 

expected to be in the area. 

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 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Protective Breathing Equipment Usage When Approved Under TSO-C99 and 

TSO-C116. The following policy, extracted from an FAA letter dated May 13, 1993, clarifies 

compliance with the requirements of § 25.1439 using equipment of the "smoke hood" variety 

approved under TSO-C116. 

(a) The FAA letter clarified whether smoke hood type PBE would be acceptable 

for meeting the requirements of § 25.1439. The FAA is not aware of any smoke hood type PBE 

being approved as meeting the requirements of § 25.1439(b)(3), which relates to the use of radio 

equipment by the flight crew. The FAA would consider a smoke hood type PBE if compliance 

with § 25.1439(b)(3) is demonstrated. 

(b) The FAA letter also clarified that a finding of equivalent safety would be 

required if an applicant wishes to use a “breathable gas,” in lieu of providing oxygen as specified 

in § 25.1439(b)(5). 

(c) The FAA letter also addressed § 25.1439(b)(6), which specifies that there must 

be a means provided to allow the crew to readily determine the quantity of oxygen available. 

PBE approved in accordance with TSO-C116 may or may not have such features. The FAA 

does not grant exemptions for the cases where the PBE is used by the flight crewmembers while 

on flight deck duty. When the PBE is intended for use in locating and combating a fire, the FAA 

will approve equipment which, when the container is intact, is considered to have a full charge of

oxygen or, for chemical oxygen generation, the canister has not been activated. The container is 

"tamper evident" and the crewmember can readily determine that the equipment is fully charged

and ready for use. This PBE would not be suitable for approval in meeting the flight 

crewmember requirements addressed by § 25.1439. 

(d) The letter provided an FAA position on PBE requirements for the first and 

second observers. The FAA treats the first observer and second observer differently. The first 

observer must have the same protection, i.e., the same equipment, as the flight crewmembers. 

The second observer may be supplied with oxygen equipment suitable for passengers. Again, 

the first observer must have the same protection (including the ability to communicate on the 

crewmember interphone) that is provided for the flight crewmembers; the second observer may 

be provided with passenger-type oxygen equipment. 

(2) Protective Breathing Equipment Serviceability. The following policy was extracted 

from an FAA memorandum dated March 9, 1990, to provide guidance regarding the approval of 

PBE which met the requirements of § 121.337 and FAA Action Notice 8150.2. 

(a) Section 121.337 was revised in 1987 and, in part, incorporated the intent of

§ 25.1439 and expanded on those requirements. This revision resulted in a new type of PBE for 

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use by any crewmember while combating fires on board an airplane. The approval basis was 

contained in Action Notice A8150.2, dated September 1, 1987, in advance of TSO-C116. 

(b) Section 121.337 clearly intends that any crewmember PBE be immediately 

ready to perform its intended function during an emergency. In fact, the rule requires that each 

certificate holder's operations manual designate at least one crewmember to check that each PBE 

unit is properly stowed and serviceable prior to the first-flight of the day. 

(c) The approval basis (TSO-C116) for crewmember PBE requires that the 

equipment has a means to indicate the serviceability of the unit in its stowed condition. Some

manufacturers of the crewmember PBE designs approved by the FAA have elected to use a 

vacuum sealed envelope as a means of protecting the PBE, and the loss of vacuum has been 

accepted as an indicator that the PBE is not serviceable. 

(d) One manufacturer recommended that when the vacuum is lost the PBE need 

not be replaced immediately, since ambient moisture will not deactivate the potassium

superoxide in the atmosphere regenerating system for a number of days. This may well be 

correct, but it is only one consideration. Since the cause of the loss of vacuum will most likely 

not be known to the crewmember making the inspection, and the extent of damage to the PBE 

itself may not be evident, the PBE should be considered not serviceable. Another consideration 

is that a design, which includes a protective envelope material that can readily be damaged, may 

not perform its intended function. 

(e) Based on the above information, ACO's should ensure that the manufacturer 

submits operating, installation, and maintenance instructions with limitations, warnings, or 

cautions for FAA approval. The information being provided the end user should be in accord 

with the safety intent of § 121.337. 

(3) Protective Breathing Equipment on the Flight Deck. The following policy was 

extracted from an FAA memorandum dated December 19, 1988, and provided additional 

guidance on protective breathing equipment for use on the flight deck. 

(a) Section 121.337 specifies provisions for protective breathing equipment with 

the airplane both pressurized and unpressurized. Section 25.1439 makes no such distinction. 

The part 25 rule specifies that protective breathing equipment must be installed as part of the 

type design if certain types of cargo compartments or isolated, occupied compartments are 

installed. If these specified compartments are not installed on the airplane, then protective 

breathing equipment is not required as part of the type design. 

(b) For an unpressurized part 25 airplane certified without these specified 

compartments and operated under Part 121, however, the decision to require protective breathing 

equipment should be based on an evaluation of the hazard. Section 25.831(d) states that if 

accumulation of hazardous quantities of smoke in the cockpit area is reasonably probable, smoke 

evacuation must be readily accomplished, starting with full pressurization and without 

depressurizing beyond safe limits. This capability must be demonstrated for certification under 

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any normal operating condition, including unpressurized flight; and smoke clearance must be 

accomplished within three minutes (Advisory Circular 25-9A, or latest revision, explains a test 

procedure in detail). Section 25.831 also specifies limits for carbon dioxide and carbon 

monoxide, but does not define unsafe levels of any other gas, vapor, or fume. Since the toxicity 

of smoke or fumes generated by a fire is unknown, the FAA has required that there be essentially 

no penetration of smoke into an occupied compartment. Since it is likely that smoke of unknown 

composition would be generated within the cockpit, and may be present for up to three minutes 

prior to its removal, installation of protective breathing equipment would be appropriate, unless 

the applicant can show that it is not necessary. 

(4) Definition of Terms Used in § 25.1439. The following policy was extracted from

an FAA memorandum dated February 26, 1985, that responded to a request for the meaning of 

the words "isolated separate compartment," as used in § 25.1439. 

(a) The preamble to the Notice 75-10, item 2-91, read: "The proposal would 

require protective breathing equipment for crewmembers expected in isolated areas." Item 2-91 

was based on the First Biennial Airworthiness Review of 1974-1975, proposal No. 812 that read: 

"(a) Protective breathing equipment must be installed for each required crewmember in isolated 

separate compartments, such as upper or lower lobe galleys, in which occupancy is permitted 

during flight." 

(b) Based on the intent of the background quoted above, the FAA concluded that 

the crew rest area, not being part of the main cabin, is similar in location to a lower lobe galley 

and, therefore, is an isolated separate compartment. However, as crewmembers are not required 

in the crew rest area except to fight a fire, only sufficient protective breathing equipment need be 

provided in the crew rest area for that function. It is also noted that, as long as the lavatory is 

located next to the flight deck door, there is little chance that a crewmember may be isolated 

there by a fire. The crew lavatory need not be considered isolated. 

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 e. References. The address for ordering the latest revision of the technical standard orders 

listed below can be found in the Appendix to this AC. 

TSO-C99, Protective Breathing Equipment. 

 TSO-C116, Crewmember Protective Breathing Equipment. 

63. SECTION 25.1441 - OXYGEN EQUIPMENT AND SUPPLY.

 a. Rule Text. 

(a) If certification with supplemental oxygen equipment is requested, the 

equipment must meet the requirements of this section and 25.1443 through 

25.1453. 

(b) The oxygen system must be free from hazards in itself, in its method of 

operation, and in its effect upon other components. 

(c) There must be a means to allow the crew to readily determine, during flight, 

the quantity of oxygen available in each source of supply. 

(d) The oxygen flow rate and the oxygen equipment for airplanes for which 

certification for operation above 40,000 feet is requested must be approved. 

b. Intent of Rule. This rule is intended to ensure that, if supplemental oxygen equipment is 

to be included in the type design of an airplane, the oxygen dispensing equipment will protect 

passengers and crewmembers from the effects of hypoxia. Many of the oxygen system

requirements are addressed in §§ 25.1443 through 25.1453, which are incorporated by reference. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in section 4b.651(a) of the CAR for oxygen systems were carried over essentially unchanged to 

§ 25.1441 of 14 CFR.

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Compliance with §§ 25.1309(c) and 25.1441(c). The following policy was 

extracted from an FAA memorandum dated February 23, 1994, and provides an FAA position on 

the requirements of § 25.1309(c) as related to a Passenger Gaseous Oxygen System that could 

erroneously indicate oxygen is available. 

(a) Section 25.1309(c) states: "Warning information must be provided to alert the 

crew to unsafe system operating conditions, and to enable them to take appropriate corrective 

action. Systems, controls, and associated monitoring and warning means must be designed to 

minimize crew errors which could create additional hazards." Section 25.1441(c) states: "There 

must be a means to allow the crew to readily determine, during flight, the quantity of oxygen 

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available in each source of supply." An applicant proposed a passenger oxygen system where 

the means is provided through an indication available to the flight crew of the oxygen pressure in 

the oxygen system plumbing between the oxygen bottles and the flow control valves. There are 

several bottles, each with its own shut-off valve, manifolded together. If, during maintenance, 

the valves on all the bottles are turned off and left in that position, but the tubing between the 

bottles and the flow control valve is not disturbed, the pressure in the lines will continue to read 

the same value that would be measured if the bottle valves were open. Therefore, the reading in 

the cockpit will erroneously indicate that oxygen is available. A concern arises because the 

flightcrew will dispatch believing that passenger oxygen is available. If a depressurization event 

occurs, the passengers will not have supplemental oxygen available as required by the 

regulations. Passengers will be at risk during the emergency descent due to lack of available 

oxygen. 

(b) The non-availability of passenger supplemental oxygen presents a hazard if a 

depressurization occurs. In most cases of depressurization, the flightcrew executes an 

emergency descent and the cabin altitude does not exceed safe limits. This is not, of course, 

always the case; loss of consciousness or more serious injury, and even death, could occur for 

some passengers if oxygen is not available after depressurization. There is usually no risk to the 

airplane itself because the flightcrew has a separate oxygen system and, assuming they act 

properly, the airplane will quickly descend to a safe altitude and proceed to the nearest airport. 

For these reasons, passenger supplemental oxygen being unavailable is not an immediate threat 

to continued safe flight and landing. However, one of the design principles suggested in 

Advisory Circular (AC) 25.1309-1A for ensuring fail-safe design concepts is error tolerance to 

allow for possible adverse effects of foreseeable errors during the airplane's design, test, 

manufacture, operation, and maintenance. This system does not appear to adhere to that 

principle. In addition, the AC also notes that a warning is required for a failure which makes it 

necessary for the flightcrew to make an unscheduled landing to reduce exposure to a more 

hazardous failure condition that would result from subsequent failures or operational conditions, 

or if the failure must be corrected before a subsequent flight. In conclusion, a warning to the 

flightcrew for non-availability of passenger oxygen is the type of warning addressed in the first 

sentence of § 25.1309(c). 

(c) The system installed to provide the crew with an indication of the amount of 

oxygen available does not appear to meet the requirement contained in the second sentence of 

§ 25.1309(c). The false indication of oxygen pressure resulting from the oxygen trapped 

between the shut off valves and the flow control valves will continue indefinitely, leading to 

dispatch after dispatch with no means to detect the problem until either a depressurization occurs 

or some maintenance action leads to the discovery that the oxygen bottle valves are closed. If 

the crew were aware of the situation, they would not dispatch or, if the problem were discovered 

in flight, they would immediately divert to an alternate airport. Clearly, the indication system is 

not designed to minimize crew errors that could create additional hazards.

(d) It is not acceptable to use probability of the event (depressurization) to 

demonstrate the probability of the unsafe condition (decompression combined with passenger 

oxygen being unavailable). In this case, the unsafe condition is not catastrophic so it is not 

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necessary to show that it is extremely improbable. As noted in the earlier advisory material 

regarding System Design Analysis (AC 25.1309-1), "If a quantitative analysis is used to help 

show compliance with Federal Aviation Regulations for equipment which is installed and 

required only for a specific operating condition for which the airplane is thereby approved, credit 

may not be taken for the fact that the operating condition does not always exist." While this note 

does not appear in AC 25.1309-1A, the basic philosophy has not changed. 

(2) Guidance, Availability of Flightcrew Oxygen. The following policy was extracted 

from an FAA memorandum dated May 11, 1992, in response to an FAA Safety 

Recommendation regarding annunciation that oxygen is available for use by the flightcrew when 

in actuality there is no oxygen available. 

(a) On a recent twin-engine transport airplane flight, the crew elected to divert 

after the first officer donned his oxygen mask and discovered he had no oxygen available. After 

landing, the crew oxygen bottle valve was found turned to the closed position. Prior to the flight, 

maintenance had started to replace the bottle due to low pressure, and the valve located at the 

bottle was turned off. When the bottle pressure was found to be acceptable, the bottle was 

reinstalled but the valve was left turned off and safety wired in that position. The maintenance 

crew checked the bottle pressure on the airplane Engine Indication and Crew Alerting System

(EICAS) status page and because the EICAS displays the pressure in the line downstream of the 

valve, the residual pressure in the line resulted in a displayed pressure that was within limits. 

(b) The existing system on this airplane measures the oxygen pressure in the line 

to the flightcrew regulator. If the shutoff valve at the pressure bottle is shut off, there is 

sufficient pressure remaining in the line to indicate a pressure high enough for dispatch. Even 

when the crew checks the masks, only a small amount of oxygen is allowed to escape, and there 

are currently no means to determine that the valve is off. Further, the oxygen pressure, and 

therefore the quantity, is observable on EICAS only when the crew selects the status page. 

(c) The FAA determined that this is unacceptable for the following reasons: 

 1. Section 25.1441(c) requires that the crew be able to determine the quantity 

of oxygen available. When the valve is in the "OFF" position, the oxygen is not available, but 

the indication on EICAS (the residual pressure in the line) is that oxygen is available if needed. 

The pressure measurement gives information about quantity only when the valve is open. If the 

valve is inadvertently left closed, the information provided is misleading. 

 2. Section 25.1309(c) states: “Warning information must be provided to 

alert the crew to unsafe system operating conditions, and to enable them to take appropriate 

corrective action. Systems, controls, and associated monitoring and warning means must be 

designed to minimize crew errors which could create additional hazards.” It is clear that the 

existing design gives misleading information to the flightcrew, which could lead them into an 

unsafe operation. If depressurization takes place at a significant altitude, and oxygen is not 

available to the crew, an unsafe condition exists. 

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 3. Further, § 121.333(c)(4) states: "Before the takeoff of a flight, each flight

crewmember shall personally preflight his oxygen equipment to insure that the oxygen mask is 

functioning, fitted properly, and connected to appropriate supply terminals, and that the oxygen 

supply and pressure are adequate for use." In theory, adherence to this rule should ensure that 

the oxygen supply is available. 

(d) In order for the scenario discussed above to be a hazard, four separate events 

would have to occur: 

 1 There would have to be maintenance performed on the airplane oxygen 

system that required the valve, located on the oxygen bottle, to be turned off for some reason. 

The airplane maintenance manual contains both visual and operational checks to ensure that the 

oxygen valve is open and the pressure is adequate. These procedures were not followed in the 

case of interest because the bottle was never removed. 

 2 The flightcrew check of their oxygen equipment would have to fail to 

indicate that oxygen is not available. This is possible with the present preflight procedure, but 

the procedure is being changed. 

 3 A decompression would have to occur that required the use of oxygen by 

the flightcrew. 

 4 Event 3 would have to occur before the flightcrew had attempted to use 

supplemental oxygen (which would demonstrate no oxygen available) or checked the EICAS 

Status Page, which would indicate low oxygen pressure. In either case, airline procedures and 

Federal Aviation Regulations require that the crew divert to fix the problem. 

(e) The manufacturer of the airplane involved in the incident which prompted this 

memorandum is changing their operational procedures, which are used by the airlines for their 

Operations Manuals, to describe an acceptable procedure to verify that adequate oxygen is 

available at the pilots mask. This procedure involves two separate flow tests (Normal and 100 

percent), followed by a pressure check on EICAS. As all flightcrew masks must be tested (at 

least two on any transport category airplane), there will be ample tests and pressure checks to 

identify a problem prior to dispatch. Based on the above considerations, the FAA has 

determined an adequate level of safety is achieved through the existing and proposed procedures 

with the certificated design. However, this mechanization is not considered to meet the 

requirements of § 25.1309(c). 

(f) Flightcrew oxygen systems should be reviewed to ensure that on future 

certification programs, a system design that can lead to misinformation is not allowed. 

(3) Oxygen Installation for Medical Use. The following policy was extracted from an 

FAA facsimile message dated January 3, 1991, that was written in response to a request for 

policy relative to the installation of shutoff valves, pressure relief devices and overboard vent 

lines installed on oxygen bottles for medical use. 

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(a) There is no current written FAA policy relative to the installation of

manifolded oxygen systems other than Civil Aeronautics Manual (CAM) 4b.651-1. CAM 

4b.651-1 recommends that low pressure oxygen systems have a pressure relief device to prevent 

over pressurization during the filling operation. 

(b) Department of Transportation (DOT) regulation (49 CFR, § 173.34(d)) 

requires that all pressure cylinders be provided with a safety device (a rupture disc) to prevent 

explosion of the bottle if the bottle is subjected to over-pressure or excessive heat. The pressure 

relief device may be installed on the bottle or on the valve of the bottle (on the pressure side of 

the valve). 

(c) To meet the requirements of § 25.1441(b) and § 1451(c) (note: Amendment 

25-72 deleted § 25.1451 and moved the requirements to § 25.869(c)), aircraft manufacturers 

have been installing oxygen systems that have a pressure relief that vents through a vent line, 

normally stainless steel, to the outside of the aircraft. A colored blowout plug is installed at the 

end of the vent line on the surface of the aircraft to indicate if the rupture disc has blown. 

(d) The inquiring FAA office indicated that they may have approved single high 

pressure bottle (air ambulance) installations that do not have the overboard drain line. Further, 

these single bottle installations are filled off the airplane. There may be other installations 

without a vent line because there doesn't appear to be a written policy on this subject. 

(e) To determine if a vent line is needed to comply with §§ 25.1441(b) and 

§ 25.869(c)(3) (formerly 25.1451(c)), the consequences of the pressure relief device venting the 

total contents of the bottle or bottles (if manifolded together) into the compartment should be 

considered, i.e., overpressurizing the compartment; exposure to grease, flammable fluids and 

ignition sources. 

(f) All the installations, which the FAA is aware of, have a high pressure shutoff 

valve located on the bottle to isolate the bottle from the oxygen system. On a manifolded 

system, these valves normally are lockwired in the open position. However, the bottle shutoff 

valve is downstream of the pressure relief device in order that the pressure relief device will 

protect the bottle from exploding if the bottle is isolated (valve shut) from the system and is 

exposed to heat or over- pressurization. 

(g) The FAA is not aware of any objection to having a common high pressure 

manifold without the bottle shutoff valves as long as there is a manifold system shutoff valve and 

a pressure relief device between the bottles and the shutoff valve. 

(h) One objection might be exposing the airplane and maintenance personnel to the 

hazards of a fully pressurized vented bottle or bottles if someone tries to remove a bottle from

the system. The total system has to be depressurized before the bottle can be removed as 

opposed to the manifold system where each bottle shutoff valve can be closed, and an individual 

bottle with its valve can be removed. 

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(i) In general, industry practice has been to separate and isolate the oxygen system

from sources of ignition. Ignition sources include pumps, motors, and electrical equipment 

associated with the medical installations as well as the aircraft equipment. Further, the 

manufacturers normally install a pressure relief device with an overboard vent line. 

 e. Reference. The address for ordering the document listed below can be found in the 

Appendix to this AC. 

The Society of Automotive Engineers (SAE) has aerospace committees that prepare 

documents reflecting industry standards and practices. The SAE A-10 Committee, Aircraft

Oxygen Equipment, has prepared a handbook, “1992 SAE Aircraft Oxygen Equipment 

Handbook,” containing all Aerospace Standards, Aerospace Recommended Practices, and 

Aerospace Information Reports related to oxygen equipment published by SAE. While not 

regulatory in nature, this handbook contains information that may be of interest to an applicant. 

64. SECTION 25.1443 - MINIMUM MASS FLOW OF SUPPLEMENTAL OXYGEN.

 a. Rule Text. 

(a) If continuous flow equipment is installed for use by flight crewmembers, the 

minimum mass flow of supplemental oxygen required for each crewmember may 

not be less than the flow required to maintain, during inspiration, a mean 

tracheal oxygen partial pressure of 149 mm. Hg. when breathing 15 liters per 

minute, BTPS, and with a maximum tidal volume of 700 cc. with a constant time 

interval between respirations. 

(b) If demand equipment is installed for use by flight crewmembers, the minimum 

mass flow of supplemental oxygen required for each crewmember may not be less 

than the flow required to maintain, during inspiration, a mean tracheal oxygen 

partial pressure of 122 mm. Hg., up to and including a cabin pressure altitude of 

35,000 feet, and 95 percent oxygen between cabin pressure altitudes of 35,000 

and 40,000 feet, when breathing 20 liters per minute BTPS. In addition, there 

must be means to allow the crew to use undiluted oxygen at their discretion. 

(c) For passengers and cabin attendants, the minimum mass flow of supplemental 

oxygen required for each person at various cabin pressure altitudes may not be 

less than the flow required to maintain, during inspiration and while using the 

oxygen equipment (including masks) provided, the following mean tracheal 

oxygen partial pressures: 

(1) At cabin pressure altitudes above 10,000 feet up to and including 18,500 feet, 

a mean tracheal oxygen partial pressure of 100 mm. Hg. when breathing 15 liters 

per minute, BTPS, and with a tidal volume of 700 cc. with a constant time interval 

between respirations. 

(2) At cabin pressure altitudes above 18,500 feet up to and including 40,000 feet, 

a mean tracheal oxygen partial pressure of 83.8 mm. Hg. when breathing 30 liters 

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per minute, BTPS, and with a tidal volume of 1,100 cc. with a constant time 

interval between respirations. 

(d) If first-aid oxygen equipment is installed, the minimum mass flow of oxygen to 

each user may not be less than four liters per minute, STPD. However, there may 

be a means to decrease this flow to not less than two liters per minute, STPD, at 

any cabin altitude. The quantity of oxygen required is based upon an average flow 

rate of three liters per minute per person for whom first-aid oxygen is required. 

(e) If portable oxygen equipment is installed for use by crewmembers, the 

minimum mass flow of supplemental oxygen is the same as specified in paragraph 

(a) or (b) of this section, whichever is applicable.

b. Intent of Rule. The requirements of this rule provide standards to ensure that oxygen 

dispensing equipment will protect passengers and crewmembers from the effects of hypoxia. 

The requirements for flight crewmembers are provided in terms of mean tracheal partial pressure 

of oxygen, which is not easily measured in certification testing. Dispensing equipment is, 

however, designed to provide the required oxygen flow. This equipment is, in most cases, 

approved through the technical standard order system. Requirements for passenger and cabin 

crewmember equipment are expressed in terms of oxygen flow rates. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in § 4b.651(b) of the CAR were carried over essentially unchanged to § 25.1443 of 14 CFR.

 d. Policy/Compliance Methods. There is no written policy or guidance in our files.

 e. Reference. The address for ordering the document listed below can be found in the 

Appendix to this AC. 

The Society of Automotive Engineers (SAE) has aerospace committees that prepare 

documents reflecting industry standards and practices. The SAE A-10 Committee, Aircraft

Oxygen Equipment, has prepared a handbook containing all Aerospace Standards, Aerospace 

Recommended Practices, and Aerospace Information Reports related to oxygen equipment 

published by SAE. While not regulatory in nature, this handbook contains valuable information 

that may be of interest to an applicant. 

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65. SECTION 25.1445 - EQUIPMENT STANDARDS FOR THE OXYGEN DISTRIBUTING 

SYSTEM.

a. Rule Text.

(a) When oxygen is supplied to both crew and passengers, the distribution system 

must be designed for either- 

(1) A source of supply for the flight crew on duty and a separate source for the 

passengers and other crewmembers; or 

(2) A common source of supply with means to separately reserve the minimum 

supply required by the flight crew on duty. 

(b) Portable walk-around oxygen units of the continuous flow, diluter-demand, 

and straight demand kinds may be used to meet the crew or passenger breathing 

requirements.

b. Intent of Rule. This rule addresses the oxygen distribution system and does not address 

oxygen dispensing equipment (masks), the standards for which are addressed in § 25.1447. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in § 4b.651(c) of the CAR were carried over essentially unchanged to § 25.1445 of 14 CFR, 

except for the standards on potable walk-around units that are found in § 25.1445(b).

 d. Policy/Compliance Methods. There is no written policy or guidance in our files. 

 e. Reference. The address for ordering the document listed below can be found in the 

Appendix to this AC. 

The Society of Automotive Engineers (SAE) has aerospace committees that prepare 

documents reflecting industry standards and practices. The SAE A-10 Committee, Aircraft

Oxygen Equipment, has prepared a handbook, “1992 SAE Aircraft Oxygen Equipment 

Handbook,” containing all Aerospace Standards, Aerospace Recommended Practices, and 

Aerospace Information Reports related to oxygen equipment published by SAE. While not 

regulatory in nature, this handbook contains information that may be of interest to an applicant. 

66. SECTION 25.1447 - EQUIPMENT STANDARDS FOR OXYGEN DISPENSING UNITS.

 a. Rule Text. 

If oxygen dispensing units are installed, the following apply: 

(a) There must be an individual dispensing unit for each occupant for whom 

supplemental oxygen is to be supplied. Units must be designed to cover the nose 

and mouth and must be equipped with a suitable means to retain the unit in 

158 
3/14/2000 AC 25-22 

position on the face. Flight crew masks for supplemental oxygen must have 

provisions for the use of communication equipment. 

(b) If certification for operation up to and including 25,000 feet is requested, an 

oxygen supply terminal and unit of oxygen dispensing equipment for the 

immediate use of oxygen by each crewmember must be within easy reach of that 

crewmember. For any other occupants, the supply terminals and dispensing 

equipment must be located to allow the use of oxygen as required by the operating 

rules in this chapter. 

(c) If certification for operation above 25,000 feet is requested, there must be 

oxygen dispensing equipment meeting the following requirements: 

(1) There must be an oxygen dispensing unit connected to oxygen supply 

terminals immediately available to each occupant, wherever seated, and at least 

two oxygen dispensing units connected to oxygen terminals in each lavatory. The 

total number of dispensing units and outlets in the cabin must exceed the number 

of seats by at least 10 percent. The extra units must be as uniformly distributed 

throughout the cabin as practicable. If certification for operation above 30,000 

feet is requested, the dispensing units providing the required oxygen flow must be 

automatically presented to the occupants before the cabin pressure altitude 

exceeds 15,000 feet. The crew must be provided with a manual means of making 

the dispensing units immediately available in the event of failure of the automatic 

system. 

(2) Each flight crewmember on flight deck duty must be provided with a quickdonning type oxygen dispensing unit connected to an oxygen supply terminal. 

This dispensing unit must be immediately available to the flight crewmember 

when seated at his station, and installed so that it: 

(i) Can be placed on the face from its ready position, properly secured, sealed, 

and supplying oxygen upon demand, with one hand, within five seconds and 

without disturbing eyeglasses or causing delay in proceeding with emergency 

duties; and 

(ii) Allows, while in place, the performance of normal communication functions. 

(3) The oxygen dispensing equipment for the flight crewmembers must be:

(i) The diluter demand or pressure demand (pressure demand mask with a diluter 

demand pressure breathing regulator) type, or other approved oxygen equipment 

shown to provide the same degree of protection, for airplanes to be operated 

above 25,000 feet. 

(ii) The pressure demand (pressure demand mask with a diluter demand pressure 

breathing regulator) type with mask-mounted regulator, or other approved 

oxygen equipment shown to provide the same degree of protection, for airplanes 

operated at altitudes where decompressions that are not extremely improbable 

may expose the flightcrew to cabin pressure altitudes in excess of 34,000 feet. 

(4) Portable oxygen equipment must be immediately available for each cabin 

attendant. 

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 

42 FR 36971, Jul. 18, 1977; Amdt. 25-87, 61 FR 28696, Jun. 5, 1996]

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3/14/2000 AC 25-22 

b. Intent of Rule. This rule addresses the oxygen dispensing units, including oxygen masks 

and regulators, and do not address oxygen distributing systems, the standards for which are 

addressed in § 25.1445. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). Many of the 

requirements in § 4b.651(d) of the CAR for oxygen dispensing units were carried over 

essentially unchanged to § 25.1447 of 14 CFR. 

 (1) Amendment 25-41 (July 18, 1977) added requirements for a manual means for the 

crew to make the dispensing units available in the event of a failure of the automatic means, and 

that the masks for use by the flight crewmembers be of the quick donning type. 

 (2) Amendment 25-87 (June 5, 1996) added flightcrew oxygen equipment requirements 

for a pressure demand type with mask-mounted regulator, or other approved equipment, for 

airplanes operated at altitudes where decompressions that are not extremely improbable may 

expose the flightcrew to cabin pressure altitudes in excess of 34,000 feet. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer 

to the preamble of this rule and the following information. 

(1) Oxygen Equipment for Forward Observer's Position. The following policy was 

extracted from an FAA memorandum dated February 1, 1988, that was written to provide 

guidance for certification of forward observer's oxygen equipment. 

(a) The requirements for a forward observer's seat are contained in § 25.785(k), 

§ 121.581(a), § 125.317(b), and § 135.75(b). In parts 121, 125, and 135 of 14 CFR, there is no 

specific list of required equipment for the first observer's seat, but rather a general statement that 

the required equipment would be "determined by the Administrator." The FAA considers a 

representative of the Administrator, occupying the first observer's seat and performing official 

duties, to be a required crewmember. This designation also applies to a company check airman, 

or other person performing official duties relating to the performance of the crew or operation of 

the airplane. This person would be expected to interact with the captain and other flight 

crewmembers, in addition to his or her normal duties relating to enroute inspection and

surveillance. For these reasons, it is important that the occupant of the observer seat be provided 

with the equipment necessary to perform his or her function, e.g., oxygen, protective breathing 

equipment, and communication via a radio and interphone panel which is the same type 

equipment provided to the flightcrew. 

(b) Section 25.785(l) does not specify what type of equipment must be provided at 

the observer seat as part of the type design. The rule states, “Each forward observer’s seat 

required by the operating rules must be shown to be suitable for use in conducting the necessary 

enroute inspection.” Section 25.785(l) has been interpreted to allow the use of passenger-type 

oxygen equipment, provided that the airplane in question is not to be used in Part 121, 125, or 

135 operation. If the airplane is to be so used, the oxygen, communication, and protective 

160 
3/14/2000 AC 25-22 

breathing requirements stated above must be provided. If the airplane is not to be used in parts 

121, 125, or 135 operation (i.e., in Part 91 operation), installation of either type of oxygen 

equipment at the first observer's seat is adequate to show compliance with the requirements of

§ 25.1441, Oxygen Equipment and Supply, and installation of communication equipment and 

protective breathing equipment would be optional. 

(c) The duration of oxygen supply should be commensurate with the crew supply, 

since protection may be necessary due to a delayed descent following decompression, protective 

breathing requirements, or other extended usage. 

 e. References. The addresses for ordering the latest revision of technical standard orders 

and other referenced documents listed below can be found in the Appendix to this AC. 

TSO-C64a, Oxygen Mask Assembly Continuous Flow, Passenger (For Air Carrier 

Aircraft). 

TSO-C78, Crewmember Demand Oxygen Masks. 

TSO-C89, Oxygen Regulators, Demand. 

TSO-C99, Protective Breathing Equipment. 

 TSO-C116, Crewmember Protective Breathing Equipment. 

The Society of Automotive Engineers (SAE) has aerospace committees which prepare 

documents reflecting industry standards and practices. The SAE A-10 Committee, Aircraft

Oxygen Equipment, has prepared a handbook, “1992 SAE Aircraft Oxygen Equipment 

Handbook,” containing all Aerospace Standards, Aerospace Recommended Practices, and 

Aerospace Information Reports related to oxygen equipment published by SAE. While not 

regulatory in nature, this handbook contains information that may be of interest to an applicant. 

67. SECTION 25.1449 - MEANS FOR DETERMINING USE OF OXYGEN.

a. Rule Text. 

There must be a means to allow the crew to determine whether oxygen is being 

delivered to the dispensing equipment. 

b. Intent of Rule. The standards in this section ensure that the flightcrew is able to 

determine that oxygen is being dispensed properly. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirement in 

§ 4b.651(e) of the CAR that there be a means for the crew to determine that oxygen is being 

delivered to the dispensing equipment (masks) were carried over essentially unchanged to 

§ 25.1449 of 14 CFR. 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer to 

the preamble of this rule and the following information. Compliance with this section is usually 

161
3/14/2000 AC 25-22 

accomplished by the use of flow indicators in the oxygen regulators (for flight crewmembers) 

and in the oxygen tubes leading to the masks (for passenger equipment). 

(1) Oxygen Equipment for Observer Position. The following policy was extracted 

from an FAA letter dated April 19, 1996, in response to an inquiry concerning the requirement 

for supplemental oxygen flow indication at the observers' stations. 

(a) An applicant proposed installing mask-mounted regulators without flow 

indication at the observers' stations. This configuration has been certified for previous versions 

of the same airplane model. It was the applicant’s position that the FAA exempts mask-mounted 

regulators from the § 25.1449 flow indication requirement. They cited FAA Technical Standard 

Order (TSO) C-89, which exempts mask-mounted regulators from the TSO requirement to 

provide a means to indicate oxygen flow from the regulator outlet. Although the TSO allows 

omission of flow indicators for mask-mounted regulators, it is not an authorization to install such 

items on transport airplanes. However, the FAA reviewed the original certification data for the 

airplane model in question, and the background of § 25.1449. Based on this review, the FAA 

determined that installation of the mask-mounted regulators without flow indicators at the 

observers' stations meets the intent of § 25.1449. 

(b) The background for § 25.1449 is found in the FAA Civil Aeronautics Manual 

(CAM) 4b. Section 25.1449 states: "There must be a means to allow the crew to determine 

whether oxygen is being delivered to the dispensing equipment." Similarly, § 4b.651 of the Civil 

Air Regulation (CAR) states: "Means shall be provided to enable the crew to determine whether 

oxygen is being delivered to the dispensing units." Although flow indicators for each individual 

passenger or flight crew station represent one method of compliance, policy guidance provided 

in § 4b.651-10 of the CAR (for crew systems) and § 4b.651-11 of the CAR (for passenger 

systems) allows operators to establish procedures for checking the oxygen flow to individual 

oxygen users. For passenger diluter-demand systems, § 4b.651-11 of the CAR states that the 

procedure may be "checking of the oxygen flow by a trained crewmember by momentarily 

moving the regulator lever to Automix ‘OFF’ (100% OXYGEN) while the mask is being worn." 

Lack of oxygen flow would be "immediately evidenced by the user's inability to inhale while 

wearing his mask." 

(c) The FAA is satisfied that § 25.1449 does not specifically require flow 

indicators at each mask. Although these indicators are generally provided, other methods of flow 

indication are not precluded. Therefore, the FAA is satisfied that flow indicators are not required 

at the observers' stations for certification, provided that the applicant incorporates suitable design 

features and procedures to ascertain flow at the observers' stations. On the model in question, 

supplemental oxygen for the flightcrew and the observers' stations is provided by a common 

source. Flow of oxygen from the source is split between the flight crew and observers' stations 

by a "T" fitting. Under normal operations, all cockpit masks are set for "100% OXYGEN." At 

this setting, lack of flow would be immediately detected by the user. The FAA is requiring 

incorporation of a preflight check of the observers' masks in the airplane flight manual (AFM). 

In addition, a positive indication of flow at the flightcrew stations by the flow indicators installed 

at those stations would indicate that oxygen is also flowing to the observers' stations. The FAA 

162 
3/14/2000 AC 25-22 

is satisfied that the system design, the preflight inspection procedure, and flow indicators at the 

flightcrew stations are adequate to meet the intent of § 25.1449. 

(2) Determination of Passenger Oxygen Supply. The following policy was extracted 

from an FAA memorandum dated January 8, 1996, which addresses compliance with § 25.1449 

for a design that will determine, from the flightdeck, that oxygen is flowing to the passenger 

supplemental oxygen masks. It is noted that 14 CFR does not require a means for the flightcrew 

to make this determination while on flight deck duty. A number of airplanes have gaseous 

supplemental oxygen systems which do not have a specific indication of oxygen flow on the 

flight deck. Compliance with 14 CFR requirements is usually shown through use of a flow 

indicating device in the oxygen mask tubing. Technical Standard Order TSO-C64a, and the 

associated Society of Automotive Engineers Aerospace Standard SAE-AS 8025, require a means 

for the crew to determine that oxygen is flowing. This is usually accomplished with a small 

device that provides a green indication when the flow rate exceeds 0.5 liters per minute. A cabin 

crewmember verifies from the indicator whether or not oxygen is flowing. If there are no cabin 

crewmembers, this would require a flight crewmember to leave their station and check. The 

FAA has determined that this would not be a viable option during an emergency such as a 

decompression. 

 e. References. None. 

68. SECTION 25.1450 - CHEMICAL OXYGEN GENERATORS.

a. Rule Text.

(a) For the purpose of this section, a chemical oxygen generator is defined as a 

device which produces oxygen by chemical reaction. 

(b) Each chemical oxygen generator must be designed and installed in 

accordance with the following requirements: 

(1) Surface temperature developed by the generator during operation may not 

create a hazard to the airplane or to its occupants. 

(2) Means must be provided to relieve any internal pressure that may be 

hazardous. 

(c) In addition to meeting the requirements in paragraph (b) of this section, each 

portable chemical oxygen generator that is capable of sustained operation by 

successive replacement of a generator element must be placarded to show- 

(1) The rate of oxygen flow, in liters per minute; 

(2) The duration of oxygen flow, in minutes, for the replaceable generator 

element; and 

(3) A warning that the replaceable generator element may be hot, unless the 

element construction is such that the surface temperature cannot exceed 100 

degrees F. 

[Amdt. 25-41, 42 FR 36971, Jul. 18, 1977]

163
3/14/2000 AC 25-22 

b. Intent of Rule. The standards in this rule ensure that chemical oxygen generators safely 

provide oxygen. 

 c. Background. The standards in this section address chemical oxygen generators. This 

subject was not addressed in either part 4b of the CAR or part 25 when it was originally codified. 

This section was added at Amendment 25-41, July 18, 1977. 

 d. Policy/Compliance Methods. There is no written policy or guidance in our files. 

 e. Reference. The address for ordering the latest revision of the document listed below can 

be found in the Appendix to this AC. 

The Society of Automotive Engineers (SAE) has aerospace committees that prepare 

documents reflecting industry standards and practices. The SAE A-10 Committee, Aircraft

Oxygen Equipment, has prepared a handbook, “1992 SAE Aircraft Oxygen Equipment 

Handbook,” containing all Aerospace Standards, Aerospace Recommended Practices, and 

Aerospace Information Reports related to oxygen equipment published by SAE. While not 

regulatory in nature, this handbook contains information that may be of interest to an applicant. 

69. SECTION 25.1453 PROTECTION OF OXYGEN EQUIPMENT FROM RUPTURE.

a. Rule Text.

Oxygen pressure tanks, and lines between tanks and the shutoff means, must be- 

(a) Protected from unsafe temperatures; and 

(b) Located where the probability and hazards of rupture in a crash landing are 

minimized. 

b. Intent of Rule. This rule ensures that the oxygen equipment is not exposed to unsafe 

temperatures and minimizes the possibility of rupture in the event of a crash landing. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in § 4b.651(g) of the CAR were carried over essentially unchanged to § 25.1453 of 14 CFR. 

d. Policy/Compliance Methods. For guidance on compliance with this requirement, refer to 

the preamble of this rule and the following information. 

(1) Flexible Plastic Oxygen Lines. The following policy was extracted from an FAA 

memorandum dated December 27, 1983, to provide guidance regarding the use of flexible plastic 

tubing for oxygen distribution lines. While this guidance was generated for Part 23 airplanes, it 

is equally applicable for part 25 airplanes. 

164 
3/14/2000 AC 25-22 

(a) The FAA has determined that the use of plastic lines for an oxygen distribution 

system that is operating under continuous pressure is not acceptable for certification in Part 23 

airplanes. 

(b) Lines constructed of combustible materials, including nylon, polyvinylchloride 

(PVC) and Teflon, may be used in oxygen lines which are pressurized only when cabin 

depressurization occurs. The following precautions should be taken when using such lines in the 

oxygen system: 

 1 Swaged metal type end fittings should be used to prevent leakage from

cold flow. 

 2 Lines should be protected from abrasion by use of a reinforcing sleeving 

of fabric braid. 

 3 Precautions should be taken to route such lines away from areas where 

they might be subjected to elevated temperatures, electrical arcing (relays and switches) and 

flammable fluids. 

 4 Refer to Advisory Circular (AC) 43.13-2A, Chapter 6, for additional 

guidance material. 

(2) Use of Plastic Oxygen Lines. The following policy was extracted from an FAA 

memorandum dated October 28, 1983, and addresses the use of plastic or nylon tubing as oxygen 

lines. 

(a) Tubing in the fuselage for high or low pressure oxygen systems that are located 

behind liners or in the walls of the fuselage are typically made of rigid stainless steel (for high 

pressure) or aluminum (for low pressure). Synthetic flexible lines connecting the oxygen mask 

to the oxygen distribution system have been accepted. Swaged metal end fitted PVC tubing, 

covered with a synthetic braid (for abrasion resistance and strength), should be used between the 

aluminum low pressure distribution line and the passenger service unit manifold of many large 

transports. These tubes and aluminum line do not contain oxygen until a depressurization 

occurs. Then, the pressure is low and for a short duration. These hoses meet the interior burn 

requirements. 

(b) Synthetic lines such as plastic or nylon cannot be recommended for oxygen 

high or low pressure lines that will be exposed to a continuous pressure (i.e., as opposed to 

pressurized when needed). These materials can cold flow. Care must be taken in the selection of 

the fitting design for exposures of even short duration. 

(c) In addition to cold flow, polyethylene and nylon will lose strength with 

increasing temperature. These materials are much more susceptible to combustion in the 

presence of oxygen than either stainless steel or aluminum. For these reasons, the FAA 

165
3/14/2000 AC 25-22 

considers polyethylene or nylon tubing inappropriate and unsafe for oxygen lines that are 

subjected to continuous oxygen pressure. 

 e. Reference. The address for ordering the document listed below can be found in the 

Appendix to this AC. 

The Society of Automotive Engineers (SAE) has aerospace committees that prepare 

documents reflecting industry standards and practices. The SAE A-10 Committee, Aircraft

Oxygen Equipment, has prepared a handbook, “1992 SAE Aircraft Oxygen Equipment 

Handbook,” containing all Aerospace Standards, Aerospace Recommended Practices, and 

Aerospace Information Reports related to oxygen equipment published by SAE. While not 

regulatory in nature, this handbook contains information that may be of interest to an applicant. 

70. SECTION 25.1455 - DRAINING OF FLUIDS SUBJECT TO FREEZING.

 a. Rule Text.

If fluids subject to freezing may be drained overboard in flight or during ground 

operation, the drains must be designed and located to prevent the formation of 

hazardous quantities of ice on the airplane as a result of the drainage. 

[Amdt. 25-23, 35 FR 5680, Apr. 8, 1970]

b. Intent of Rule. This rule addresses fluids subject to freezing that normally drain from

the airplane either in flight or during ground operation. The intent is to prevent damage to the 

airplane when the ice disengages and falls free. This section addresses "gray" water from the 

sinks and drains in galleys and lavatories. Section 25.1455 does not apply to lavatory drain 

systems that are normally drained when the airplane is parked, rather than being in "ground 

operation." This requirement is usually met by designing the drain masts (usually heated to 

prevent the fluid from freezing) so that the fluid does not impact the airframe or engines when it 

is draining. 

 c. Background. Effective February 1, 1965, part 25 was added to Title 14, Code of Federal 

Regulations (14 CFR) to replace part 4b of the Civil Air Regulations (CAR). The requirements 

in section 4b.660 of the CAR were carried over to § 25.1455 of 14 CFR. 

 (1) Amendment 25-23 (April 8, 1970) changed the intent of the rule from preventing 

the formation of ice to preventing the formation of hazardous quantities of ice on the airplane as 

a result of drainage.

 d. Policy/Compliance Methods. There is no written policy or guidance in our files.

 e. Reference. The address for ordering the latest revision of the advisory circular listed 

below can be found in the Appendix to this AC. 

166 
3/14/2000 AC 25-22 

AC 25.1455-1, Waste Water/Potable Water Drain System Certification Testing.

71. SECTION 25.1461 - EQUIPMENT CONTAINING HIGH ENERGY ROTORS.

 a. Rule Text.

(a) Equipment containing high energy rotors must meet paragraph (b), (c), or (d)

of this section. 

(b) High energy rotors contained in equipment must be able to withstand damage 

caused by malfunctions, vibration, abnormal speeds, and abnormal temperatures. 

In addition- 

(1) Auxiliary rotor cases must be able to contain damage caused by the failure of

high energy rotor blades; and 

(2) Equipment control devices, systems, and instrumentation must reasonably 

ensure that no operating limitations affecting the integrity of high energy rotors 

will be exceeded in service. 

(c) It must be shown by test that equipment containing high energy rotors can 

contain any failure of a high energy rotor that occurs at the highest speed 

obtainable with the normal speed control devices inoperative. 

(d) Equipment containing high energy rotors must be located where rotor failure

will neither endanger the occupants nor adversely affect continued safe flight. 

[Amdt. 25-41, 42 FR 36971, Jul. 18, 1977]

b. Intent of Rule. This rule ensures that failures of high energy rotating equipment will not 

adversely affect systems, structure, or occupants in the event that the rotating components fail at 

high speeds. Some of the equipment that contains high energy rotors are turbine engine starters, 

air cycle machines, air driven hydraulic pumps, cooling fans for galley equipment and electronic 

bays, cabin ventilation recirculation fans, and high speed electrically driven hydraulic pumps. 

This section does not pertain to equipment covered by Subpart E, Propulsion, such as engines 

and auxiliary power units, which are covered by the provisions of § 25.903(d). 

167
3/14/2000 AC 25-22 

 c. Background. 

(1) Amendment 4b-8 of the Civil Air Regulations (CAR). This subject was first 

addressed in part 4b of the CAR under Amendment 4b-8, effective May 17, 1958. A new 

§ 4b.659 was included which required appropriate protection of the airplane against failure of 

high energy rotors when such rotors were incorporated in any equipment on the airplanes. 

Section 4b.659 read "Equipment incorporating high energy rotors shall be demonstrated as 

capable of containing a failed rotor or shall be so located that failure will not affect the ability of

the airplane to continue safe flight." 

(2) Amendment 4b-12 of the Civil Aeronautics Manual (CAM). Amendment 4b-12, 

effective May 13, 1962, deleted § 4b.659 because it was surmised that its substance was covered

by the provisions of § 4b.606 (which later became § 25.1309) which is concerned with the 

reliability of all equipment, systems, and installations. 

 (3) Amendment 25-41. Amendment 25-41 added § 25.1461 as a result of an 

Airworthiness Review Program conducted in 1975. This new section added requirements for 

protection against the failure of equipment containing high energy rotors. Experience had 

demonstrated that failures which release the energy stored in these rotors may result in engine or 

structural damage, fires, or injury to occupants. 

 d. Policy/Compliance Methods. Compliance to §25.1461(c) is accomplished if it can be 

shown by test that equipment containing high-energy rotors can contain any failure of the rotor 

that occurs at the highest speed obtainable with normal speed control devices inoperative. In 

general, three phase induction motor driven fans do not have speed control devices. Fan speed is 

a function of the input power frequency supplied to the device. 

(1) Because of the absence of speed controls, these fans can be treated as an extension 

of the aircraft power generating system. To comply with § 25.1461(c), the minimum test speed 

would be at the highest speed obtainable with power generating system speed control devices 

inoperative (speed control device is the mechanical system that governs generator shaft speed). 

In the event of such a failure, generating system controls will limit frequency. Exceeding those 

limits would require further multiple failures. These higher order failures do not have to be 

considered to comply with § 25.1461(c). 

(2) For external ground power operation, the same frequency limits are applicable. The 

cart (or other) providing power must fail, the cart protective overfrequency protection must fail, 

and the airplane protective systems must fail to trip the external power contactor. In addition, an 

uncontained fan rotor burst on the ground is classified differently as it is not a safety of flight 

issue. 

 e. References. None. 

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3/14/2000 AC 25-22 

Chapter 4. OPERATING LIMITATIONS AND INFORMATION

Section 1. OPERATING LIMITATIONS 

72. SECTION 25.1529 - INSTRUCTIONS FOR CONTINUED AIRWORTHINESS.

a. Rule Text. 

The applicant must prepare instructions for Continued Airworthiness in 

accordance with Appendix H to this Part that are acceptable to the Administrator. 

The instructions may be incomplete at type certification if a program exists to 

[Amdt. 25-54, 45 FR 60173, Sep. 11, 1980] ensure their completion prior to 

delivery of the first airplane for issuance of a standard certificate of 

airworthiness, whichever occurs later.

H25.4 Airworthiness Limitations section. 

The Instructions for Continued Airworthiness must contain a section titled 

Airworthiness Limitations that is segregated and clearly distinguishable from the 

rest of the document. This section must set forth each mandatory replacement 

time, structural inspection interval, and related structural inspection procedures 

approved under § 25.571. If the Instructions for Continued Airworthiness consist 

of multiple documents, the section required by this paragraph must be included in 

the principal manual. This section must contain a legible statement in a 

prominent location that reads: "The Airworthiness Limitations section is FAA 

approved and specifies maintenance required under §§ 43.16 and 91.403 of the 

Federal Aviation Regulations unless an alternative program has been FAA 

approved." 

[Amdt. 25-54, 45 FR 60173, Sep. 11, 1980; Amdt. 25-68, 54 FR 34329, Aug. 18, 

1989]

b. Intent of Rule. The purpose of this rule is to ensure continued airworthiness of 

the airplane by requiring that inspections, checks, and replacement of parts are performed 

in accordance with the requirements established for type certification. 

 c. Background. This subject was not addressed in either section 4b of the Civil Air 

Regulations (CAR) or part 25 of Title 14, Code of Federal Regulations (14 CFR). This 

rule was added, along with Appendix H, by Amendment 25-54. 

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3/14/2000 AC 25-22 

 d. Policy/Compliance Methods. For guidance on compliance with this requirement, 

refer to the preamble of this rule and the airworthiness limitations document. 

e. Reference. The address for ordering the latest revision of the advisory circular 

listed below can be found in the Appendix to this AC. 

Advisory Circular 25-19, Certification Maintenance Requirements. 

170 
3/14/2000 AC 25-22 

APPENDIX 1, CROSS REFERENCE FOR SELECTED SECTIONS OF 

PART 25, CAR 4b/CAM 4b

DISTRIBUTION TABLE 

Revised Section Former Section Revised

Section

Former Section

25.671 4b.320 (less (b)) 25.869(b) 25.1443(b) & (c) [4b.658] 

25.672 not applicable 25.869(c) 25.1451 [4b.651(f)] 

25.699 4b.323(e) & (f) 25.1001 4b.437 (less (e) (last sent.)) 

25.701 4b.324 25.1183 4b.333, 4b.483

25.703 not applicable 25.1185 4b.481

25.729 4b.334, -2 25.1189 4b.382

25.731 4b.335(a) & (b) 25.1301 4b.600, 4b.601 

25.733 4b.336 25.1309 4b.606

25.735 4b.335 less (a), (b) 25.1416 not applicable 

4b.337, -4 (1st sent.) 25.1419 4b.640 (less intro, 3rd sent. 

25.771 4b.350 25.1431 4b.650

25.773 4b.351 25.1433 4b.658

25.775 4b.352 25.1435 4b.653, 4b.654, 4b.655 

25.777 4b.353 25.1438 not applicable

25.783 4b.356, -1(last sent) 25.1439 4b.651 (less (a)-(g) 

4b.356-2(a)/1st sent 25.1441 4b.651(a)

25.831 4b.371, 4b.371-1 25.1443 4b.651(b)

25.832 not applicable 25.1445 4b.651(c), 4b.651-5a 

25.833 4b.372 25.1147 4b.651(d)

25.841 4b.374, 4b.375, -1 25.1449 4b.651(e) 

25.843 4b.376 25.1450 not applicable

25.851 4b.380 (less (c)), 4b.383 

(2nd sent. of (a) & 

(b)(3) 

25.1453 4b.651(g)

25.854 not applicable 25.1455 4b.660

25.855 4b.382, 4b.384 25.1461 not applicable 

25.857 not applicable

25.858 not applicable

25.863 4b.385

171
3/14/2000 AC 25-22 

APPENDIX 2, GLOSSARY OF ACRONYMS

AC Advisory Circular 

ACO Aircraft Certification Office 

AD Airworthiness Directive

AFM Airplane Flight Manual 

AIA Aerospace Industries Association 

AIR Aircraft Certification Service 

ARAC Aviation Rulemaking Advisory Committee 

ARP Aerospace Recommended Practice (SAE) 

BTMS Aircraft Brake Temperature Monitor System

CAA Civil Aeronautics Administration (US) 

CAR Civil Air Regulations 

CAM Civil Aeronautics Manual 

CFR Code of Federal Regulations 

DER Designated Engineering Representative 

ECS Environmental Control System

EICAS Engine Indication and Crew Alerting System

FAR Federal Aviation Regulations 

IIDSS Inflight Ice Detection Sensing Systems 

JAA Joint Aviation Authorities 

JAR Joint Airworthiness Requirements 

KE Kinetic Energy 

NOE Non-original Equipment 

NTSB National Transportation Safety Board 

OAT Outside Air Temperature 

OBOGS On-Board Oxygen Generating System 

OEM Original Equipment Manufacturer 

PBE Protective Breathing Equipment 

PMA Parts Manufacturer Approval 

PPM Parts Per Million 

PPMV Parts Per Million by Volume

QTR Qualification Test Report 

R&D Research & Development 

RTO Rejected Take-off 

SAE Society of Automotive Engineers 

SLD Supercooled Large Droplets 

SLE Sea Level Equivalent 

STC Supplemental Type Certificate 

TAD Transport Airplane Directorate 

TAT Total Air Temperature 

TC Type Certificate 

TSO Technical Standard Order 

TWA Time-Weighted-Average 

171 
3/14/2000 AC 25-22 

APPENDIX 3, INDEX OF ADVISORY CIRCULAR REFERENCES

The advisory circulars listed below contain information relevant to the approval of mechanical 

systems on transport category airplanes. They can be obtained from the U.S. Department of 

Transportation, Subsequent Distribution Office, SVC-121.23, Ardmore East Business Center, 

3341 Q 75th Avenue, Landover, MD 20785, USA. 

AC No. Title Section 

 Reference

20-32B Carbon Monoxide (CO) Contamination in Aircraft- 25.831 

Detection and Prevention, November 24, 1972, or latest 

 revision, ACE-110 

20-33B Technical Information Regarding Civil Aeronautics All sections

Manuals 1, 3, 4a, 4b, 5, 6, 7, 8, 9, 13 and 14, May 1, 1975, 

 or latest revision, AIR-200 

20-34D Prevention of Retractable Landing Gear Failures, 25.729 

August 8, 1980, or latest revision, AFS-340 

20-36S Index of Articles (Materials, Parts, Processes and All sections 

 Appliances) Certified Under the Technical Standard Order 

System, September 1, 1993, or latest revision, AFS-613 

20-41A Substitute Technical Standard Order (TSO) Aircraft All sections

Equipment, April 5, 1977, or latest revision, AIR-120 

20-42C Hand Fire Extinguishers for use in Aircraft, 25.851 

March 7, 1984, or latest revision, ACE-110 

20-62D Eligibility, Quality, and Identification of Approved All sections 

Aeronautical Replacement Parts, May 24, 1996, 

or latest revision, AFS-340 

20-73 Aircraft Ice Protection, April 21, 1971, or latest revision, 25.1419 

 AIR-120 

20-97A High-Speed Tire Maintenance and Operational Practices, 25.733 

May 13, 1987, or latest revision, AFS-340 

171
3/14/2000 AC 25-22 

20-99 Antiskid and Associated Systems, May 27, 1977, or 25.735 

 latest revision, ANM-100 

20-110J Index of Aviation Technical Standard Orders, All sections 

May 30, 1997, or latest revision, AIR-120 

20-117 Hazards Following Ground Deicing and Ground 25.1419 

 Operations in Conditions Conducive to Aircraft Icing, 

December 17, 1982; Chg. 1, April 15, 1983, or 

 latest revision, AFS-200 

20-128A Design Considerations for Minimizing Hazards Caused 25.1435 

by Uncontained Turbine Engine and Auxiliary Power 

Unit Rotor and Fan Blade Failures, March 25, 1997, or 

 latest revision, ANM-114 

21-16C Radio Technical Commission for Aeronautics (now called All sections

RTCA Inc.) Document DO-160C, February 14, 1990, 

 or latest revision, AIR-120 

21-23 Airworthiness Certification of Civil Aircraft Engine, 25.735 

Propellers, and Related Products, July 7, 1987, or 

 latest revision, AIR-4 

23.729-1 Landing Gear Doors and Retraction Mechanism, March 26, 23.729 

1984, or latest revision, ACE-110. For information only. 

23-1419-2 Certification of Part 23 Airplanes for Flight in Icing 23.1419 

 Conditions, January 3, 1992, or latest revision, ACE-100 

25-7A Flight Test Guide for Certification of 25.729, 25.731 

Transport Category Airplanes, March 31, 1998, or 25.733, 25.735 

 latest revision, ANM-110 25.773 

25-9A Smoke Detection, Penetrating, and Evacuation Tests and 25.831, 25.854 

Related Flight Manual Emergency Procedures, 25.855, 25.858 

January 6, 1994, or latest revision, ANM-110 

25-14 High Lift and Drag Devices, May 4, 1988, or latest 25.699, 25.703 

 revision, ANM-112 25.671, 25.672 

25-16 Electrical Fault and Fire Prevention and Protection, 25.869 

April 5, 1991, or latest revision, ANM-111 

25-17 Transport Airplane Cabin Interiors Crashworthiness 25.773, 25.857 

Handbook, July 15, 1991, or latest revision, ANM-114 25.858, 25.869 

172 
3/14/2000 AC 25-22 

25-18 Transport Category Airplanes Modified for Cargo Service, 25.857 

January 6, 1994, or latest revision, ANM-114 

25-19 Certification Maintenance Requirements, November 28, 25.1309 

1994, or latest revision, ANM-113 

25-20 Pressurization, Ventilation, and Oxygen Systems 25.831, 25.841 

Assessment for Subsonic Flight Including High Altitude 

Operation, September 10, 1996, or latest revision, ANM-111 

25-21 Certification of Transport Airplane Structure, 25.671, 25.672 

September 1, 1999, ANM-110 25.701 

25.672-1 Active Flight Controls, November 15, 1983, or latest 25.672 

 revision, ANM-110B 

25.703-1 Takeoff Configuration Warning Systems, March 17, 1993, 25.703 

 or latest revision, ANM-111 

25.773-1 Pilot Compartment View for Transport Category Airplanes, 25.773 

January 8, 1993, or latest revision, ANM-111 

25.783-1 Fuselage Doors, Hatches, and Exits, December 10, 1986, 25.783 

 or latest revision, ANM-110 

25.963--1 Fuel Tank Access Covers, July 29, 1992, or latest 25.729 

 revision, ANM-112 

25.1309-1A System Design Analysis, June 21, 1988, or latest 25.1309, 25.1435 

 revision, ANM-112 

25.1419-1 Certification of Transport Category Airplanes for 25.1419 

Flight in Icing Conditions, August 18, 1999, or latest 

revision, ANM-112 

25.1455-1 Waste Water/Potable Water Drain System Certification 25.1455 

Testing, March 11, 1985, or latest revision, ANM-112 

27-1A Certification of Normal Category Rotorcraft, dated 25.731 

July 30, 1997, or latest revision, ASW-111 

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3/14/2000 AC 25-22 

29-2B Certification of Transport Category Rotorcraft, July 30, 1997, 25.731 

 or latest revision, ASW-111 

43.13-1A Acceptable Methods, Techniques and Practices 25.729, 25.731, 

Aircraft Inspection and Repair (with Errata Sheet), dated 25.733, 25.735 

April 17, 1972, or latest revision, AFS-600 

43.13-2A Acceptable Methods, Techniques, and Practices Aircraft All sections 

 Alterations, includes Chg. 1, dated June 9, 1977; Chg. 2, 

dated Oct. 30, 1989, or latest revision, AFS 340 

91-6A Water, Slush, and Snow on the Runway, May 24, 1978, 25.109, 25.735 

 or latest revision, AFS-430 

91-13C Cold Weather Operation of Aircraft, July 24, 1979, 25.109, 25.735 

 or latest revision, AFS-806 

91-51A Effect of Icing on Aircraft Control and Airplane De-ice and 25.1419 

Anti-Ice Systems, July 19, 1996, or latest revision, AFS-820 

120-38 Transport Category Airplanes Cabin Ozone Concentrations, 

October 10, 1980, or latest revision, AFS-260 25.832 

120-39 Hazards of Waste Water Ice Accumulation Separating From 25.1455 

 Aircraft in Flight, October 31, 1980, or latest revision, AFS-331 

120-42A Extended Range Operation with Two Engine Airplanes 25.1435 

(ETOPS), December 30, 1988, or latest revision, AFS-400 

120-58 Pilot Guide for Large Aircraft Ground Deicing, September 30, 25.1419 

1992, or latest revision, AFS-420 

121.195-1A Operational Landing Distances for Wet Runways: Transport 25.109, 25.735 

 Category Airplanes, June 19, 1990, or latest revision, AFS-430 

145-4 Inspection, Retread, Repair and Alterations of Aircraft Tires, 

September 27, 1982, or latest revision, AFS-340 25.733 

174 
3/14/2000 AC 25-22 

*25-XX Cargo Compartment Fire Extinguishing or Suppression 

 Systems, date TBD 25.857 

*25-XX Class B and F Cargo Compartments, date TBD 25.857 

*25-XX Propulsion Systems Handbook 25.863, 25.943, 

(Mega AC) date TBD 25.1001, 25.1183, 

 25.1185, 25.1189 

*25-XX Electrical Systems Handbook (Mega AC) 25.1301, 25.1309, 

 date TBD 25.1325 

*NOTE: The designation "25-XX" means that a draft AC exists for the subject. Upon approval, 

the "XX" will be replaced with a numerical reference identification code.

175
3/14/2000 AC 25-22 

APPENDIX 4, INDEX OF INCORPORATED GUIDANCE MATERIAL AND 

REFERENCES

1. FAA Memorandums and Letters Referenced in this AC:

Copies of the memorandums and letters are available from the FAA Transport Airplane 

Directorate, 1601 Lind Ave. SW, Renton, WA 98055, USA. 

SectionSubject Date 

25.729 Protection of Equipment in Wheel Wells, Dec. 4, 1997 

 § 25.729(f)(1) 

25.729 Landing Gear Position Indication System - July 12, 1988 

“Backup Requirement” - Section 25.729(e) 

25.729 Flap System/Landing Gear Warning System Dec. 19, 1983 

 Tie-In 

25.729 Landing Gear Position Indication System June 3, 1983 

25.729 Landing Gear Slush Tests April 22, 1983 

25.733 Approval of Retreaded Tires by Similarity March 1, 1995 

25.733 Approval of Nonretreadable Tires - August 18, 1988 

 Qualification Testing 

25.733 Approval Method for Substitute Tires April 14, 1988 

25.733 Replacement of Bias Ply Tires with Radial Feb. 7, 1984 

 Ply Tires

25.733 Certification Program for Replacing Bias Jan. 12, 1984 

Tires with Radial Tires, Part 25 Transports; 

 All Weight Categories 

25.733 Certification of Radial Tire Installation March 28, 1984 

25.733 Critical Conditions & Maximum Ramp Jan. 26, 1981 

 Weight Definitions 

25.735 Carbon Brake Refurbishment April 11, 1996 

 (redensification) 

25.735 Parts Manufacturer Approval for a April 9, 1996 

Non-U.S. Location: § 21.303(g) 

25.735 Certification of Replacement Brakes: May 19, 1992 

 Non-U.S. Location 

25.735 Parking Brake Testing March 20, 1992, 

25.735 Worn Brake Requirements for Non- Feb. 18, 1992 

 Original-Equipment Parts (letter)

25.735 Determination of Allowable Feb. 23, 1990, Mar. 2, 1990 

 Worn Brake Limits 

25.735 Credit for Reverse Thrust in Worn April 28, 1989 

 Brake Testing 

171 
3/14/2000 AC 25-22 

25.735 Certification of Brake Replacement Mar. 16, 1989, Dec. 27, 1988, 

 Components - Performance and March 16, 1988 

25.735 Ground Equipment-Brake Cooling Unit Dec. 27, 1988 

25.735 Certification Requirements for Asbestos- Dec. 16, 1987 

 Free Brake Linings 

25.735 Determination of Certification Requirements Oct. 26, 1987 

for Brake Components - Performance 

25.735 Approval of Brake Components Utilizing July 27, 1987 

New Brake Friction Material 

25.735 Maximum Quick Turnaround Times August 19, 1983 

25.735 Dynamometer Test Guidelines to Establish March 2, 1990 

Brake Wear Limits (letter) 

25.831 Ventilation During Takeoff with the May 20, 1983 

Environmental Control System Turned Off 

25.831 Fresh Air and Normal Operating Conditions Sept. 10, 1997 

25.831 Cabin Altitude Limit; § 25.841(a) Jan. 12, 1994 

25.832 Use of Operational Limitations March 31, 1997 

25.832 Meaning of the Term Time-Weighted Average Oct. 26, 1987 

in § 25.832 of the FAR 

25.833 Standards for Certification of Combustion March 14, 1948 

 Heaters 

25.841 Rapid Equalization of Pressure June 21, 1982 

25.841 Access to a Pressurized Cargo Compartment June 25, 1986 

in the Engine Burst Zone During High 

 Altitude Operation 

25.841 Cabin Altitude Limit; § 25.841(a) Jan. 12, 1994 

25.843 Pressure Tests for Compliance with 25.843(a) Feb. 21, 1990 

25.854 Approval of Lavatory Smoke Detectors Oct. 28, 1991 

25.854 Approval of Lavatory Fire Extinguishers March 31, 1997 

Containing Agents Other Than Halon 

25.857 Main Deck Cargo Compartment Fire June 6, 1997 

 Protection Certification Procedures 

25.857 Cargo Compartments Sept. 13, 1988 

25.857 Class A Cargo Compartments February 8, 1996 

25.857 Class E Cargo Compartments March 21, 1991 

25.857 Class E Cargo Compartments July 3, 1990 

25.858 Smoke Detection Certification Testing June 18, 1997 

25.858 Smoke Detection Certification Testing Feb. 11, 1993 

25.858 Smoke Detection Certification Testing Dec. 3, 1992, Dec. 15, 1992 

25.1419 Inflight Ice Detection Systems Jan. 13, 1998 

25.1419 Aircraft Mounted Ground Ice Detection Oct. 10, 1996 

 Systems 

25.1419 Roll Control in Supercooled Large Droplets July 23, 1997 

25.1435 Hydraulic Fuel Contamination April 20, 1995 

172
3/14/2000 AC 25-22 

25.1435 Certification of Hydraulic Lines for April 23, 1992 

 Temporary Repairs 

25.1435 Use of Hydraulic Lines for Temporary Repairs Jan. 25, 1988 

25.1435 Hydraulic System Certification Philosophy March 5, 1982 

25.1435 Interpretation of 25.1425(a)(4) of the FAR Sept. 14, 1978 

25.1438 Pneumatic Versus Pressurization System Oct. 19, 1993 

 Identification 

25.1439 Protective Breathing Equipment Usage May 13, 1993 

When Approved Under TSO-C99 

 and TSO-C116 

25.1439 Protective Breathing Equipment March 9, 1990 

 Serviceability 

25.1439 Protective Breathing Equipment on the Dec. 19, 1988 

 Flight Deck 

25.1439 Definition of Terms Used in § 25.1439 Feb. 26, 1985 

25.1441 Certification of On-Board Oxygen Jan. 31, 1996 

 Generation Systems 

25.1441 Compliance with §§ 25.1309(c) and 25.1441(c) Feb. 23, 1994 

25.1441 Guidance Availability of Flightcrew Oxygen May 11, 1992 

25.1441 Oxygen Installation for Medical Use Jan. 3, 1991 

25.1445 Isolation of Flight Crewmember Oxygen Jan. 8, 1996 

 Supply 

25.1447 Oxygen Equipment for Forward Observer’s Feb. 1, 1988 

 Position 

25.1447 Oxygen Equipment for Observer Position Jan. 24, 1983 

25.1449 Oxygen Equipment for Observer Position April 19, 1996 

25.1450 Life Limit on Solid State Oxygen Aug. 15, 1983 

25.1453 Use of Plastic Oxygen Lines Dec. 27, 1983 

25.1453 Use of Plastic Oxygen Lines Oct. 28, 1983 

173 
3/14/2000 AC 25-22 

APPENDIX 4

2. FAA Technical Standard Orders (TSO) Referenced in this AC.

The Technical Standard Orders listed below contain information relevant to mechanical system

components installed on transport category airplanes. They can be obtained from the U.S. 

Department of Transportation, Subsequent Distribution Office, SVC-121.23, Ardmore East 

Business Center, 3341 Q 75th Avenue, Landover, MD 20785, USA. 

TSO Date Title Section

 Reference

TSO-C1c July 10, 1987 Cargo and Baggage Compartment Smoke 25.854 

 Detection Instruments

TSO-C19b Portable Water-Solution Type Fire 25.851 

 Extinguishers 

TSO-C20 Combustion Heaters 25.833 

TSO-C26 Mar. 15, 1952 25.731 

TSO-C26a June 1, 1961 25.731 

TSO C26b 1962 Aircraft Tires 25.733 

TSO-C26b Jan. 21, 1971 25.731 

TSO-C26c Dec. 31, 1979 Aircraft Tires 25.731 

TSO-C26c May 18, 1984 Aircraft Wheels and Wheel-Brake 25.731, 25.735 

 Assemblies, with Addendum I 

TSO-C45a Feb. 28, 1995 Manifold Pressure Indicating Instruments 25.869 

TSO-C47 Pressure Instruments - Fuel, Oil, and 25.1435 

 Hydraulic

TSO-C48 Carbon Monoxide Detector Instruments 25.831 

TSO C62c Dec. 31, 1979 Aircraft Tires 25.733 

TSO-C62c Sept. 12, 1984 Aircraft Tires, with Addendum 1 25.733 

TSO-C62d Sept. 7, 1990 Aircraft Tires 25.733 

TSO-C64a Aug. 25, 1989 Oxygen Mask Assembly Continuous 25.1445, 25.1447 

 Flow, Passenger (for air carrier aircraft) 

TSO-C75 Hydraulic Hose Assemblies 25.1435 

TSO-C78 Crewmember Demand Oxygen Masks 25.1447 

TSO-C89 Oxygen Regulators, Demand 25.1447, 25.1449 

TSO-C99 June 27, 1983 Protective Breathing Equipment 25.1439 

TSO-116 March 1, 1990 Crewmember Protective Breathing 25.1439, 25.1447 

 Equipment

174
3/14/2000 AC 25-22 

APPENDIX 4

3. FAA Orders Referenced in this AC.

The FAA Orders listed below contain information relevant to mechanical system components 

installed on transport category airplanes. They can be obtained from the U.S. Department of 

Transportation, Subsequent Distribution Office, SVC-121.23, Ardmore East Business Center, 

3341 Q 75th Avenue, Landover, MD 20785, USA. 

Number Title Section Reference 

8000.40D Maintenance of Pressure Cylinders In Use As 25.857, 25.1435 

Aircraft Equipment, Dec. 26, 1995, AFS-350 

8000.54 Process Specifications for Retreading Tires and 25.733 

Inspection Procedures Manuals for Tire Retreaders, 

Nov. 2, 1982, AFS-350 

8000.64 Qualification of Aircraft Radial Tires for Use on 25.733 

Aircraft and for Retreading, Dec. 4, 1986, AFS-350 

8110.4A Type Certification Process, March 2, 1995, AIR-110 21.11 to 21.53 

8110.8 Engineering Flight Test Guide for Transport Category 25.729, 25.735 

Airplanes, Chg. 5, July 17, 1986. This order was 

canceled by AC 25-7A, March 31, 1998, ANM-100 

8150.1A Technical Standard Order Procedures, Sept. 21, 1987, 21.60 to 21.621 

 AIR-120 

4. FAA Videos Referenced in this AC:

These videos are available from the William J. Hughes (FAA) Technical Center, 

Atlantic City International Airport, Atlantic City, NJ 08405, USA. 

Title Section

 Reference

FAA’s Tire Approval Process Video, MTS 422/422.1, 42:15 minutes, 

September, 1994 25.733 

FAA’s Wheels & Brakes Approval Process Videos, Parts 1 & 2, 

September, 1994 25.735 

FAA’s Smoke Quantities to Certify Smoke Detection Systems in 25.858 

Cargo Areas, June, 1997 

175 
3/14/2000 AC 25-22 

APPENDIX 4

5. SAE Documents Referenced in this AC.

The Society of Automotive Engineers (SAE) has aerospace committees that prepare documents 

reflecting industry standards and practices. While not regulatory in nature, this handbook 

contains information that may be of interest to an applicant. These documents provide additional 

information, guidance, and/or standards, and are available from the Society of Automotive 

Engineers, Inc., 400 Commonwealth Drive, Warrendale, PA 15096-0001, USA. 

Number Title Section

 Reference

SAE AIR-737E Hydraulic and Pneumatic Specifications and Standards, 25.1435 

 May 1995 

SAE AIR-786A Elastomer Compatibility Considerations Relative to 25.1435 

O-Ring and Sealant Selection, July 1992 

SAE AIR-811B Disposition of Wheels which have been Overheated, 25.731 

 April 1, 1996 

SAE AIR-1047C A Guide for Selection of Quick Disconnect Couplings for 25.1435 

Aerospace Fluid Systems, April 1994 

SAE AIR-1064C Brake Dynamics, March 1, 1993 25.735 

SAE AIR-1083B Airborne Hydraulic and Control System Survivability 25.1435 

 for Military Aircraft, June 1994 

SAE AIR-1116 Fluid Properties, July 1992 25.1435 

SAE AIR-1362 Physical Properties of Hydraulic Fluids, Nov. 1991 25.1435 

SAE AIR-1739 Information on Antiskid Systems, March 1, 1993 25.735 

SAE AIR-1899 Aircraft Hydraulic System Characteristics, Nov. 1991 25.1435 

SAE AIR-1904 Tire Spray Suppression-Airplane, Design Consideration 25.733 

and Testing, Jan. 1, 1997 

SAE AIR-1918 Comparison of Hydraulic System Cleanliness Procedures 25.1435 

and Requirements for Ten Aerospace Companies, Nov. 1991 

SAE AIR-1934A Use of Carbon Heat Sink Brakes on Aircraft, March 1, 1995 25.735 

SAE AIR-4003 Turbine Engine Containment, Jan. 1991 25.1435 

SAE AIR-4150 Inspection of In-Service Airborne Accumulators, Feb. 1993 25.1435 

SAE AIR-4566 Crashworthiness Landing Gear Design, July 1992 25.729 

SAE ARP-24B Determination of Hydraulic Pressure Drop, Nov. 1991 25.1435 

SAE ARP-219 Procedure and Method for Conducting Test of Hydraulic 25.1435 

Components in Contamination Controlled System 

SAE ARP-490E Electro-Hydraulic Servovalves, Sept. 1993 25.1435 

SAE ARP 507C Wheels and Brakes, Supplementary Criteria for Design 25.735 

 Endurance-Civil Transport Aircraft 

SAE ARP-597C Wheels and Brakes, Supplementary Criteria for Design 25.735 

 Endurance, Civil Transport Aircraft, April 1, 1996 

SAE ARP-598B The Determination of Particulate Contamination in Liquids 25.1435 

176
3/14/2000 AC 25-22 

 by the Particle Count Method 

SAE ARP-763 Accumulators, Ground, Hydropneumatic Pressure, 25.1435 

 Oct. 15, 1962 

SAE ARP-785 Procedure for the Determination of Particulate 25.1435 

Contamination in Hydraulic Fluids by the Control Filter 

 Gravimetric Procedure 

SAE ARP 813A Maintainability Recommendations for Aircraft Wheels 25.735 

 and Brakes 

SAE ARP-813B Maintainability Recommendations for Aircraft Wheels 25.735 

 and Brakes, April 1, 1993 

SAE ARP-862A Skid Control Performance, April 1, 1996 25.735 

SAE ARP-994 Design of Tubing Installations for Aerospace Fluid 25.1435 

 Power Systems, Nov. 1991 

SAE ARP 1064B Brake Dynamics 25.735 

SAE ARP-1070B Design and Testing of Antiskid Brake Control, 25.735 

 April 1, 1996 

SAE ARP-1084 Hydraulic External Leakage for In-Service Components, 25.1435 

 November 1991 

SAE ARP-1280A Aerospace Application Guide for Hydraulic Power 25.1435 

 Transfer Units, July 1994 

SAE ARP-1281B Actuators: Aircraft Flight Controls, Power Operated, 25.1435 

 Hydraulic, General Specification, May, 1993 

SAE ARP-1311A Landing Gear - Aircraft, January 1995 25.729 

SAE-ARP 1322 Overpressurization Release Devices, March 1, 1992 25.731, 25.733 

SAE ARP-1619 Replacement and Modified Brakes and Wheels, 25.735 

 April 1, 1993 

SAE APR-1709 Coupling Assembly, Hydraulic Self Sealing, Quick 25.1435 

 Disconnect, April 1994 

SAE ARP-1786 Wheel Roll on Rim Criteria for Aircraft Application, 25.731 

 July 1, 1994 

SAE ARP-1832 Color Identification for O-Ring Seals, Dec. 1992 25.1435 

SAE-ARP-1907 Automatic Braking Systems Requirements, April 1, 1993 25.735 

SAE ARP-4379 Accumulator, Hydraulic, Cylindrical Aircraft, June 1991 25.1435 

SAE ARP-4752 Aerospace - Design and Installation of Commercial 25.1435 

Transport Aircraft Hydraulic Systems, Sept. 1994 

SAE-ARP-4834 Recommended Practice for Retreaded Aircraft Tires- 25.733 

Radial and Bias, Nov. 1, 1995 

SAE AS-483A Skid Control Equipment, March 1, 1992 25.735 

SAE AS 595B Civil Type Aircraft and Variable Delivery Hydraulic 25.1435 

 Pump, March 1995 

SAE AS-707B Thermal Sensitive Inflation Pressure Release Devices 25.731, 25.733 

for Tubeless Aircraft Wheels, March 1, 1992 

177 
3/14/2000 AC 25-22 

SAE AS-1145A Aircraft Brake Temperature Monitor System (BTMS), 25.735 

 March 1, 1992 

SAE-AS-1188 Aircraft Tire Inflation-Deflation Equipment, 25.733 

 April 1, 1993 

SAE AS-1241B Fire Resistant Phosphate Ester Hydraulic Fluid for 25.1435 

 Aircraft, Feb. 18, 1992 

SAE AS-4059 Cleanliness Classification for Hydraulic Fluids 25.1435 

SAE-AS-4833 Aircraft New Tire Standard-Bias and Radial, June 1, 1995 25.733 

SAE A-6 Committee has prepared the Aerospace Fluid Power 25.1435 

Actuation & Control Technologies documents, related to 

methods for testing and measurement of fluid contamination. 

SAE A-10 Committee has prepared the Aircraft Oxygen Equipment 25.1441, 25.1443, 

Handbook, 1992, or latest revision, containing all Aerospace 25.1445, 25.1447, 

Standards, Aerospace Recommended Practices, and Aerospace 25.1453 

Information Reports related to oxygen equipment published 

 by SAE 

SAE publication “Aircraft Flight Control Actuation System Design,” 25.1435 

by E. T. Raymond and C. C. Chenoweth 

178
3/14/2000 AC 25-22 

APPENDIX 4

6. Miscellaneous References in this AC:

These documents provide additional information, guidance, and/or standards, and are available 

from the National Technical Information Service, 5285 Port Royal Road, Springfield, VA 22161, 

USA. 

Number Title Section

 Reference

DOT/FAA/AR-96/122 Development of a Minimum Performance Standard 25.854 

for Lavatory Trash Receptacle Automatic Fire 

 Extinguishers, February 1997 

DOT/FAA/CT-88/8-1 Aircraft Icing Handbook (three volumes) 25.1419 

DOT/FAA/CT/83/1 FAA Technical Center Report, Analysis of 25.857 

Dissipation of Gaseous Extinguisher Agents in 

 Ventilated Compartments, May 1983

CAA Technical 

Development Report 

No. 146, Appendix I Evaluation of Flight Fire Protection Means 25.857 

 for Inaccessible Aircraft Baggage Compartments, 

 June 1951 

FAA Technical 

Report ADS-4 Engineering Summary of Airframe Icing 25.1419 

 Technical Data, Dec. 10, 1963 

FAA Report 

FAA-AEQ-77-13 Ozone Concentration By Latitude, Altitude, 25.832 

 and Month, Near 80° West, August 1977 

FAA Report 

FAA-AM-79-20 Effects of Ozone on Exercising and Sedentary 25.832 

 Adult Men and Women Representative of the 

Flight Attendant Population, October 1979 

Generic Issue Paper PCARGOE.DOC, Protection of Critical Systems 25.855 

& Equipment within Class E Cargo Compartments 

179 
3/14/2000 AC 25-22 

180

NPRM 89-31 Minimum Air Flow, Requirements, Fresh Air 25.831 

NPRM 93-8 Determination of Allowable Worn Brake 25.735 

 Limits 

AD-93-07-15 Class B Fire Extinguishing Methods 25.857 

The Tire and Rim Association, Inc. prepares a yearbook which lists aircraft tire 25.731 

 and rim sizes and ratings. 25.735 

 The address is: Tire and Rim Association, Inc. 

 175 Montrose West Ave., Suite 150 

 Copley, OH 44321, USA. 
3/14/2000 AC 25-22 

THIS PAGE INTENTIONALLY LEFT BLANK.

188 
3/14/2000 AC 25-22 

APPENDIX 4

7. Cross Reference for Other Handbooks.While not regulatory in nature, these handbooks contain information that may be of interest to an applicant. Theadditional information, guidance, and/or standards. Copies of the handbooks are available from the FAA TranspDirectorate, 1601 Lind Ave. SW, Renton, WA 98055, USA. Part 25 Airworthiness Standards: Transport Category Airplanes Mechanical Systems

Subpart B - Performance

Section Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

25.109 Accelerate-stop

distance 

x x x

25.125 Landing x x xSubpart C - Structures
3/14/2000 AC 25-22 

Section Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

25.581 Lightning protection x xSubpart D - Design and Construction

Section Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

Control Systems 25.671 General x x x25.672 Stability

augmentation 

x x

25.677 Trim systems x25.699 Lift and drag device 

ind. 

x

25.701 Flap interconnection x x x x25.703 Takeoff warning

system

x

Landing Gear 25.729 Retracting mechanism x x x x25.731 Wheels x x x25.733 Tires x x25.735 Brakes x x x x190 
3/14/2000 AC 25-22 

Personnel/Cargo Accommodation 25.773 Pilot compartment

view 

x x x

25.783 Doors x x xVentilation and Heating 25.831 Ventilation x x25.832 Cabin ozone

concentration 

x

25.833 Combustion heating

systems x 

Pressurization 25.841 Pressurized cabins x x25.843 Tests for pressurized 

cabins x x 

Section Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

Fire Protection 25.851 Fire extinguishers x x25.854 Lavatory fire

protection 

x

25.855 Cargo or baggage 

compartments x x 

25.857 Cargo compartment

classification x x 

25.858 Cargo Compartment
3/14/2000 AC 25-22 

fire detection systems x 

25.863 Flammable fluid fire 

protection x x 

25.865 Fire protection of 

flight controls x x 

25.869 Fire protection:

systems 

x x x

Subpart E - PowerplantSection Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

General25.943 Negative acceleration x x xFuel System Components 25.1001 Fuel jettisoning

system

x x x x

192 
3/14/2000 AC 25-22 

Section Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

Powerplant Fire Protection25.1183 Flammable fluidcarrying components x x 25.1185 Flammable fluids x x25.1189 Shutoff means x xSubpart F - EquipmentSection Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

General25.1301 Function and

installation 

x x x x x

25.1309 Equipment, system, &

installation x x x x x 

25.1316 System lightning

protection x x 

25.1322 Warning, caution, & 

advisory x x 

25.1351 System lightning

protection x x 
3/14/2000 AC 25-22 

25.1359 Electrical system fire 

& smoke protection (now 25.869) x x 

Safety Equipment 25.1419 Ice protection x xSection Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

Miscellaneous Equipment 25.1431 Electronic equipment x x x x25.1433 Vacuum systems x25.1435 Hydraulic systems x25.1438 Pressurization &

pneumatic systems x 

25.1439 Protective breathing

equipment x 

25.1441 Oxygen equipment

and supply x 

25.1443 Minimum mass flow

of supplemental oxygen 

x 

25.1445 Equipment standards

for oxygen distributing system x 

25.1447 Equipment standards

for oxygen dispensing units 

x 

25.1449 Means for194 
3/14/2000 AC 25-22 

determining use of oxygen 

x 

25.1450 Chemical oxygen

generators x 

25.1451 Fire protection: 

oxygen equipment (now 25.869) 

x 

25.1453 Protection of oxygen 

equipment fromrupture 

x 

25.1455 Draining of fluids 

subject to freezing x 

25.1461 Equipment containing

high energy rotors x 
3/14/2000 AC 25-22 

Subpart G- Operating LimitationsSection Subject

Order 8110.8

AC 25-7A

AC Draft Propulsion Handbook

AC 25-

21

AC Draft Crashworthiness Handbook

AC Draft Electrical Systems 

Handbook

AMSH

General 25.1529 Instructions for

continued airworthiness 

x 

25.1541 Markings and

placards 

x

25.1557 Miscellaneous

markings and placards 

x 

25.1581 Airplane flight

manual 

x

196 
3/14/2000 AC 25-22 

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